WorldWideScience

Sample records for solid rocket nozzle

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

    Science.gov (United States)

    2016-02-16

    Laboratory, Edwards, CA Abstract In a solid rocket motor (SRM), when the aluminum based propellant combusts, the fuel is oxidized into alumina (Al2O3...34Chemical Erosion of Refractory-Metal Nozzle Inserts in Solid - Propellant Rocket Motors," J. Propulsion and Power, Vol. 25, no.1,, 2009. [4] E. Y. Wong...34 Solid Rocket Nozzle Design Summary," in 4th AIAA Propulsion Joint Specialist Conference, Cleveland, OH, 1968. [5] Nayfeh, A. H.; Saric, W. S

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

    Science.gov (United States)

    Nayani, Kishore Nath; Bajaj, Dinesh Kumar

    2017-10-01

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

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

    Science.gov (United States)

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

    2018-02-01

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

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

    Science.gov (United States)

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

    1987-01-01

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

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

    Science.gov (United States)

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

    1996-11-01

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

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

    Science.gov (United States)

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

    1991-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Lin Sun

    2016-06-01

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

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

    Science.gov (United States)

    Yen, Yi-Hsin

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

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

    Science.gov (United States)

    Clayton, J. Louie

    2012-01-01

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

  10. Integrated Composite Rocket Nozzle Extension, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — ORBITEC proposes to develop and demonstrate an Integrated Composite Rocket Nozzle Extension (ICRNE) for use in rocket thrust chambers. The ICRNE will utilize an...

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

    Science.gov (United States)

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

    1993-06-01

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

  12. Analysis of film cooling in rocket nozzles

    Science.gov (United States)

    Woodbury, Keith A.

    1992-01-01

    Computational Fluid Dynamics (CFD) programs are customarily used to compute details of a flow field, such as velocity fields or species concentrations. Generally they are not used to determine the resulting conditions at a solid boundary such as wall shear stress or heat flux. However, determination of this information should be within the capability of a CFD code, as the code supposedly contains appropriate models for these wall conditions. Before such predictions from CFD analyses can be accepted, the credibility of the CFD codes upon which they are based must be established. This report details the progress made in constructing a CFD model to predict the heat transfer to the wall in a film cooled rocket nozzle. Specifically, the objective of this work is to use the NASA code FDNS to predict the heat transfer which will occur during the upcoming hot-firing of the Pratt & Whitney 40K subscale nozzle (1Q93). Toward this end, an M = 3 wall jet is considered, and the resulting heat transfer to the wall is computed. The values are compared against experimental data available in Reference 1. Also, FDNS's ability to compute heat flux in a reacting flow will be determined by comparing the code's predictions against calorimeter data from the hot firing of a 40K combustor. The process of modeling the flow of combusting gases through the Pratt & Whitney 40K subscale combustor and nozzle is outlined. What follows in this report is a brief description of the FDNS code, with special emphasis on how it handles solid wall boundary conditions. The test cases and some FDNS solution are presented next, along with comparison to experimental data. The process of modeling the flow through a chamber and a nozzle using the FDNS code will also be outlined.

  13. Design methods in solid rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    1987-03-01

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

  14. History of Solid Rockets

    Science.gov (United States)

    Green, Rebecca

    2017-01-01

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

  15. Optimization of Tape Winding Process Parameters to Enhance the Performance of Solid Rocket Nozzle Throat Back Up Liners using Taguchi's Robust Design Methodology

    Science.gov (United States)

    Nath, Nayani Kishore

    2017-08-01

    The throat back up liners is used to protect the nozzle structural members from the severe thermal environment in solid rocket nozzles. The throat back up liners is made with E-glass phenolic prepregs by tape winding process. The objective of this work is to demonstrate the optimization of process parameters of tape winding process to achieve better insulative resistance using Taguchi's robust design methodology. In this method four control factors machine speed, roller pressure, tape tension, tape temperature that were investigated for the tape winding process. The presented work was to study the cogency and acceptability of Taguchi's methodology in manufacturing of throat back up liners. The quality characteristic identified was Back wall temperature. Experiments carried out using L 9 ' (34) orthogonal array with three levels of four different control factors. The test results were analyzed using smaller the better criteria for Signal to Noise ratio in order to optimize the process. The experimental results were analyzed conformed and successfully used to achieve the minimum back wall temperature of the throat back up liners. The enhancement in performance of the throat back up liners was observed by carrying out the oxy-acetylene tests. The influence of back wall temperature on the performance of throat back up liners was verified by ground firing test.

  16. Space Shuttle solid rocket booster

    Science.gov (United States)

    Hardy, G. B.

    1979-01-01

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

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

    Science.gov (United States)

    Suksila, Thada

    2018-01-01

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

  18. Nuclear thermal rocket nozzle testing and evaluation program

    International Nuclear Information System (INIS)

    Davidian, K.O.; Kacynski, K.J.

    1993-01-01

    Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. In this report, the Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis Research Center is outlined and the advantages of a plug nozzle are described. A facility description, experimental designs and schematics are given. Results of pretest performance analyses show that high nozzle performance can be attained despite substantial nozzle length reduction through the use of plug nozzles as compared to a convergent-divergent nozzle. Pretest measurement uncertainty analyses indicate that specific impulse values are expected to be within plus or minus 1.17%

  19. Thiokol Solid Rocket Motors

    Science.gov (United States)

    Graves, S. R.

    2000-01-01

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

  20. Technology for low cost solid rocket boosters.

    Science.gov (United States)

    Ciepluch, C.

    1971-01-01

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

  1. Analysis of film cooling in rocket nozzles

    Science.gov (United States)

    Woodbury, Keith A.

    1993-01-01

    This report summarizes the findings on the NASA contract NAG8-212, Task No. 3. The overall project consists of three tasks, all of which have been successfully completed. In addition, some supporting supplemental work, not required by the contract, has been performed and is documented herein. Task 1 involved the modification of the wall functions in the code FDNS (Finite Difference Navier-Stokes) to use a Reynolds Analogy-based method. This task was completed in August, 1992. Task 2 involved the verification of the code against experimentally available data. The data chosen for comparison was from an experiment involving the injection of helium from a wall jet. Results obtained in completing this task also show the sensitivity of the FDNS code to unknown conditions at the injection slot. This task was completed in September, 1992. Task 3 required the computation of the flow of hot exhaust gases through the P&W 40K subscale nozzle. Computations were performed both with and without film coolant injection. This task was completed in July, 1993. The FDNS program tends to overpredict heat fluxes, but, with suitable modeling of backside cooling, may give reasonable wall temperature predictions. For film cooling in the P&W 40K calorimeter subscale nozzle, the average wall temperature is reduced from 1750R to about 1050R by the film cooling. The average wall heat flux is reduced by a factor of 3.

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

    Science.gov (United States)

    Evans, Brian

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

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

    Science.gov (United States)

    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.

  4. Longitudinal acoustic instabilities in slender solid propellant rockets : linear analysis

    OpenAIRE

    García Schafer, Juan Esteban; Liñán Martínez, Amable

    2001-01-01

    To describe the acoustic instabilities in the combustion chambers of laterally burning solid propellant rockets the interaction of the mean flow with the acoustic waves is analysed, using multiple scale techniques, for realistic cases in which the combustion chamber is slender and the nozzle area is small compared with the cross-sectional area of the chamber. Associated with the longitudinal acoustic oscillations we find vorticity and entropy waves, with a wavelength typically small compared ...

  5. The Advanced Solid Rocket Motor

    Science.gov (United States)

    Mitchell, Royce E.

    1992-01-01

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

  6. Fracture Characteristics of C/SiC Composites for Rocket Nozzle at Elevated Temperature

    Energy Technology Data Exchange (ETDEWEB)

    Yoon, Dong Hyun; Lee, Jeong Won; Kim, Jae Hoon [Chungnam Nat’l Univ., Daejeon (Korea, Republic of); Sihn, Ihn Cheol; Lim, Byung Joo [Dai-Yang Industries Co., Daejeon (Korea, Republic of)

    2016-11-15

    In a solid propulsion system, the rocket nozzle is exposed to high temperature combustion gas. Hence, choosing an appropriate material that could demonstrate adequate performance at high temperature is important. As advanced materials, carbon/silicon carbide composites (C/SiC) have been studied with the aim of using them for the rocket nozzle throat. However, when compared with typical structural materials, C/SiC composites are relatively weak in terms of both strength and toughness, owing to their quasi-brittle behavior and oxidation at high temperatures. Therefore, it is important to evaluate the thermal and mechanical properties of this material before using it in this application. This study presents an experimental method to investigate the fracture behavior of C/SiC composite material manufactured using liquid silicon infiltration (LSI) method at elevated temperatures. In particular, the effects of major parameters, such as temperature, loading, oxidation conditions, and fiber direction on strength and fracture characteristics were investigated. Fractography analysis of the fractured specimens was performed using an SEM.

  7. Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

    Science.gov (United States)

    Jones, Daniel S.; Bui, Trong T.; Ruf, Joseph H.

    2013-01-01

    For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

  8. Modification of Bonding Strength Test of WC HVOF Thermal Spray Coating on Rocket Nozzle

    Directory of Open Access Journals (Sweden)

    Bondan Sofyan

    2010-10-01

    Full Text Available One way to reduce structural weight of RX-100 rocket is by modifying the nozzle material and processing. Nozzle is the main target in weight reduction due to the fact that it contributes 30 % to the total weight of the structur. An alternative for this is by substitution of massive graphite, which is currently used as thermal protector in the nozzle, with thin layer of HVOF (High Velocity Oxy-Fuel thermal spray layer. This paper presents the characterization of nozzle base material as well as the modification of bonding strength test, by designing additional jig to facilitate testing processes while maintaining level of test accuracy. The results showed that the material used for  RX-100 rocket nozzle is confirmed to be S45C steel. Modification of the bonding strength test was conducted by utilizing chains, which improve test flexibility and maintains level of accuracy of the test.

  9. Solid rocket motor cost model

    Science.gov (United States)

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

    1972-01-01

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

  10. Design and Experimental Study on Spinning Solid Rocket Motor

    Science.gov (United States)

    Xue, Heng; Jiang, Chunlan; Wang, Zaicheng

    The study on spinning solid rocket motor (SRM) which used as power plant of twice throwing structure of aerial submunition was introduced. This kind of SRM which with the structure of tangential multi-nozzle consists of a combustion chamber, propellant charge, 4 tangential nozzles, ignition device, etc. Grain design, structure design and prediction of interior ballistic performance were described, and problem which need mainly considered in design were analyzed comprehensively. Finally, in order to research working performance of the SRM, measure pressure-time curve and its speed, static test and dynamic test were conducted respectively. And then calculated values and experimental data were compared and analyzed. The results indicate that the designed motor operates normally, and the stable performance of interior ballistic meet demands. And experimental results have the guidance meaning for the pre-research design of SRM.

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

    Science.gov (United States)

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

    1961-01-01

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

  12. Infrared Imagery of Solid Rocket Exhaust Plumes

    Science.gov (United States)

    Moran, Robert P.; Houston, Janice D.

    2011-01-01

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

  13. Structural strengthening of rocket nozzle extension by means of laser metal deposition

    Science.gov (United States)

    Honoré, M.; Brox, L.; Hallberg, M.

    2012-03-01

    Commercial space operations strive to maximize the payload per launch in order to minimize the costs of each kg launched into orbit; this yields demand for ever larger launchers with larger, more powerful rocket engines. Volvo Aero Corporation in collaboration with Snecma and Astrium has designed and tested a new, upgraded Nozzle extension for the Vulcain 2 engine configuration, denoted Vulcain 2+ NE Demonstrator The manufacturing process for the welding of the sandwich wall and the stiffening structure is developed in close cooperation with FORCE Technology. The upgrade is intended to be available for future development programs for the European Space Agency's (ESA) highly successful commercial launch vehicle, the ARIANE 5. The Vulcain 2+ Nozzle Extension Demonstrator [1] features a novel, thin-sheet laser-welded configuration, with laser metal deposition built-up 3D-features for the mounting of stiffening structure, flanges and for structural strengthening, in order to cope with the extreme load- and thermal conditions, to which the rocket nozzle extension is exposed during launch of the 750 ton ARIANE 5 launcher. Several millimeters of material thickness has been deposited by laser metal deposition without disturbing the intricate flow geometry of the nozzle cooling channels. The laser metal deposition process has been applied on a full-scale rocket nozzle demonstrator, and in excess of 15 kilometers of filler wire has been successfully applied to the rocket nozzle. The laser metal deposition has proven successful in two full-throttle, full-scale tests, firing the rocket engine and nozzle in the ESA test facility P5 by DLR in Lampoldshausen, Germany.

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

    International Nuclear Information System (INIS)

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

    2004-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2004-07-01

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

  16. Flow-Structural Interaction in Solid Rocket Motors

    National Research Council Canada - National Science Library

    Murdock, John

    2004-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

  18. Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

    Science.gov (United States)

    Marshall, Joel H.

    A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

  19. A Optimal Design of the Rocket Nozzle Wall by the Numerical Method

    Directory of Open Access Journals (Sweden)

    Jin-Won Kim

    1986-06-01

    Full Text Available It is the aims of this study to choose the materials and determine the material thickness of laminated Rocket Nozzle Wall operating at high pressure and high temperature. The heat conduction analysis of each layer was performed by Crank Nicolson method changing the thickness and the materials for the input data of Tungsten, Graphite, Alumina, Aluminum, Molybdenum, Plastic laminate. The results of the study of the study for pressure of 93.5 kg/cm^2 and temperature of 3000 degC in the nozzle dia of 40 cm are as follows.

  20. Research Amplitudo Vibration On Holder Due To The Process Of Lathe Nozzle Rocket RX 450

    Science.gov (United States)

    Ediwan; Budi Djatmiko, Agus; Dody Arisandi, EfFendy; Purnomo, Heri; Ibadi, Mahfud

    2018-04-01

    The main function of the rocket nozzle is to convert the enthalpy efficiency from combustion gas to kinetic energy and also to make high velocity out of the gas. The rocket nozzle usually consists of a converging and diverging part. With a smaller area on the neck and enlarged at the exit area. The velocity flow through the nozzle enlarges into the speed of sound through the neck and then becomes super sonic in the divergent part. Nozzle making or machining using conventional lathes, first performed is drilling on a massive metal that is bonded to the veneer, then after a sufficient gap is done deep-boring. At the time of the process of lathe in the nozzle RX 450 there is an obstacle that is vibrating tool holder chisel or holder so it is worried about not precision of the process of lathe. This should not happen because it can cause failure in the latter for it needs to be studied and studied further so that the lathe process goes accordingly. The holder material of ST 60 with a modulus of elasticity 200 GPa and a nozzle material of AISI 4340 alloy steel with σyield = 470 MPa, Shear Modulus G = 80 GPa. The purpose of this research is to observe the amplitude of vibration on the holder due to RX- 450 nozzle lathe processing for the purpose of amplitude that occurs in accordance with the desired so that the nozzle structure is no damage process. The result of the research was obtained holder with length (L) 80cm, profile width (B) 5 cm, height of profile (H) 10 cm, turning machine ω = 8.98 rad / sec and natural holder frequency ωn = 89.8 rad / second, Amplitude of vibration of δ = 1.21 mm, while the amplitude of the design X = 1.22 mm From the results of this study it can be said that the holder of a chisel or holder can be used as a tool at the time of RX nozzle retrieval process and is quite safe because it works under the condition ω/ω n Rocket Payload "AKPV Engineering University of Wyoming 2009 )

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

    Science.gov (United States)

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

    1990-01-01

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

  2. Manufacturing Process Developments for Regeneratively-Cooled Channel Wall Rocket Nozzles

    Science.gov (United States)

    Gradl, Paul; Brandsmeier, Will

    2016-01-01

    Regeneratively cooled channel wall nozzles incorporate a series of integral coolant channels to contain the coolant to maintain adequate wall temperatures and expand hot gas providing engine thrust and specific impulse. NASA has been evaluating manufacturing techniques targeting large scale channel wall nozzles to support affordability of current and future liquid rocket engine nozzles and thrust chamber assemblies. The development of these large scale manufacturing techniques focus on the liner formation, channel slotting with advanced abrasive water-jet milling techniques and closeout of the coolant channels to replace or augment other cost reduction techniques being evaluated for nozzles. NASA is developing a series of channel closeout techniques including large scale additive manufacturing laser deposition and explosively bonded closeouts. A series of subscale nozzles were completed evaluating these processes. Fabrication of mechanical test and metallography samples, in addition to subscale hardware has focused on Inconel 625, 300 series stainless, aluminum alloys as well as other candidate materials. Evaluations of these techniques are demonstrating potential for significant cost reductions for large scale nozzles and chambers. Hot fire testing is planned using these techniques in the future.

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

    Science.gov (United States)

    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.

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

    OpenAIRE

    Cramer, Robert Grewelle.

    1982-01-01

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

  5. Analysis of plume backflow around a nozzle lip in a nuclear rocket

    International Nuclear Information System (INIS)

    Chung, C.H.; Kim, S.C.; Stubbs, R.M.; De Witt, K.J.

    1993-06-01

    The structure of the flow around a nuclear thermal rocket nozzle lip has been investigated using the direct simulation Monte Carlo method. Special attention has been paid to the behavior of a small amount of harmful particles that may be present in the rocket exhaust gas. The harmful fission product particles are modeled by four inert gases whose molecular weights are in a range of 4 131. Atomic hydrogen, which exists in the flow due to the extremely high nuclear fuel temperature in the reactor, is also included. It is shown that the plume backflow is primarily determined by the thin subsonic fluid layer adjacent to the surface of the nozzle lip, and that the inflow boundary in the plume region has negligible effect on the backflow. It is also shown that a relatively large amount of the lighter species is scattered into the backflow region while the amount of the heavier species becomes negligible in this region due to extreme separation between the species. Results indicate that the backscattered molecules are very energetic and are fast-moving along the surface in the backflow region near the nozzle lip. 22 refs

  6. Solid Rocket Testing at AFRL (Briefing Charts)

    Science.gov (United States)

    2016-10-21

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

  7. Analysis of plasma behavior in a magnetic nozzle of laser fusion rocket

    International Nuclear Information System (INIS)

    Nagamine, Yoshihiko; Yoshimi, Naofumi; Nakama, Yuji; Muranaka, Takanobu; Mayumi, Takao; Nakashima, Hideki

    1997-01-01

    A magnetic nozzle concept in a laser fusion rocket is suitable for controlling the fusion plasma flow and it has an advantage that thermalization with wall structures in a thrust chamber can be avoided. Rayleigh-Taylor instability would occur at the surface of expanding plasma and it would lead to the degradation of thrust efficiency, due to diffusion of the plasma through ambient decelerating magnetic field. A 3D hybrid particle-in-cell code has been developed to analyze the plasma instability in the magnetic nozzle. The resultant linear growth rate γ of the instability is found to be 2.96 x 10 6 and it is in good agreement with the theoretical value from conventional Rayleigh Taylor instability. (author)

  8. Efficient solid rocket propulsion for access to space

    Science.gov (United States)

    Maggi, Filippo; Bandera, Alessio; Galfetti, Luciano; De Luca, Luigi T.; Jackson, Thomas L.

    2010-06-01

    Space launch activity is expected to grow in the next few years in order to follow the current trend of space exploitation for business purpose. Granting high specific thrust and volumetric specific impulse, and counting on decades of intense development, solid rocket propulsion is a good candidate for commercial access to space, even with common propellant formulations. Yet, some drawbacks such as low theoretical specific impulse, losses as well as safety issues, suggest more efficient propulsion systems, digging into the enhancement of consolidated techniques. Focusing the attention on delivered specific impulse, a consistent fraction of losses can be ascribed to the multiphase medium inside the nozzle which, in turn, is related to agglomeration; a reduction of agglomerate size is likely. The present paper proposes a model based on heterogeneity characterization capable of describing the agglomeration trend for a standard aluminized solid propellant formulation. Material microstructure is characterized through the use of two statistical descriptors (pair correlation function and near-contact particles) looking at the mean metal pocket size inside the bulk. Given the real formulation and density of a propellant, a packing code generates the material representative which is then statistically analyzed. Agglomerate predictions are successfully contrasted to experimental data at 5 bar for four different formulations.

  9. The Chameleon Solid Rocket Propulsion Model

    International Nuclear Information System (INIS)

    Robertson, Glen A.

    2010-01-01

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

  10. Solid Rocket Motor Design Using Hybrid Optimization

    Directory of Open Access Journals (Sweden)

    Kevin Albarado

    2012-01-01

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

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

    Science.gov (United States)

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

    1988-01-01

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

  12. Hot-gas-side heat transfer characteristics of subscale, plug-nozzle rocket calorimeter chamber

    Science.gov (United States)

    Quentmeyer, Richard J.; Roncace, Elizabeth A.

    1993-01-01

    An experimental investigation was conducted to determine the hot-gas-side heat transfer characteristics for a liquid-hydrogen-cooled, subscale, plug-nozzle rocket test apparatus. This apparatus has been used since 1975 to evaluate rocket engine advanced cooling concepts and fabrication techniques, to screen candidate combustion chamber liner materials, and to provide data for model development. In order to obtain the data, a water-cooled calorimeter chamber having the same geometric configuration as the plug-nozzle test apparatus was tested. It also used the same two showerhead injector types that were used on the test apparatus: one having a Rigimesh faceplate and the other having a platelet faceplate. The tests were conducted using liquid oxygen and gaseous hydrogen as the propellants over a mixture ratio range of 5.8 to 6.3 at a nominal chamber pressure of 4.14 MPa abs (600 psia). The two injectors showed similar performance characteristics with the Rigimesh faceplate having a slightly higher average characteristic-exhaust-velocity efficiency of 96 percent versus 94.4 percent for the platelet faceplate. The throat heat flux was 54 MW/m(sup 2) (33 Btu/in.(sup 2)-sec) at the nominal operating condition, which was a chamber pressure of 4.14 MPa abs (600 psia), a hot-gas-side wall temperature of 730 K (1314 R), and a mixture ratio of 6.0. The chamber throat region correlation coefficient C(sub g) for a Nusselt number correlation of the form Nu =C(sub g)Re(sup 0.8)Pr(sup 0.3) averaged 0.023 for the Rigimesh faceplate and 0.026 for the platelet faceplate.

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

    Science.gov (United States)

    Sutton, George P.

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    1989-04-01

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

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    1972-01-01

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

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

    Science.gov (United States)

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

    2001-01-01

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

  18. Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

    Science.gov (United States)

    Gradl, Paul; Valentine, Peter; Crisanti, Matthew; Greene, Sandy Elam

    2016-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

  19. Stochastic rocket dynamics under random nozzle side loads: Ornstein-Uhlenbeck boundary layer separation and its coarse grained connection to side loading and rocket response

    Energy Technology Data Exchange (ETDEWEB)

    Keanini, R.G.; Srivastava, N.; Tkacik, P.T. [Department of Mechanical Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Weggel, D.C. [Department of Civil and Environmental Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Knight, P.D. [Mitchell Aerospace and Engineering, Statesville, North Carolina 28677 (United States)

    2011-06-15

    A long-standing, though ill-understood problem in rocket dynamics, rocket response to random, altitude-dependent nozzle side-loads, is investigated. Side loads arise during low altitude flight due to random, asymmetric, shock-induced separation of in-nozzle boundary layers. In this paper, stochastic evolution of the in-nozzle boundary layer separation line, an essential feature underlying side load generation, is connected to random, altitude-dependent rotational and translational rocket response via a set of simple analytical models. Separation line motion, extant on a fast boundary layer time scale, is modeled as an Ornstein-Uhlenbeck process. Pitch and yaw responses, taking place on a long, rocket dynamics time scale, are shown to likewise evolve as OU processes. Stochastic, altitude-dependent rocket translational motion follows from linear, asymptotic versions of the full nonlinear equations of motion; the model is valid in the practical limit where random pitch, yaw, and roll rates all remain small. Computed altitude-dependent rotational and translational velocity and displacement statistics are compared against those obtained using recently reported high fidelity simulations [Srivastava, Tkacik, and Keanini, J. Appl. Phys. 108, 044911 (2010)]; in every case, reasonable agreement is observed. As an important prelude, evidence indicating the physical consistency of the model introduced in the above article is first presented: it is shown that the study's separation line model allows direct derivation of experimentally observed side load amplitude and direction densities. Finally, it is found that the analytical models proposed in this paper allow straightforward identification of practical approaches for: (i) reducing pitch/yaw response to side loads, and (ii) enhancing pitch/yaw damping once side loads cease. (Copyright copyright 2011 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim)

  20. Evaluation of the Effect of Exhausts from Liquid and Solid Rockets on Ozone Layer

    Science.gov (United States)

    Yamagiwa, Yoshiki; Ishimaki, Tetsuya

    This paper reports the analytical results of the influences of solid rocket and liquid rocket exhausts on ozone layer. It is worried about that the exhausts from solid propellant rockets cause the ozone depletion in the ozone layer. Some researchers try to develop the analytical model of ozone depletion by rocket exhausts to understand its physical phenomena and to find the effective design of rocket to minimize its effect. However, these models do not include the exhausts from liquid rocket although there are many cases to use solid rocket boosters with a liquid rocket at the same time in practical situations. We constructed combined analytical model include the solid rocket exhausts and liquid rocket exhausts to analyze their effects. From the analytical results, we find that the exhausts from liquid rocket suppress the ozone depletion by solid rocket exhausts.

  1. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    OpenAIRE

    Abdelaziz Almostafa; Guozhu Liang; Elsayed Anwer

    2018-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    1995-10-01

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

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

    Directory of Open Access Journals (Sweden)

    Xiaodong He

    2017-04-01

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

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

    Vyas, Anand B.

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

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

    Science.gov (United States)

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

    2013-12-01

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

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

    Science.gov (United States)

    Elliott, T. S.; Majdalani, J.

    2014-11-01

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

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

    International Nuclear Information System (INIS)

    Elliott, T S; Majdalani, J

    2014-01-01

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

  9. Scale Effects on Solid Rocket Combustion Instability Behaviour

    OpenAIRE

    David R. Greatrix

    2011-01-01

    The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combusti...

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

    OpenAIRE

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

    2017-01-01

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

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

    Science.gov (United States)

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

    2017-09-01

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

  12. Nutation instability of spinning solid rocket motor spacecraft

    Directory of Open Access Journals (Sweden)

    Dan YANG

    2017-08-01

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

  13. A novel kind of solid rocket propellant

    Energy Technology Data Exchange (ETDEWEB)

    Lo, R.E. [Berlin University of Technology (Germany). Rocket Technology at the Aerospace Institute (ILR)

    1998-09-01

    Cryogenic Solid Propellants (CSPs) combine the simplicity of conventional solid propulsion with the high performance of liquid propulsion. By introducing materials that require cooling for remaining solid, CSPs offer an almost unlimited choice of propellant constituents that mights be selected with respect to specific impulse, density or environmental protection. The prize to be paid for these advantages is the necessity of constant cooling and the requirement of special design features that provide combustion control by moving from deflagration to hybrid like boundary layer combustion. This is achieved by building the solid propellant grains out of macroscopic elements rather than using the quasi homogeneous mixture of conventional composites. The elements may be coated, providing protection and support. Different elements may be designed for individual tasks and serve as modules for ignition, sustained combustion, gas generation, combustion efficiency enhancement, etc. Modular dissected grains offer many new ways of interaction inside the combustion chamber and new degrees of freedom for the designer of such `multiple internal hybrid grains`. At a preliminary level, a study finished in Germany 1997 demonstrated large payload gains when the US space Shuttle and the ARIANE 5 boosters were replaced by CSP-boosters. A very preliminary cost analysis resulted in development costs in the usual magnitude (but not in higher ones). Costs of operation were identified as crucial, but not established. Some experimental work in Germany is scheduled to begin in 1998, almost all details in this article (and many more that were not mentioned - most prominent cost analyses of CSP development and operations) wait for deeper analysis and verification. Actually, a whole new world new of world of chemical propulsion awaits exploration. The topic can be looked up and discussed at the web site of the Advanced Propulsion Workshop of the International Academy of Astronautics. The author

  14. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    Directory of Open Access Journals (Sweden)

    Abdelaziz Almostafa

    2018-01-01

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

  15. Specific Impulses Losses in Solid Propellant Rockets

    Science.gov (United States)

    1974-12-17

    the solid we go back to the wall flux. Platinum film thermometric probes [77, 78], developed for somewhat similar problems, were used without succ...AS: E Paper 63-41A 207, 1964. [83) A.D. KIDEIR .A XU - J.A. CAHILL - The density of liquid aluminium oxide. J. Inovg. luc.. Chem. vol.14, no 3-4, p...IELLOR - I. GLL.ASIJ - Augo.nted ignition effi ciency for aluminium . Combustion Science rand Tcchnology, vol. I, p. 437, 1970. [90s EIJ. IJJ tBAU

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

    Science.gov (United States)

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

    1976-01-01

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

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

    Science.gov (United States)

    Laubacher, Brian A.

    2000-01-01

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

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

    Science.gov (United States)

    Moore, Dennis R.; Phelps, Willie J.

    2011-01-01

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

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

    Science.gov (United States)

    Colbaugh, Lauren

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

  20. MEMS-Based Solid Propellant Rocket Array Thruster

    Science.gov (United States)

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

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

  1. Rocket nozzle expansion ratio analysis for dual-fuel earth-to-orbit vehicles

    Science.gov (United States)

    Martin, James A.

    1989-01-01

    Results are reported from a recent study of the effects of Space Shuttle Main Engine expansion ratio modifications, in the cases of both single-stage and two-stage systems. Two-position nozzles were employed; after varying the lower expansion ratio while the higher was held constant at 120, the lower expansion ratio was held constant at 40 or 60 while the higher expansion ratio was varied. The expansion ratios for minimum vehicle dry mass are different for single-stage and two-stage systems. For two-stage systems, a single expansion ratio of 77.5 provides a lower dry mass than any two-position nozzle.

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

    Science.gov (United States)

    Fischbach, Sean

    2014-01-01

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

  3. Rocket Solid Propellant Alternative Based on Ammonium Dinitramide

    Directory of Open Access Journals (Sweden)

    Grigore CICAN

    2017-03-01

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

  4. Transient Three-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

    Science.gov (United States)

    Wang, Ten-See

    2004-01-01

    Three-dimensional numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, and pressure-based computational fluid dynamics formulation, and a simulated inlet condition based on a system calculation. Finite-rate chemistry was used throughout the study so that combustion effect is always included, and the effect of wall cooling on side load physics is studied. The side load physics captured include the afterburning wave, transition from free- shock to restricted-shock separation, and lip Lambda shock oscillation. With the adiabatic nozzle, free-shock separation reappears after the transition from free-shock separation to restricted-shock separation, and the subsequent flow pattern of the simultaneous free-shock and restricted-shock separations creates a very asymmetric Mach disk flow. With the cooled nozzle, the more symmetric restricted-shock separation persisted throughout the start-up transient after the transition, leading to an overall lower side load than that of the adiabatic nozzle. The tepee structures corresponding to the maximum side load were addressed.

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

    National Research Council Canada - National Science Library

    Hanning-Lee, M

    1996-01-01

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

  6. Five-Segment Solid Rocket Motor Development Status

    Science.gov (United States)

    Priskos, Alex S.

    2012-01-01

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

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

    Science.gov (United States)

    Quing, Xinlin; Beard, Shawn; Zhang, Chang

    2011-01-01

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

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

    Science.gov (United States)

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

    1982-01-01

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

  9. Scale effects on solid rocket combustion instability behaviour

    Energy Technology Data Exchange (ETDEWEB)

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

    2011-07-01

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

  10. Scale Effects on Solid Rocket Combustion Instability Behaviour

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2011-01-01

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

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

    Science.gov (United States)

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

    2000-01-01

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

  12. High-speed schlieren imaging of rocket exhaust plumes

    Science.gov (United States)

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

    2016-11-01

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

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

    Science.gov (United States)

    Sambamurthi, Jay K.

    1995-01-01

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

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

    Science.gov (United States)

    Sambamurthi, Jay K.

    1995-01-01

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

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

    National Research Council Canada - National Science Library

    Cubbage, James M

    1965-01-01

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

  16. Experimental validation of solid rocket motor damping models

    Science.gov (United States)

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

    2017-12-01

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

  17. Experimental validation of solid rocket motor damping models

    Science.gov (United States)

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

    2018-06-01

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

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

    Science.gov (United States)

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

    1994-01-01

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

  19. Shock wave fabricated ceramic-metal nozzles

    NARCIS (Netherlands)

    Carton, E.P.; Stuivinga, M.E.C.; Keizers, H.L.J.; Verbeek, H.J.; Put, P.J. van der

    1999-01-01

    Shock compaction was used in the fabrication of high temperature ceramic-based materials. The materials' development was geared towards the fabrication of nozzles for rocket engines using solid propellants, for which the following metal-ceramic (cermet) materials were fabricated and tested: B4C-Ti

  20. A Multiconstrained Ascent Guidance Method for Solid Rocket-Powered Launch Vehicles

    Directory of Open Access Journals (Sweden)

    Si-Yuan Chen

    2016-01-01

    Full Text Available This study proposes a multiconstrained ascent guidance method for a solid rocket-powered launch vehicle, which uses a hypersonic glide vehicle (HGV as payload and shuts off by fuel exhaustion. First, pseudospectral method is used to analyze the two-stage launch vehicle ascent trajectory with different rocket ignition modes. Then, constraints, such as terminal height, velocity, flight path angle, and angle of attack, are converted into the constraints within height-time profile according to the second-stage rocket flight characteristics. The closed-loop guidance method is inferred by different spline curves given the different terminal constraints. Afterwards, a thrust bias energy management strategy is proposed to waste the excess energy of the solid rocket. Finally, the proposed method is verified through nominal and dispersion simulations. The simulation results show excellent applicability and robustness of this method, which can provide a valuable reference for the ascent guidance of solid rocket-powered launch vehicles.

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

    Science.gov (United States)

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

    2018-03-01

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

  2. Working-cycle processes in solid-propellant rocket engines (Handbook). Rabochie protsessy v raketnykh dvigateliakh tverdogo topliva /Spravochnik/

    Energy Technology Data Exchange (ETDEWEB)

    Shishkov, A.A.; Panin, S.D.; Rumiantsev, B.V.

    1989-01-01

    Physical and mathematical models of processes taking place in solid-propellant rocket engines and gas generators are presented in a systematic manner. The discussion covers the main types of solid propellants, the general design and principal components of solid-propellant rocket engines, the combustion of a solid-propellant charge, thermodynamic calculation of combustion and outflow processes, and analysis of gasdynamic processes in solid-propellant rocket engines. 40 refs.

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

    National Research Council Canada - National Science Library

    2003-01-01

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

  4. Construction and design of solid-propellant rocket engines. Konstruktsiia i proektirovanie raketnykh dvigatelei tverdogo topliva

    Energy Technology Data Exchange (ETDEWEB)

    Fakhrutdinov, I.K.; Kotel' nikov, A.V.

    1987-01-01

    Methods for assessing the durability of different components of solid-propellant rocket engines are presented. The following aspects of engine development are discussed: task formulation, parameter calculation, construction scheme selection, materials, and durability assessment. 45 references.

  5. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    National Research Council Canada - National Science Library

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

    2007-01-01

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

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

    Science.gov (United States)

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

    2012-01-01

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

  7. Solid Rocket Booster Large Main and Drogue Parachute Reliability Analysis

    Science.gov (United States)

    Clifford, Courtenay B.; Hengel, John E.

    2009-01-01

    The parachutes on the Space Transportation System (STS) Solid Rocket Booster (SRB) are the means for decelerating the SRB and allowing it to impact the water at a nominal vertical velocity of 75 feet per second. Each SRB has one pilot, one drogue, and three main parachutes. About four minutes after SRB separation, the SRB nose cap is jettisoned, deploying the pilot parachute. The pilot chute then deploys the drogue parachute. The drogue chute provides initial deceleration and proper SRB orientation prior to frustum separation. At frustum separation, the drogue pulls the frustum from the SRB and allows the main parachutes that are mounted in the frustum to unpack and inflate. These chutes are retrieved, inspected, cleaned, repaired as needed, and returned to the flight inventory and reused. Over the course of the Shuttle Program, several improvements have been introduced to the SRB main parachutes. A major change was the replacement of the small (115 ft. diameter) main parachutes with the larger (136 ft. diameter) main parachutes. Other modifications were made to the main parachutes, main parachute support structure, and SRB frustum to eliminate failure mechanisms, improve damage tolerance, and improve deployment and inflation characteristics. This reliability analysis is limited to the examination of the SRB Large Main Parachute (LMP) and drogue parachute failure history to assess the reliability of these chutes. From the inventory analysis, 68 Large Main Parachutes were used in 651 deployments, and 7 chute failures occurred in the 651 deployments. Logistic regression was used to analyze the LMP failure history, and it showed that reliability growth has occurred over the period of use resulting in a current chute reliability of R = .9983. This result was then used to determine the reliability of the 3 LMPs on the SRB, when all must function. There are 29 drogue parachutes that were used in 244 deployments, and no in-flight failures have occurred. Since there are no

  8. Numerical Study on Similarity of Plume’s Infrared Radiation from Reduced Scaling Solid Rocket

    Directory of Open Access Journals (Sweden)

    Xiaoying Zhang

    2015-01-01

    Full Text Available Similarity of plume radiation between reduced scaling solid rocket models and full scale ones in ground conditions has been taken for investigation. Flow and radiation of plume from solid rockets with scaling ratio from 0.1 to 1 have been computed. The radiative transfer equation (RTE is solved by the finite volume method (FVM in infrared band 2~6 μm. The spectral characteristics of plume gases have been calculated with the weighted-sum-of-gray-gas (WSGG model, and those of the Al2O3 particles have been solved by the Mie scattering model. Our research shows that, with the decreasing scaling ratio of the rocket engine, the radiation intensity of the plume decreases with 1.5~2.5 power of the scaling ratio. The infrared radiation of the plume gases shows a strong spectral dependency, while that of the Al2O3 particles shows grey property. Spectral radiation intensity of the high temperature core of the solid rocket plume increases greatly in the peak absorption spectrum of plume gases. Al2O3 particle is the major radiation composition in the rocket plume, whose scattering coefficient is much larger than its absorption coefficient. There is good similarity between spectral variations of plumes from different scaling solid rockets. The directional plume radiation rises with the increasing azimuth angle.

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

    Science.gov (United States)

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

    2014-11-01

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

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

    International Nuclear Information System (INIS)

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

    2014-01-01

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

  11. Particle Impact Erosion. Volume 4. User’s Manual Erosion Prediction Procedure for Rocket Nozzle Expansion Region

    Science.gov (United States)

    1983-05-01

    empirical erosion model, with use of the debris-layer model optional. 1.1 INTERFACE WITH ISPP ISPP is a collection of computer codes designed to calculate...expansion with the ODK code, 4. A two-dimensional, two-phase nozzle expansion with the TD2P code, 5. A turbulent boundary layer solution along the...INPUT THERMODYNAMIC DATA FOR TEMPERATURESBELOW 300°K OIF NEEDED) NO A• 11 READ SSP NAMELIST (ODE. BAL. ODK . TD2P. TEL. NOZZLE GEOMETRY) PROfLM 2

  12. Solid propellant processing factor in rocket motor design

    Science.gov (United States)

    1971-01-01

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

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

    Data.gov (United States)

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

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

    Science.gov (United States)

    Sforzini, R. H.

    1973-01-01

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

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

    National Research Council Canada - National Science Library

    Blaine, Jeanette M; Patterson, Mark; Zhang, Xiaohong; Hilmas, Greg; Fehrenholtz, Bill

    2008-01-01

    "Next generation" aluminized propellants have become more energetic in order to impart a higher specific impulse to the system, resulting in higher temperatures and pressures that need to be contained...

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

    OpenAIRE

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

    2014-01-01

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

  17. High-Pressure Burning Rate Studies of Solid Rocket Propellants

    Science.gov (United States)

    2013-01-01

    monopropellant burning rate. The self-de§agration rates of neat AP are plotted in Fig. 2 for both pressed pellets and single crystals. There is agreement...rate data from various investigators: 1 ¡ [2]; pressed pellets : 2 ¡ [3], 3 ¡ [4], and 4 ¡ [2]; and single crystals: 5 ¡ [5], and 6 ¡ [6]. Line ¡ AP...7]. Strand or window burners have had more use in the solid propellant community. There are numerous types and styles of combustion vessels, but they

  18. Seismic tests at the HDR facility using explosives and solid propellant rockets

    International Nuclear Information System (INIS)

    Corvin, P.; Steinhilber, H.

    1981-01-01

    In blast tests the HDR reactor building and its mechanical equipment were subjected to earthquake-type excitations through the soil and the foundation. A series of six tests was carried out, two tests being made with HDR facility under operating conditions (BWR conditions, 285 0 C, 70 bar). The charges were placed in boreholes at a depth of 4 to 10 m and a distance of 16 to 25 m from the reactor building. The tests with solid propellant rockets were made in order to try a new excitation technique. The rockets used in these tests were of compact design and had a short combustion period (500 ms) at high constant thrust (100 kN per combustion chamber). These rockets were fixed to the concrete dome of the building in such a way that the thrust generated during the combustion of the propellant resulted in an impulsive load acting on the building. This type of excitation was selected with a view to investigating the global effects of the load case 'aircraft impact' on the building and the mechanical equipment. In the four tests made so far, up to four rockets were ignited simultaneously, so that the maximum impulse was 2 x 10 5 Ns. The excitation level can be markedly increased by adding further rockets. This excitation technique was characterised by excellent reproducibility of the load parameters. (orig./HP)

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

    OpenAIRE

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

    1996-01-01

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

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

    Science.gov (United States)

    Hua, Zenggong

    1992-06-01

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

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

    OpenAIRE

    David Greatrix

    2015-01-01

    The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave ...

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

    OpenAIRE

    David R. Greatrix

    2009-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Zhang Xiaoying

    2016-08-01

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

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

    Science.gov (United States)

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

    2005-01-01

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

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

    Science.gov (United States)

    Harris, Joshua A.; Patterson-Hine, Ann

    2013-01-01

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

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

    Science.gov (United States)

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

    2011-01-01

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

  7. Lessons from half a century experience of Japanese solid rocketry since Pencil rocket

    Science.gov (United States)

    Matogawa, Yasunori

    2007-12-01

    50 years have passed since a tiny rocket "Pencil" was launched horizontally at Kokubunji near Tokyo in 1955. Though there existed high level of rocket technology in Japan before the end of the second World War, it was not succeeded by the country after the War. Pencil therefore was the substantial start of Japanese rocketry that opened the way to the present stage. In the meantime, a rocket group of the University of Tokyo contributed to the International Geophysical Year in 1957-1958 by developing bigger rockets, and in 1970, the group succeeded in injecting first Japanese satellite OHSUMI into earth orbit. It was just before the launch of OHSUMI that Japan had built up the double feature system of science and applications in space efforts. The former has been pursued by ISAS (the Institute of Space and Astronautical Science) of the University of Tokyo, and the latter by NASDA (National Space Development Agency). This unique system worked quite efficiently because space activities in scientific and applicational areas could develop rather independently without affecting each other. Thus Japan's space science ran up rapidly to the international stage under the support of solid propellant rocket technology, and, after a 20 year technological introduction period from the US, a big liquid propellant launch vehicle, H-II, at last was developed on the basis of Japan's own technology in the early 1990's. On October 1, 2003, as a part of Governmental Reform, three Japanese space agencies were consolidated into a single agency, JAXA (Japan Aerospace Exploration Agency), and Japan's space efforts began to walk toward the future in a globally coordinated fashion, including aeronautics, astronautics, space science, satellite technology, etc., at the same time. This paper surveys the history of Japanese rocketry briefly, and draws out the lessons from it to make a new history of Japan's space efforts more meaningful.

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

    Science.gov (United States)

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

    1985-01-01

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

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

    Science.gov (United States)

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

    1995-01-01

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

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Kai Qu

    2013-03-01

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

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

    Science.gov (United States)

    Culick, F. E. C.

    1972-01-01

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

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

    Science.gov (United States)

    1972-01-01

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

  15. An improved heat transfer configuration for a solid-core nuclear thermal rocket engine

    International Nuclear Information System (INIS)

    Clark, J.S.; Walton, J.T.; Mcguire, M.L.

    1992-07-01

    Interrupted flow, impingement cooling, and axial power distribution are employed to enhance the heat-transfer configuration of a solid-core nuclear thermal rocket engine. Impingement cooling is introduced to increase the local heat-transfer coefficients between the reactor material and the coolants. Increased fuel loading is used at the inlet end of the reactor to enhance heat-transfer capability where the temperature differences are the greatest. A thermal-hydraulics computer program for an unfueled NERVA reactor core is employed to analyze the proposed configuration with attention given to uniform fuel loading, number of channels through the impingement wafers, fuel-element length, mass-flow rate, and wafer gap. The impingement wafer concept (IWC) is shown to have heat-transfer characteristics that are better than those of the NERVA-derived reactor at 2500 K. The IWC concept is argued to be an effective heat-transfer configuration for solid-core nuclear thermal rocket engines. 11 refs

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

    Science.gov (United States)

    Nelson, Linda A.

    2006-01-01

    After every Shuttle mission, the Solid Rocket Boosters (SRBs) are recovered and observed for missing material. Most of the SRB is covered with a cork-based thermal protection material (MCC-l). After the most recent shuttle mission, STS-114, the forward section of the booster appeared to have been impacted during flight. The darkened fracture surfaces indicated that this might have occurred early in flight. The scope of the analysis included microscopic observations to assess the degree of heat effects and locate evidence of the impact source as well as chemical analysis of the fracture surfaces and recovered foreign material using Fourier Transform Infrared Spectroscopy and Scanning Electron Microscopy/Energy Dispersive Spectroscopy. The amount of heat effects and presence of soot products on the fracture surface indicated that the material was impacted prior to SRB re-entry into the atmosphere. Fragments of graphite fibers found on these fracture surfaces were traced to slag inside the Solid Rocket Motor (SRM) that forms during flight as the propellant is spent and is ejected throughout the descent of the SRB after separation. The direction of the impact mark matches with the likely trajectory of SRBs tumbling prior to re-entry.

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

    Science.gov (United States)

    Yang, H. Q.; West, Jeff

    2014-01-01

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

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

    International Nuclear Information System (INIS)

    Marimuthu, R.; Nageswara Rao, B.

    2013-01-01

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

  19. Shelf life extension for the lot AAE nozzle severance LSCs

    Science.gov (United States)

    Cook, M.

    1990-01-01

    Shelf life extension tests for the remaining lot AAE linear shaped charges for redesigned solid rocket motor nozzle aft exit cone severance were completed in the small motor conditioning and firing bay, T-11. Five linear shaped charge test articles were thermally conditioned and detonated, demonstrating proper end-to-end charge propagation. Penetration depth requirements were exceeded. Results indicate that there was no degradation in performance due to aging or the linear shaped charge curving process. It is recommended that the shelf life of the lot AAE nozzle severance linear shaped charges be extended through January 1992.

  20. On use of hybrid rocket propulsion for suborbital vehicles

    Science.gov (United States)

    Okninski, Adam

    2018-04-01

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

  1. Altitude Compensating Nozzle

    Science.gov (United States)

    Ruf, Joseph H.; Jones, Daniel

    2015-01-01

    The dual-bell nozzle (fig. 1) is an altitude-compensating nozzle that has an inner contour consisting of two overlapped bells. At low altitudes, the dual-bell nozzle operates in mode 1, only utilizing the smaller, first bell of the nozzle. In mode 1, the nozzle flow separates from the wall at the inflection point between the two bell contours. As the vehicle reaches higher altitudes, the dual-bell nozzle flow transitions to mode 2, to flow full into the second, larger bell. This dual-mode operation allows near optimal expansion at two altitudes, enabling a higher mission average specific impulse (Isp) relative to that of a conventional, single-bell nozzle. Dual-bell nozzles have been studied analytically and subscale nozzle tests have been completed.1 This higher mission averaged Isp can provide up to a 5% increase2 in payload to orbit for existing launch vehicles. The next important step for the dual-bell nozzle is to confirm its potential in a relevant flight environment. Toward this end, NASA Marshall Space Flight Center (MSFC) and Armstrong Flight Research Center (AFRC) have been working to develop a subscale, hot-fire, dual-bell nozzle test article for flight testing on AFRC's F15-D flight test bed (figs. 2 and 3). Flight test data demonstrating a dual-bell ability to control the mode transition and result in a sufficient increase in a rocket's mission averaged Isp should help convince the launch service providers that the dual-bell nozzle would provide a return on the required investment to bring a dual-bell into flight operation. The Game Changing Department provided 0.2 FTE to ER42 for this effort in 2014.

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

    Science.gov (United States)

    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

  3. Optimization of the stand for test of hybrid rocket engines of solid fuel

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2017-01-01

    Full Text Available In the paper the laboratory experimental stand of the hybrid rocket engine of solid fuel to study ballistic parameters of the engine at burning of high-energy materials in flow of hot gas is presented. Mixture of air with nitrogen with a specified content of active oxygen is used as a gaseous oxidizer. The experimental stand has modular design and consists of system of gas supply, system of heating of gas, system for monitoring gas parameters, to which a load cell with a model engine was connected. The modular design of the stand allows to change its configuration under specific objective. This experimental stand allows to conduct a wide range of the pilot studies at interaction of a hot stream of gas with samples high-energy materials.

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

    Science.gov (United States)

    Ahmed, Waqas; Maqsood, Adnan; Riaz, Rizwan

    2018-03-01

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

  5. Structural optimization of an alternate design for the Space Shuttle solid rocket booster field joint

    Science.gov (United States)

    Barthelemy, Jean-Francois M.; Rogers, James L., Jr.; Chang, Kwan J.

    1987-01-01

    A structural optimization procedure is used to determine the shape of an alternate design for the Shuttle's solid rocket booster field joint. In contrast to the tang and clevis design of the existing joint, this alternate design consists of two flanges bolted together. Configurations with 150 studs of 1 1/8 in diameter and 135 studs of 1 3/16 in diameter are considered. Using a nonlinear programming procedure, the joint weight is minimized under constraints on either von Mises or maximum normal stresses, joint opening and geometry. The procedure solves the design problem by replacing it by a sequence of approximate (convex) subproblems; the pattern of contact between the joint halves is determined every few cycles by a nonlinear displacement analysis. The minimum weight design has 135 studs of 1 3/16 in diameter and is designed under constraints on normal stresses. It weighs 1144 lb per joint more than the current tang and clevis design.

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

    Science.gov (United States)

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

    1988-01-01

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

  7. Computation of turbulent reacting flow in a solid-propellant ducted rocket

    Science.gov (United States)

    Chao, Yei-Chin; Chou, Wen-Fuh; Liu, Sheng-Shyang

    1995-05-01

    A mathematical model for computation of turbulent reacting flows is developed under general curvilinear coordinate systems. An adaptive, streamline grid system is generated to deal with the complex flow structures in a multiple-inlet solid-propellant ducted rocket (SDR) combustor. General tensor representations of the k-epsilon and algebraic stress (ASM) turbulence models are derived in terms of contravariant velocity components, and modification caused by the effects of compressible turbulence is also included in the modeling. The clipped Gaussian probability density function is incorporated in the combustion model to account for fluctuations of properties. Validation of the above modeling is first examined by studying mixing and reacting characteristics in a confined coaxial-jet problem. This is followed by study of nonreacting and reacting SDR combustor flows. The results show that Gibson and Launder's ASM incorporated with Sarkar's modification for compressible turbulence effects based on the general curvilinear coordinate systems yields the most satisfactory prediction for this complicated SDR flowfield.

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

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2012-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2010-11-15

    The ability to predict with some accuracy a given solid rocket motor's performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size). The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor's internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor's internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale. (author)

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

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2009-03-01

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

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

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

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

    OpenAIRE

    S. Saha; D. Chakraborty

    2016-01-01

    Combustion instability in solid propellant rocket motor is numerically simulated by implementing propellant response function with quasi steady homogeneous one dimensional formulation. The convolution integral of propellant response with pressure history is implemented through a user defined function in commercial computational fluid dynamics software. The methodology is validated against literature reported motor test and other simulation results. Computed amplitude of pressure fluctuations ...

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

    Science.gov (United States)

    Stefanski, Philip L.

    2015-01-01

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

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

    Science.gov (United States)

    Heitkotter, Robert H

    1956-01-01

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

  16. Effect of ammonium perchlorate grain sizes on the combustion of solid rocket propellant

    Energy Technology Data Exchange (ETDEWEB)

    Hegab, A.; Balabel, A. [Menoufia Univ., Menoufia (Egypt). Faculty of Engineering

    2007-07-01

    The combustion of heterogeneous solid rocket propellant consisting of ammonium perchlorate (AP) particles was discussed with reference to the chemical and physical complexity of the propellant and the microscopic scale of the combustion zone. This study considered the primary flame between the decomposition products of the binder and the AP oxidizer; the primary diffusion flame from the oxidizer; density and conductivity of the AP and binder; temperature-dependent gas-phase transport properties; and, an unsteady non-planer regression surface. Three different random packing disc models for the AP particles imbedded in a matrix of a hydroxyl terminated polybutadience (HTPB) fuel-binder were used as a base of the combustion code. The models have different AP grain sizes and distribution with the fuel binder. A 2D calculation was developed for the combustion of heterogeneous solid propellant, accounting for the gas phase physics, the solid phase physics and an unsteady non-planar description of the regressing propellant surface. The mathematical model described the unsteady burning of a heterogeneous propellant by simultaneously solving the combustion fields in the gas phase and the thermal field in the solid phase with appropriate jump condition across the gas/solid interface. The gas-phase kinetics was represented by a two-step reaction mechanism for the primary premixed flame and the primary diffusion flame between the decomposition products of the HTPB and the oxidizer. The essentially-non-oscillatory (ENO) scheme was used to describe the propagation of the unsteady non-planer regression surface. The results showed that AP particle size has a significant effect on the combustion surface deformation as well as on the burning rate. This study also determined the effect of various parameters on the surface propagation speed, flame structure, and the burning surface geometry. The speed by which the combustion surface recedes was found to depend on the exposed pressure

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

    Science.gov (United States)

    1983-01-01

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

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

    Science.gov (United States)

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

    2001-01-01

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

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

    Directory of Open Access Journals (Sweden)

    David Greatrix

    2015-02-01

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

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

    Science.gov (United States)

    Clayton, Louie

    2011-01-01

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

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

    Science.gov (United States)

    Lingbloom, Mike; Plaia, Jim; Newman, John

    2006-01-01

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

  2. Fracture tolerance analysis of the solid rocket booster servo-actuator for the space shuttle

    Energy Technology Data Exchange (ETDEWEB)

    Smith, S.H.; Ghadiali, N.D.; Zahoor, A.; Wilson, M.R.

    1982-01-01

    The results of an evaluation of the fracture tolerance of three components of the thrust vector control servo-actuator for the solid rocket booster of the space shuttle are described. These components were considered as being potentially fracture critical and therefore having the potential to fall short of a desired service life of 80 missions (that is, a service life factor of 4.0 on a basic service life of 20 missions). Detailed stress analysis of the rod end, cylinder, and feedback link components was accomplished by three-dimensional finite-element stress analysis methods. A dynamic structural model of the feedback system was used to determine the dynamic inertia loads and reactions to apply to the finite-element model of the feedback link. Twenty mission stress spectra consisting of lift-off, boost, re-entry, and water impact mission segments were developed for each component based on dynamic loadings. Most components were determined to have the potential of reaching a service life of 80 missions or service life factor of 4.0. 22 refs.

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

    Science.gov (United States)

    Horstman, Matthew F.; Mulrooney, Mark

    2009-01-01

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

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

    Science.gov (United States)

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

    2016-07-01

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

  5. Development of nuclear rocket engine technology

    International Nuclear Information System (INIS)

    Gunn, S.V.

    1989-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2000-09-01

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

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

    Science.gov (United States)

    Rønningen, Jan-Erik; Kubberud, Nils

    2005-08-01

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

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

    Science.gov (United States)

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

    1971-01-01

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

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

    Science.gov (United States)

    Martin, Heath T.

    2015-01-01

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

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

    Science.gov (United States)

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

    2001-01-01

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

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

    Science.gov (United States)

    McDonald, Brian A.

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

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

    Science.gov (United States)

    Wahid, Mastura Ab; Ali, Wan Khairuddin Wan

    2010-06-01

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

  13. Metallized solid rocket propellants based on AN/AP and PSAN/AP for access to space

    Science.gov (United States)

    Levi, S.; Signoriello, D.; Gabardi, A.; Molinari, M.; Galfetti, L.; Deluca, L. T.; Cianfanelli, S.; Klyakin, G. F.

    2009-09-01

    Solid rocket propellants based on dual mixes of inorganic crystalline oxidizers (ammonium nitrate (AN) and ammonium perchlorate (AP)) with binder and a mixture of micrometric-nanometric aluminum were investigated. Ammonium nitrate is a low-cost oxidizer, producing environment friendly combustion products but with lower specific impulse compared to AP. The better performance obtained with AP and the low quantity of toxic emissions obtained by using AN have suggested an interesting compromise based on a dual mixture of the two oxidizers. To improve the thermal response of raw AN, different types of phase stabilized AN (PSAN) and AN/AP co-crystals were investigated.

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

    Science.gov (United States)

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

    2014-05-01

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

  15. Hybrids - Best of both worlds. [liquid and solid propellants mated for safe reliable and low cost launch vehicles

    Science.gov (United States)

    Goldberg, Ben E.; Wiley, Dan R.

    1991-01-01

    An overview is presented of hybrid rocket propulsion systems whereby combining solids and liquids for launch vehicles could produce a safe, reliable, and low-cost product. The primary subsystems of a hybrid system consist of the oxidizer tank and feed system, an injector system, a solid fuel grain enclosed in a pressure vessel case, a mixing chamber, and a nozzle. The hybrid rocket has an inert grain, which reduces costs of development, transportation, manufacturing, and launch by avoiding many safety measures that must be taken when operating with solids. Other than their use in launch vehicles, hybrids are excellent for simulating the exhaust of solid rocket motors for material development.

  16. Mean Flow Augmented Acoustics in Rocket Systems

    Science.gov (United States)

    Fischbach, Sean R.

    2014-01-01

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

  17. Nozzle seal

    International Nuclear Information System (INIS)

    Herman, R.F.

    1977-01-01

    In an illustrative embodiment of the invention, a nuclear reactor pressure vessel, having an internal hoop from which the heated coolant emerges from the reactor core and passes through to the reactor outlet nozzles, is provided with sealing members operatively disposed between the outlet nozzle and the hoop. The sealing members are biased against the pressure vessel and the hoop and are connected by a leak restraining member establishing a leak-proof condition between the inlet and outlet coolants in the region about the outlet nozzle. Furthermore, the flexible responsiveness of the seal assures that the seal will not structurally couple the hoop to the pressure vessel

  18. Nozzle seal

    International Nuclear Information System (INIS)

    Walling, G.A.

    1977-01-01

    In an illustrative embodiment of the invention, a nuclear reactor pressure vessel, having an internal hoop from which the heated coolant emerges from the reactor core and passes through to the reactor outlet nozzles, is provided with sealing rings operatively disposed between the outlet nozzles and the hoop. The sealing rings connected by flexible members are biased against the pressure vessel and the hoop, establishing a leak-proof condition between the inlet and outlet coolants in the region about the outlet nozzle. Furthermore, the flexible responsiveness of the seal assures that the seal will not structurally couple the hoop to the pressure vessel. 4 claims, 2 figures

  19. RSRM Nozzle-to-Case Joint J-leg Development

    Science.gov (United States)

    Albrechtsen, Kevin U.; Eddy, Norman F.; Ewing, Mark E.; McGuire, John R.

    2003-01-01

    Since the beginning of the Space Shuttle Reusable Solid Rocket Motor (RSRM) program, nozzle-to-case joint polysulfide adhesive gas paths have occurred on several flight motors. These gas paths have allowed hot motor gases to reach the wiper O-ring. Even though these motors continue to fly safely with this condition, a desire was to reduce such occurrences. The RSRM currently uses a J-leg joint configuration on case field joints and igniter inner and outer joints. The J-leg joint configuration has been successfully demonstrated on numerous RSRM flight and static test motors, eliminating hot gas intrusion to the critical O-ring seals on these joints. Using the proven technology demonstrated on the case field joints and igniter joints, a nozzle-to-case joint J-leg design was developed for implementation on RSRM flight motors. This configuration provides an interference fit with nozzle fixed housing phenolics at assembly, with a series of pressurization gaps incorporated outboard of the joint mating surface to aid in joint pressurization and to eliminate any circumferential flow in this region. The joint insulation is bonded to the nozzle phenolics using the same pressure sensitive adhesive used in the case field joints and igniter joints. An enhancement to the nozzle-to-case joint J-leg configuration is the implementation of a carbon rope thermal barrier. The thermal barrier is located downstream of the joint bondline and is positioned within the joint in a manner where any hot gas intrusion into the joint passes through the thermal barrier, reducing gas temperatures to a level that would not affect O-rings downstream of the thermal barrier. This paper discusses the processes used in reaching a final nozzle-to-case joint J-leg design, provides structural and thermal results in support of the design, and identifies fabrication techniques and demonstrations used in arriving at the final configuration.

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

    Science.gov (United States)

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

    2017-06-01

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

  1. Inverse estimation of heat flux and temperature on nozzle throat-insert inner contour

    Energy Technology Data Exchange (ETDEWEB)

    Chen, Tsung-Chien [Department of Power Vehicle and Systems Engineering, Chung Cheng Institute of Technology, National Defense University, Ta-Hsi, Tao-Yuan 33509 (China); Liu, Chiun-Chien [Chung Shan Institute of Science and Technology, Lung-Tan, Tao-Yuan 32526 (China)

    2008-07-01

    During the missile flight, the jet flow with high temperature comes from the heat flux of propellant burning. An enormous heat flux from the nozzle throat-insert inner contour conducted into the nozzle shell will degrade the material strength of nozzle shell and reduce the nozzle thrust efficiency. In this paper, an on-line inverse method based on the input estimation method combined with the finite-element scheme is proposed to inversely estimate the unknown heat flux on the nozzle throat-insert inner contour and the inner wall temperature by applying the temperature measurements of the nozzle throat-insert. The finite-element scheme can easily define the irregularly shaped boundary. The superior capability of the proposed method is demonstrated in two major time-varying estimation cases. The computational results show that the proposed method has good estimation performance and highly facilitates the practical implementation. An effective analytical method can be offered to increase the operation reliability and thermal-resistance layer design in the solid rocket motor. (author)

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

    Science.gov (United States)

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

    1987-01-01

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

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

    International Nuclear Information System (INIS)

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

    2003-01-01

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

  4. Computational study of performance characteristics for truncated conical aerospike nozzles

    Science.gov (United States)

    Nair, Prasanth P.; Suryan, Abhilash; Kim, Heuy Dong

    2017-12-01

    Aerospike nozzles are advanced rocket nozzles that can maintain its aerodynamic efficiency over a wide range of altitudes. It belongs to class of altitude compensating nozzles. A vehicle with an aerospike nozzle uses less fuel at low altitudes due to its altitude adaptability, where most missions have the greatest need for thrust. Aerospike nozzles are better suited to Single Stage to Orbit (SSTO) missions compared to conventional nozzles. In the current study, the flow through 20% and 40% aerospike nozzle is analyzed in detail using computational fluid dynamics technique. Steady state analysis with implicit formulation is carried out. Reynolds averaged Navier-Stokes equations are solved with the Spalart-Allmaras turbulence model. The results are compared with experimental results from previous work. The transition from open wake to closed wake happens in lower Nozzle Pressure Ratio for 20% as compared to 40% aerospike nozzle.

  5. Interpretation of Core Length in Shear Coaxial Rocket Injectors from X-ray Radiography Measurements

    Science.gov (United States)

    2014-06-01

    interrogating the near field of a number of dense sprays including diesel injectors , aerated liquid jets, solid-cone sprays, impinging-jet sprays and gas...Measurements of Mass Distributions in the Near- Nozzle Region of Sprays form Standard Multi-hole Common-rail Diesel Injection Systems,” 11th Triennial...Shear Coaxial Rocket Injectors from X-ray Radiography Measurements 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) 5d. PROJECT

  6. Advanced Flow Analysis Tools for Transient Solid Rocket Motor Simulations, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The challenges of designing, developing, and fielding man-rated propulsion systems continue to increase as NASA's mission moves forward with evolving solid...

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

    Science.gov (United States)

    Ellis, Russell A.; Berdoyes, Michel

    2002-07-01

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

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

    Science.gov (United States)

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

    2018-05-01

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

  9. The development of a solid-state hydrogen sensor for rocket engine leakage detection

    Science.gov (United States)

    Liu, Chung-Chiun

    1994-01-01

    Hydrogen propellant leakage poses significant operational problems in the rocket propulsion industry as well as for space exploratory applications. Vigorous efforts have been devoted to minimizing hydrogen leakage in assembly, test, and launch operations related to hydrogen propellant. The objective has been to reduce the operational cost of assembling and maintaining hydrogen delivery systems. Specifically, efforts have been made to develop a hydrogen leak detection system for point-contact measurement. Under the auspices of Lewis Research Center, the Electronics Design Center at Case Western Reserve University, Cleveland, Ohio, has undertaken the development of a point-contact hydrogen gas sensor with potential applications to the hydrogen propellant industry. We envision a sensor array consisting of numbers of discrete hydrogen sensors that can be located in potential leak sites. Silicon-based microfabrication and micromachining techniques are used in the fabrication of these sensor prototypes. Evaluations of the sensor are carried out in-house at Case Western Reserve University as well as at Lewis Research Center and GenCorp Aerojet, Sacramento, California. The hydrogen gas sensor is not only applicable in a hydrogen propulsion system, but also usable in many other civilian and industrial settings. This includes vehicles or facility use, or in the production of hydrogen gas. Dual space and commercial uses of these point-contacted hydrogen sensors are feasible and will directly meet the needs and objectives of NASA as well as various industrial segments.

  10. The development of a solid-state hydrogen sensor for rocket engine leakage detection

    Science.gov (United States)

    Liu, Chung-Chiun

    Hydrogen propellant leakage poses significant operational problems in the rocket propulsion industry as well as for space exploratory applications. Vigorous efforts have been devoted to minimizing hydrogen leakage in assembly, test, and launch operations related to hydrogen propellant. The objective has been to reduce the operational cost of assembling and maintaining hydrogen delivery systems. Specifically, efforts have been made to develop a hydrogen leak detection system for point-contact measurement. Under the auspices of Lewis Research Center, the Electronics Design Center at Case Western Reserve University, Cleveland, Ohio, has undertaken the development of a point-contact hydrogen gas sensor with potential applications to the hydrogen propellant industry. We envision a sensor array consisting of numbers of discrete hydrogen sensors that can be located in potential leak sites. Silicon-based microfabrication and micromachining techniques are used in the fabrication of these sensor prototypes. Evaluations of the sensor are carried out in-house at Case Western Reserve University as well as at Lewis Research Center and GenCorp Aerojet, Sacramento, California. The hydrogen gas sensor is not only applicable in a hydrogen propulsion system, but also usable in many other civilian and industrial settings. This includes vehicles or facility use, or in the production of hydrogen gas. Dual space and commercial uses of these point-contacted hydrogen sensors are feasible and will directly meet the needs and objectives of NASA as well as various industrial segments.

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

    OpenAIRE

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

    1983-01-01

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

  12. Common Cause Case Study: An Estimated Probability of Four Solid Rocket Booster Hold-Down Post Stud Hang-ups

    Science.gov (United States)

    Cross, Robert

    2005-01-01

    Until Solid Rocket Motor ignition, the Space Shuttle is mated to the Mobil Launch Platform in part via eight (8) Solid Rocket Booster (SRB) hold-down bolts. The bolts are fractured using redundant pyrotechnics, and are designed to drop through a hold-down post on the Mobile Launch Platform before the Space Shuttle begins movement. The Space Shuttle program has experienced numerous failures where a bolt has hung up. That is, it did not clear the hold-down post before liftoff and was caught by the SRBs. This places an additional structural load on the vehicle that was not included in the original certification requirements. The Space Shuttle is currently being certified to withstand the loads induced by up to three (3) of eight (8) SRB hold-down experiencing a "hang-up". The results of loads analyses performed for (4) stud hang-ups indicate that the internal vehicle loads exceed current structural certification limits at several locations. To determine the risk to the vehicle from four (4) stud hang-ups, the likelihood of the scenario occurring must first be evaluated. Prior to the analysis discussed in this paper, the likelihood of occurrence had been estimated assuming that the stud hang-ups were completely independent events. That is, it was assumed that no common causes or factors existed between the individual stud hang-up events. A review of the data associated with the hang-up events, showed that a common factor (timing skew) was present. This paper summarizes a revised likelihood evaluation performed for the four (4) stud hang-ups case considering that there are common factors associated with the stud hang-ups. The results show that explicitly (i.e. not using standard common cause methodologies such as beta factor or Multiple Greek Letter modeling) taking into account the common factor of timing skew results in an increase in the estimated likelihood of four (4) stud hang-ups of an order of magnitude over the independent failure case.

  13. Development of the Astrobee F sounding rocket system.

    Science.gov (United States)

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

    1973-01-01

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

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

    Science.gov (United States)

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

    2016-11-05

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

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

    Energy Technology Data Exchange (ETDEWEB)

    Terry, Brandon C., E-mail: terry13@purdue.edu [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Sippel, Travis R. [Department of Mechanical Engineering, Iowa State University, 2025 Black Engineering, Ames, IA 50011 (United States); Pfeil, Mark A. [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Gunduz, I.Emre; Son, Steven F. [School of Mechanical Engineering, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States)

    2016-11-05

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

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

    International Nuclear Information System (INIS)

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

    2016-01-01

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

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

    Science.gov (United States)

    Wang, Qun-Zhen

    2003-01-01

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

  18. A six degree-of-freedom thrust sensor for a labscale hybrid rocket

    International Nuclear Information System (INIS)

    Wright, Ann M; Born, Traig; Strickland, Ryan; Wright, Andrew B

    2013-01-01

    A six degree-of-freedom thrust sensor was designed, constructed, calibrated, and tested using the labscale hybrid rocket at the University of Arkansas at Little Rock. The system consisted of six independent legs: one parallel to the axis of symmetry of the rocket for main thrust measurement, two vertical legs near the nozzle end of the rocket, one vertical leg near the oxygen input end of the rocket, and two separated horizontal legs near the nozzle end. Each leg was composed of a rotational bearing, a load cell, and a universal joint above and below the load cell. The leg was designed to create point contact along only one direction and minimize the non-axial forces applied to the load cell. With this system, force in each direction and moments for roll, pitch, and yaw can be measured. The system was calibrated and tested using a labscale hybrid rocket using gaseous oxygen and hydroxyl-terminated polybutadiene solid fuel. The thrust stand proved to be stable during calibration tests. Thrust force vector components and roll, pitch, and yaw moments were calculated for test firings with an oxygen mass flow rate range of 0.0174–0.0348 kg s −1 . (paper)

  19. A six degree-of-freedom thrust sensor for a labscale hybrid rocket

    Science.gov (United States)

    Wright, Ann M.; Wright, Andrew B.; Born, Traig; Strickland, Ryan

    2013-12-01

    A six degree-of-freedom thrust sensor was designed, constructed, calibrated, and tested using the labscale hybrid rocket at the University of Arkansas at Little Rock. The system consisted of six independent legs: one parallel to the axis of symmetry of the rocket for main thrust measurement, two vertical legs near the nozzle end of the rocket, one vertical leg near the oxygen input end of the rocket, and two separated horizontal legs near the nozzle end. Each leg was composed of a rotational bearing, a load cell, and a universal joint above and below the load cell. The leg was designed to create point contact along only one direction and minimize the non-axial forces applied to the load cell. With this system, force in each direction and moments for roll, pitch, and yaw can be measured. The system was calibrated and tested using a labscale hybrid rocket using gaseous oxygen and hydroxyl-terminated polybutadiene solid fuel. The thrust stand proved to be stable during calibration tests. Thrust force vector components and roll, pitch, and yaw moments were calculated for test firings with an oxygen mass flow rate range of 0.0174-0.0348 kg s-1.

  20. Design and analysis approach for linear aerospike nozzle

    International Nuclear Information System (INIS)

    Khan, S.U.; Khan, A.A.; Munir, A.

    2014-01-01

    The paper presents an aerodynamic design of a simplified linear aerospike nozzle and its detailed exhaust flow analysis with no spike truncation. Analytical method with isentropic planar flow was used to generate the nozzle contour through MATLAB . The developed code produces a number of outputs comprising nozzle wall profile, flow properties along the nozzle wall, thrust coefficient, thrust, as well as amount of nozzle truncation. Results acquired from design code and numerical analyses are compared for observing differences. The numerical analysis adopted an inviscid model carried out through commercially available and reliable computational fluid dynamics (CFD) software. Use of the developed code would assist the readers to perform quick analysis of different aerodynamic design parameters for the aerospike nozzle that has tremendous scope of application in future launch vehicles. Keyword: Rocket propulsion, Aerospike Nozzle, Control Design, Computational Fluid Dynamics. (author)

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

    Science.gov (United States)

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

    2017-12-01

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

  2. Fuel nozzle assembly

    Science.gov (United States)

    Johnson, Thomas Edward [Greer, SC; Ziminsky, Willy Steve [Simpsonville, SC; Lacey, Benjamin Paul [Greer, SC; York, William David [Greer, SC; Stevenson, Christian Xavier [Inman, SC

    2011-08-30

    A fuel nozzle assembly is provided. The assembly includes an outer nozzle body having a first end and a second end and at least one inner nozzle tube having a first end and a second end. One of the nozzle body or nozzle tube includes a fuel plenum and a fuel passage extending therefrom, while the other of the nozzle body or nozzle tube includes a fuel injection hole slidably aligned with the fuel passage to form a fuel flow path therebetween at an interface between the body and the tube. The nozzle body and the nozzle tube are fixed against relative movement at the first ends of the nozzle body and nozzle tube, enabling the fuel flow path to close at the interface due to thermal growth after a flame enters the nozzle tube.

  3. HPLC Characterization of Phenol-Formaldehyde Resole Resin Used in Fabrication of Shuttle Booster Nozzles

    Science.gov (United States)

    Young, Philip R.

    1999-01-01

    A reverse phase High Performance Liquid Chromatographic method was developed to rapidly fingerprint a phenol-formaldehyde resole resin similar to Durite(R) SC-1008. This resin is used in the fabrication of carbon-carbon composite materials from which Space Shuttle Solid Rocket Booster nozzles are manufactured. A knowledge of resin chemistry is essential to successful composite processing and performance. The results indicate that a high quality separation of over 35 peaks in 25 minutes were obtained using a 15 cm Phenomenex LUNA C8 bonded reverse phase column, a three-way water-acetonitrile-methanol nonlinear gradient, and LTV detection at 280 nm.

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

    Science.gov (United States)

    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

  5. Laval nozzles for cluster-jet targets

    Energy Technology Data Exchange (ETDEWEB)

    Grieser, Silke; Bonaventura, Daniel; Hergemoeller, Ann-Katrin; Hetz, Benjamin; Koehler, Esperanza; Lessmann, Lukas; Khoukaz, Alfons [Institut fuer Kernphysik, Westfaelische Wilhelms-Universitaet Muenster, 48149 Muenster (Germany)

    2016-07-01

    Cluster-jet targets are highly suited for storage ring experiments due to the fact that they provide high and constant beam densities. Therefore, a cluster-jet target is planned to be the first internal target for the PANDA experiment at FAIR. A cluster source generates a continuous flow of cryogenic solid clusters by the expansion of pre-cooled gases within fine Laval nozzles. For the production of clusters the geometry of the nozzle is crucial. The production of such nozzles with their complex inner geometry represents a major technical challenge. The possibility to produce new fine Laval nozzles ensures the operation of cluster-jet targets, e.g. for the PANDA experiment, and opens the way for future investigations on the cluster production process to match the required targets performance. Optimizations on the recently developed production process and the fabrication of new glass nozzles were done. Initial measurements of these nozzles at the PANDA cluster-jet target prototype and the investigation of the cluster beam origin within the nozzle will be presented and discussed. For the future more Laval nozzles with different geometries will be produced and additional measurements with these new nozzles at the PANDA cluster-jet target prototype towards higher performance will be realized.

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

    Science.gov (United States)

    Schramm, Harry F.; Sullivan, Kenneth W.

    1991-01-01

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

  7. Manufacturing Advanced Channel Wall Rocket Liners, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This SBIR will adapt and demonstrate a low cost flexible method of manufacturing channel wall liquid rocket nozzles and combustors, while providing developers a...

  8. Nozzle airfoil having movable nozzle ribs

    Science.gov (United States)

    Yu, Yufeng Phillip; Itzel, Gary Michael

    2002-01-01

    A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

  9. AFRPL Graphite Performance Prediction Program. Improved Capability for the Design and Ablation Performance Prediction of Advanced Air Force Solid Propellant Rocket Nozzles

    Science.gov (United States)

    1976-12-01

    the molecular weight of molecules are shown, it is not always possible to unequivically identify the molecule. For instance, a molecular weight at 28...tabular form in Table A-7. Also shown in this table are possible chemical species that could be represented by each molecular weight. Only molecular weights...pyrolytic graphite is 2.2 gr/cc. The heat of formation is assumed to be zero (the value for elemental carbon). The surface emisivity (total

  10. Rocket Flight.

    Science.gov (United States)

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

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

  11. Program ELM: A tool for rapid thermal-hydraulic analysis of solid-core nuclear rocket fuel elements

    International Nuclear Information System (INIS)

    Walton, J.T.

    1992-11-01

    This report reviews the state of the art of thermal-hydraulic analysis codes and presents a new code, Program ELM, for analysis of fuel elements. ELM is a concise computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in a nuclear thermal rocket reactor with axial coolant passages. The program was developed as a tool to swiftly evaluate various heat transfer coefficient and friction factor correlations generated for turbulent pipe flow with heat addition which have been used in previous programs. Thus, a consistent comparison of these correlations was performed, as well as a comparison with data from the NRX reactor experiments from the Nuclear Engine for Rocket Vehicle Applications (NERVA) project. This report describes the ELM Program algorithm, input/output, and validation efforts and provides a listing of the code

  12. Rocket science

    International Nuclear Information System (INIS)

    Upson Sandra

    2011-01-01

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

  13. Unique nuclear thermal rocket engine

    International Nuclear Information System (INIS)

    Culver, D.W.; Rochow, R.

    1993-06-01

    In January, 1992, a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars was introduced (Culver, 1992). This NTRE promises to be both shorter and lighter in weight than conventionally designed engines, because its forward flowing reactor is located within an expansion-deflection rocket nozzle. The concept has matured during the year, and this paper discusses a nearer term version that resolves four open issues identified in the initial concept: (1) the reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2) elimination need for a new, uncooled nozzle throat material suitable for long life application; (3) a practical provision for reactor power control; and (4) use of near-term, long-life turbopumps

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

    Science.gov (United States)

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

    2017-06-01

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

  15. Cold spray nozzle design

    Science.gov (United States)

    Haynes, Jeffrey D [Stuart, FL; Sanders, Stuart A [Palm Beach Gardens, FL

    2009-06-09

    A nozzle for use in a cold spray technique is described. The nozzle has a passageway for spraying a powder material, the passageway having a converging section and a diverging section, and at least the diverging section being formed from polybenzimidazole. In one embodiment of the nozzle, the converging section is also formed from polybenzimidazole.

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

    Science.gov (United States)

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

    2012-01-01

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

  17. Cold Flow Determination of the Internal Flow Environment Around the Submerged TVC Nozzle for the Space Shuttle SRM

    Science.gov (United States)

    Whitesides, R. H.; Ghosh, A.; Jenkins, S. L.; Bacchus, D. L.

    1989-01-01

    A series of subscale cold flow tests was performed to quantify the gas flow characteristics at the aft end of the Space Shuttle Solid Rocket Motor. This information was used to support the analyses of the redesigned nozzle/case joint. A portion of the thermal loads at the joint are due to the circumferential velocities and pressure gradients caused primarily by the gimbaling of the submerged nose TVC nozzle. When the nozzle centerline is vectored with respect to the motor centerline, asymmetries are set up in the flow field under the submerged nozzle and immediately adjacent to the nozzle/case joint. Specific program objectives included: determination of the effects of nozzle gimbal angle and propellant geometry on the circumferential flow field; measurement of the static pressure and gas velocities in the vicinity of the nozzle/case joint; use of scaling laws to apply the subscale cold flow data to the full scale SRM; and generation of data for use in validation of 3-D computational fluid dynamic, CFD, models of the SRM flow field. These tests were conducted in the NASA Marshall Space Flight Center Airflow Facility with a 7.5 percent scale model of the aft segment of the SRM. Static and dynamic pressures were measured in the model to quantify the flow field. Oil flow data was also acquired to obtain qualitative visual descriptions of the flow field. Nozzle gimbal angles of 0, 3.5, and 7 deg were used with propellant grain configurations corresponding to motor burn times of 0, 9, 19, and 114 seconds. This experimental program was successful in generating velocity and pressure gradient data for the flow field around the submerged nose nozzle of the Space Shuttle SRM at various burn times and gimbal angles. The nature of the flow field adjacent to the nozzle/case joint was determined with oil droplet streaks, and the velocity and pressure gradients were quantified with pitot probes and wall static pressure measurements. The data was applied to the full scale SRM thru

  18. Theodore von Karman - Rocket Scientist

    Indian Academy of Sciences (India)

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

  19. Feasibility of using neutron radiography to inspect the Space Shuttle solid rocket booster aft skirt, forward skirt and frustum. Part 1: Summary report

    Science.gov (United States)

    Barton, J. P.; Bader, J. W.; Brenizer, J. S.; Hosticka, B.

    1992-01-01

    The space shuttle's solid rocket boosters (SRB) include components made primarily of aluminum that are parachuted back for retrieval from the ocean and refurbished for repeated usage. Nondestructive inspection methods used on these aging parts to reduce the risk of unforeseen problems include x-ray, ultrasonics, and eddy current. Neutron radiography tests on segments of an SRB component show that entrapped moisture and naturally occurring aluminum corrosion can be revealed by neutron radiography even if present in only small amounts. Voids in sealant can also be evaluated. Three alternatives are suggested to follow-up this study: (1) take an SRB component to an existing neutron radiography system; (2) take an existing mobile neutron radiography system to the NASA site; or (3) plan a dedicated system custom designed for NASA applications.

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

    Science.gov (United States)

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

    2011-09-01

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

  1. Aeroelastic Modeling of a Nozzle Startup Transient

    Science.gov (United States)

    Wang, Ten-See; Zhao, Xiang; Zhang, Sijun; Chen, Yen-Sen

    2014-01-01

    Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development during test. While three-dimensional, transient, turbulent, chemically reacting computational fluid dynamics methodology has been demonstrated to capture major side load physics with rigid nozzles, hot-fire tests often show nozzle structure deformation during major side load events, leading to structural damages if structural strengthening measures were not taken. The modeling picture is incomplete without the capability to address the two-way responses between the structure and fluid. The objective of this study is to develop a tightly coupled aeroelastic modeling algorithm by implementing the necessary structural dynamics component into an anchored computational fluid dynamics methodology. The computational fluid dynamics component is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, while the computational structural dynamics component is developed under the framework of modal analysis. Transient aeroelastic nozzle startup analyses at sea level were performed, and the computed transient nozzle fluid-structure interaction physics presented,

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

    Science.gov (United States)

    Zibit, Benjamin Seth

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

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

    Science.gov (United States)

    Ramsey, P. E.

    1976-01-01

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

  4. Advanced exhaust nozzle technology

    Energy Technology Data Exchange (ETDEWEB)

    Glidewell, R J; Warburton, R E

    1981-01-01

    Recent developments in turbine engine exhaust nozzle technology include nonaxisymmetric nozzles, thrust reversing, and thrust vectoring. Trade studies have been performed to determine the impact of these developments on the thrust-to-weight ratio and specific fuel consumption of an advanced high performance, augmented turbofan engine. Results are presented in a manner which provides an understanding of the sources and magnitudes of differences in the basic elements of nozzle internal performance and weight as they relate to conventional, axisymmetric nozzle technology. Conclusions are presented and recommendations are made with regard to future directions of advanced development and demonstration. 5 refs.

  5. Firefighter Nozzle Reaction

    DEFF Research Database (Denmark)

    Chin, Selena K.; Sunderland, Peter B.; Jomaas, Grunde

    2017-01-01

    to anchor forces, the hose becomes straight. The nozzle reaction is found to equal the jet momentum flow rate, and it does not change when an elbow connects the hose to the nozzle. A forward force must be exerted by a firefighter or another anchor that matches the forward force that the jet would exert...... on a perpendicular wall. Three reaction expressions are derived, allowing it to be determined in terms of hose diameter, jet diameter, flow rate, and static pressure upstream of the nozzle. The nozzle reaction predictions used by the fire service are 56% to 90% of those obtained here for typical firefighting hand...

  6. Hybrid rocket engine research program at Ryerson University

    Energy Technology Data Exchange (ETDEWEB)

    Karpynczyk, J.; Greatrix, D.R. [Ryerson Polytechnic Univ., Toronto, ON (Canada). Dept. of Aerospace Engineering

    2007-07-01

    Hybrid rocket engines (HREs) are a combination of solid and liquid propellant rocket engine designs. A solid fuel grain is located in the main combustion chamber and nozzle aft, while a stored liquid or gaseous oxidizer source supplies the required oxygen content through a throttle valve, for combustion downstream in the main chamber. HREs have drawn significant interest in certain flight applications, as they can be advantageous in terms of cost, ease and safety in storage, controllability in flight, and availability of propellant constituents. Key factors that will lead to further practical usage of HREs for flight applications are their predictability and reproducibility of operational performance. This paper presented information on studies being conducted at Ryerson University aimed at analyzing and testing the performance of HREs. It discussed and illustrated the conventional HRE and analyzed engine performance considerations such as the fuel regression rate, mass flux about the fuel surface, burning rate, and zero transformation parameter. Other factors relating to HRE performance that were presented included induced forward and aft oxidizer flow swirl effects as a means for augmenting the fuel regression rate, stoichiometric grain length issues, and feed system stability. Last, the paper presented a simplified schematic diagram of a proposed thrust/test stand for HRE test firings. 2 refs., 3 figs.

  7. ELM - A SIMPLE TOOL FOR THERMAL-HYDRAULIC ANALYSIS OF SOLID-CORE NUCLEAR ROCKET FUEL ELEMENTS

    Science.gov (United States)

    Walton, J. T.

    1994-01-01

    ELM is a simple computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in nuclear thermal rockets. Written for the nuclear propulsion project of the Space Exploration Initiative, ELM evaluates the various heat transfer coefficient and friction factor correlations available for turbulent pipe flow with heat addition. In the past, these correlations were found in different reactor analysis codes, but now comparisons are possible within one program. The logic of ELM is based on the one-dimensional conservation of energy in combination with Newton's Law of Cooling to determine the bulk flow temperature and the wall temperature across a control volume. Since the control volume is an incremental length of tube, the corresponding pressure drop is determined by application of the Law of Conservation of Momentum. The size, speed, and accuracy of ELM make it a simple tool for use in fuel element parametric studies. ELM is a machine independent program written in FORTRAN 77. It has been successfully compiled on an IBM PC compatible running MS-DOS using Lahey FORTRAN 77, a DEC VAX series computer running VMS, and a Sun4 series computer running SunOS UNIX. ELM requires 565K of RAM under SunOS 4.1, 360K of RAM under VMS 5.4, and 406K of RAM under MS-DOS. Because this program is machine independent, no executable is provided on the distribution media. The standard distribution medium for ELM is one 5.25 inch 360K MS-DOS format diskette. ELM was developed in 1991. DEC, VAX, and VMS are trademarks of Digital Equipment Corporation. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

  8. Nuclear rockets

    International Nuclear Information System (INIS)

    Sarram, M.

    1972-01-01

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

  9. Nuclear rockets

    Energy Technology Data Exchange (ETDEWEB)

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

    1972-02-01

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

  10. Fluid flow nozzle energy harvesters

    Science.gov (United States)

    Sherrit, Stewart; Lee, Hyeong Jae; Walkemeyer, Phillip; Winn, Tyler; Tosi, Luis Phillipe; Colonius, Tim

    2015-04-01

    Power generation schemes that could be used downhole in an oil well to produce about 1 Watt average power with long-life (decades) are actively being developed. A variety of proposed energy harvesting schemes could be used to extract energy from this environment but each of these has their own limitations that limit their practical use. Since vibrating piezoelectric structures are solid state and can be driven below their fatigue limit, harvesters based on these structures are capable of operating for very long lifetimes (decades); thereby, possibly overcoming a principle limitation of existing technology based on rotating turbo-machinery. An initial survey [1] identified that spline nozzle configurations can be used to excite a vibrating piezoelectric structure in such a way as to convert the abundant flow energy into useful amounts of electrical power. This paper presents current flow energy harvesting designs and experimental results of specific spline nozzle/ bimorph design configurations which have generated suitable power per nozzle at or above well production analogous flow rates. Theoretical models for non-dimensional analysis and constitutive electromechanical model are also presented in this paper to optimize the flow harvesting system.

  11. Firefighter Nozzle Reaction

    DEFF Research Database (Denmark)

    Chin, Selena K.; Sunderland, Peter B.; Jomaas, Grunde

    2017-01-01

    Nozzle reaction and hose tension are analyzed using conservation of fluid momentum and assuming steady, inviscid flow and a flexible hose in frictionless contact with the ground. An expression that is independent of the bend angle is derived for the hose tension. If this tension is exceeded owing...... to anchor forces, the hose becomes straight. The nozzle reaction is found to equal the jet momentum flow rate, and it does not change when an elbow connects the hose to the nozzle. A forward force must be exerted by a firefighter or another anchor that matches the forward force that the jet would exert...... on a perpendicular wall. Three reaction expressions are derived, allowing it to be determined in terms of hose diameter, jet diameter, flow rate, and static pressure upstream of the nozzle. The nozzle reaction predictions used by the fire service are 56% to 90% of those obtained here for typical firefighting hand...

  12. Rocket observations

    Science.gov (United States)

    1984-05-01

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

  13. Metal Digital Direct Manufacturing (MDDM) for Close-Out of Combustion Chambers and Nozzle Fabrications, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — This NASA sponsored STTR project will investigate methods for close-out of large, liquid rocket engine, nickel or stainless steel nozzle, coolant channels utilizing...

  14. Multi-Stage Hybrid Rocket Conceptual Design for Micro-Satellites Launch using Genetic Algorithm

    Science.gov (United States)

    Kitagawa, Yosuke; Kitagawa, Koki; Nakamiya, Masaki; Kanazaki, Masahiro; Shimada, Toru

    The multi-objective genetic algorithm (MOGA) is applied to the multi-disciplinary conceptual design problem for a three-stage launch vehicle (LV) with a hybrid rocket engine (HRE). MOGA is an optimization tool used for multi-objective problems. The parallel coordinate plot (PCP), which is a data mining method, is employed in the post-process in MOGA for design knowledge discovery. A rocket that can deliver observing micro-satellites to the sun-synchronous orbit (SSO) is designed. It consists of an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank and a nozzle. The objective functions considered in this study are to minimize the total mass of the rocket and to maximize the ratio of the payload mass to the total mass. To calculate the thrust and the engine size, the regression rate is estimated based on an empirical model for a paraffin (FT-0070) propellant. Several non-dominated solutions are obtained using MOGA, and design knowledge is discovered for the present hybrid rocket design problem using a PCP analysis. As a result, substantial knowledge on the design of an LV with an HRE is obtained for use in space transportation.

  15. MHD thrust vectoring of a rocket engine

    Science.gov (United States)

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

    2016-09-01

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

  16. Fundamentals of aircraft and rocket propulsion

    CERN Document Server

    El-Sayed, Ahmed F

    2016-01-01

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

  17. Air-Powered Rockets.

    Science.gov (United States)

    Rodriguez, Charley; Raynovic, Jim

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

  18. Thermohydraulic modeling of nuclear thermal rockets: The KLAXON code

    International Nuclear Information System (INIS)

    Hall, M.L.; Rider, W.J.; Cappiello, M.W.

    1992-01-01

    The hydrogen flow from the storage tanks, through the reactor core, and out the nozzle of a Nuclear Thermal Rocket is an integral design consideration. To provide an analysis and design tool for this phenomenon, the KLAXON code is being developed. A shock-capturing numerical methodology is used to model the gas flow (the Harten, Lax, and van Leer method, as implemented by Einfeldt). Preliminary results of modeling the flow through the reactor core and nozzle are given in this paper

  19. Cold water injection nozzles

    International Nuclear Information System (INIS)

    Kura, Masaaki; Maeda, Masamitsu; Endo, Takio.

    1979-01-01

    Purpose: To inject cold water in a reactor without applying heat cycles to a reactor container and to the inner wall of a feedwater nozzle by securing a perforated plate at the outlet of the cold water injection nozzle. Constitution: A disc-like cap is secured to the final end of a return nozzle of a control rod drive. The cap prevents the flow of a high temperature water flowing downward in the reactor from entering into the nozzle. The cap is perforated with a plurality of bore holes for injecting cold water into the reactor. The cap is made to about 100 mm in thickness so that the cold water passing through the bore holes is heated by the heat conduction in the cap. Accordingly, the flow of high temperature water flowing downwardly in the reactor is inhibited by the cap from backward flowing into the nozzle. Moreover, the flow of the cold water in the nozzle is controlled and rectified when passed through the bore holes in the cap and then injected into the reactor. (Yoshino, Y.)

  20. Structural Evaluation of the RSRM Nozzle Replacement Adhesive

    Science.gov (United States)

    Batista-Rodriguez, A.; McLennan, M. L.; Palumbos, A. V.; Richardson, D. E.

    1999-01-01

    This paper describes the structural performance evaluation of a replacement adhesive for the Reusable Solid Rocket Motor (RSRM) nozzle utilizing finite element analysis. Due to material obsolescence and industrial safety issues, the two current structural adhesives, EA 913 and EA 946 are to be replaced with a new adhesive. TIGA 321. The structural evaluation in support of the adhesive replacement effort includes residual stress, transportation, and flight analyses. Factors of safety are calculated using the stress response from each analysis. The factors of safety are used as the limiting criteria to compare the replacement adhesive against the current adhesives. Included in this paper are the analytical approach, assumptions and modeling techniques as well as the results of the evaluation. An important factor to the evaluation is the similarity in constitutive material properties (elastic modulus and Poisson's ratio) between TIGA 321 and EA 913. This similarity leads to equivalent material response from the two adhesives. However, TIGA 321 surpasses EA 913's performance due to higher material capabilities. Conversely, the change in stress response from EA 946 to TIGA 321 is more apparent: this is primarily attributed to the difference in the modulii of the two adhesives, which differ by two orders of magnitude. The results of the bondline evaluation indicate that the replacement adhesive provides superior performance than the current adhesives with only minor exceptions. Furthermore, TIGA 321 causes only a minor chance in the response of the phenolic and metal components.

  1. Hybrid rocket propulsion systems for outer planet exploration missions

    Science.gov (United States)

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

    2016-11-01

    Outer planet exploration missions require significant propulsive capability, particularly to achieve orbit insertion. Missions to explore the moons of outer planets place even more demanding requirements on propulsion systems, since they involve multiple large ΔV maneuvers. Hybrid rockets present a favorable alternative to conventional propulsion systems for many of these missions. They typically enjoy higher specific impulse than solids, can be throttled, stopped/restarted, and have more flexibility in their packaging configuration. Hybrids are more compact and easier to throttle than liquids and have similar performance levels. In order to investigate the suitability of these propulsion systems for exploration missions, this paper presents novel hybrid motor designs for two interplanetary missions. Hybrid propulsion systems for missions to Europa and Uranus are presented and compared to conventional in-space propulsion systems. The hybrid motor design for each of these missions is optimized across a range of parameters, including propellant selection, O/F ratio, nozzle area ratio, and chamber pressure. Details of the design process are described in order to provide guidance for researchers wishing to evaluate hybrid rocket motor designs for other missions and applications.

  2. Sounding rockets explore the ionosphere

    International Nuclear Information System (INIS)

    Mendillo, M.

    1990-01-01

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

  3. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

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

  4. Characterisation of subsonic axisymmetric nozzles

    Czech Academy of Sciences Publication Activity Database

    Tesař, Václav

    2008-01-01

    Roč. 86, č. 11 (2008), s. 1253-1262 ISSN 0263-8762 R&D Projects: GA AV ČR IAA200760705 Institutional research plan: CEZ:AV0Z20760514 Keywords : nozzle * characterisation * nozzle properties * nozzle invariants Subject RIV: BK - Fluid Dynamics Impact factor: 0.989, year: 2008

  5. Duplex tab exhaust nozzle

    Science.gov (United States)

    Gutmark, Ephraim Jeff (Inventor); Martens, Steven (nmn) (Inventor)

    2012-01-01

    An exhaust nozzle includes a conical duct terminating in an annular outlet. A row of vortex generating duplex tabs are mounted in the outlet. The tabs have compound radial and circumferential aft inclination inside the outlet for generating streamwise vortices for attenuating exhaust noise while reducing performance loss.

  6. Numerical Simulation of Reactive Flows in Overexpanded Supersonic Nozzle with Film Cooling

    Directory of Open Access Journals (Sweden)

    Mohamed Sellam

    2015-01-01

    Full Text Available Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2 gas (GH2 in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2 film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are also analyzed in terms of cooling efficiency and the results already give an idea on their magnitude. It was also shown that slightly increasing the film injection pressure can avoid the reignition phenomena by moving the separation shock towards the nozzle exit section.

  7. Numerical Investigation of Twin-Nozzle Rocket Plume Phenomenology

    National Research Council Canada - National Science Library

    Ebrahimi, Houshang

    1998-01-01

    .... The Van Leer Flux Splitting option has been successfully implemented into the existing GIFS model and provides a more robust solution scheme, making application of the model more reasonable for engineering applications...

  8. THRUST AUGMENTED NOZZLE (TAN) the New Paradigm for Booster Rockets

    Science.gov (United States)

    2006-07-12

    station. The engine has to throttle to 34 percent (3X or 1020 psia) to keep from exceeding the acceleration limits. Figure 6. Baseline SSTO ...vehicle powered by seven up-sized SSME class engines. Figure 7. Baseline SSTO vehicle trajectory. With a payload fraction of 1 percent, it does not...want to invest in such a risky endeavor. American Institute of Aeronautics and Astronautics 6 B. TAN-Powered SSTO Vehicle For the Dual Fuel TAN

  9. Carbon Nano-Composite Ablative Rocket Nozzles, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The constantly evolving science of nanotechnology keeps coming around to old ideas re-tooled with new technologies. Though much work has been done examining the...

  10. Premixed direct injection nozzle

    Science.gov (United States)

    Zuo, Baifang [Simpsonville, SC; Johnson, Thomas Edward [Greer, SC; Lacy, Benjamin Paul [Greer, SC; Ziminsky, Willy Steve [Simpsonville, SC

    2011-02-15

    An injection nozzle having a main body portion with an outer peripheral wall is disclosed. The nozzle includes a plurality of fuel/air mixing tubes disposed within the main body portion and a fuel flow passage fluidly connected to the plurality of fuel/air mixing tubes. Fuel and air are partially premixed inside the plurality of the tubes. A second body portion, having an outer peripheral wall extending between a first end and an opposite second end, is connected to the main body portion. The partially premixed fuel and air mixture from the first body portion gets further mixed inside the second body portion. The second body portion converges from the first end toward said second end. The second body portion also includes cooling passages that extend along all the walls around the second body to provide thermal damage resistance for occasional flame flash back into the second body.

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

    Science.gov (United States)

    Davidian, Kenneth O.

    1990-01-01

    The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

  12. Limit loads in nozzles

    International Nuclear Information System (INIS)

    Zouain, N.

    1983-01-01

    The static method for the evaluation of the limit loads of a perfectly elasto-plastic structure is presented. Using the static theorem of Limit Analysis and the Finite Element Method, a lower bound for the colapso load can be obtained through a linear programming problem. This formulation if then applied to symmetrically loaded shells of revolution and some numerical results of limit loads in nozzles are also presented. (Author) [pt

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

    Science.gov (United States)

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

    2015-09-01

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

  14. Numerical study on drop formation through a micro nozzle

    International Nuclear Information System (INIS)

    Kim, Sung Il; Son, Gi Hun

    2005-01-01

    The drop ejection process from a micro nozzle is investigated by numerically solving the conservation equations for mass and momentum. The liquid-gas interface is tracked by a level set method which is extended for two-fluid flows with irregular solid boundaries. Based on the numerical results, the liquid jet breaking and droplet formation behavior is found to depend strongly on the pulse type of forcing pressure and the contact angle at the gas-liquid-solid interline. The negative pressure forcing can be used to control the formation of satelite droplets. Also, various nozzle shapes are tested to investigate their effect on droplet formation

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

    Science.gov (United States)

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

    1954-01-01

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

  16. Two stage turbine for rockets

    Science.gov (United States)

    Veres, Joseph P.

    1993-01-01

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

  17. Performance of a RBCC Engine in Rocket-Operation

    Science.gov (United States)

    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.

  18. Transient Side Load Analysis of Out-of-Round Film-Cooled Nozzle Extensions

    Science.gov (United States)

    Wang, Ten-See; Lin, Jeff; Ruf, Joe; Guidos, Mike

    2012-01-01

    There was interest in understanding the impact of out-of-round nozzle extension on the nozzle side load during transient startup operations. The out-of-round nozzle extension could be the result of asymmetric internal stresses, deformation induced by previous tests, and asymmetric loads induced by hardware attached to the nozzle. The objective of this study was therefore to computationally investigate the effect of out-of-round nozzle extension on the nozzle side loads during an engine startup transient. The rocket engine studied encompasses a regeneratively cooled chamber and nozzle, along with a film cooled nozzle extension. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and transient inlet boundary flow properties derived from an engine system simulation. Six three-dimensional cases were performed with the out-of-roundness achieved by three different degrees of ovalization, elongated on lateral y and z axes: one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation line jump was the primary source of the peak side loads. Comparing to the peak side load of the perfectly round nozzle, the peak side loads increased for the slightly and more ovalized nozzle extensions, and either increased or decreased for the two significantly ovalized nozzle extensions. A theory based on the counteraction of the flow destabilizing effect of an exacerbated asymmetrical flow caused by a lower degree of ovalization, and the flow stabilizing effect of a more symmetrical flow, created also by ovalization, is presented to explain the observations obtained in this effort.

  19. Thermal Analysis of the Fastrac Chamber/Nozzle

    Science.gov (United States)

    Davis, Darrell

    2001-01-01

    This paper will describe the thermal analysis techniques used to predict temperatures in the film-cooled ablative rocket nozzle used on the Fastrac 60K rocket engine. A model was developed that predicts char and pyrolysis depths, liner thermal gradients, and temperatures of the bondline between the overwrap and liner. Correlation of the model was accomplished by thermal analog tests performed at Southern Research, and specially instrumented hot fire tests at the Marshall Space Flight Center. Infrared thermography was instrumental in defining nozzle hot wall surface temperatures. In-depth and outboard thermocouple data was used to correlate the kinetic decomposition routine used to predict char and pyrolysis depths. These depths were anchored with measured char and pyrolysis depths from cross-sectioned hot-fire nozzles. For the X-34 flight analysis, the model includes the ablative Thermal Protection System (TPS) material that protects the overwrap from the recirculating plume. Results from model correlation, hot-fire testing, and flight predictions will be discussed.

  20. Thermal Analysis of the MC-1 Chamber/Nozzle

    Science.gov (United States)

    Davis, Darrell W.; Phelps, Lisa H. (Technical Monitor)

    2001-01-01

    This paper will describe the thermal analysis techniques used to predict temperatures in the film-cooled ablative rocket nozzle used on the MC-1 60K rocket engine. A model was developed that predicts char and pyrolysis depths, liner thermal gradients, and temperatures of the bondline between the overwrap and liner. Correlation of the model was accomplished by thermal analog tests performed at Southern Research, and specially instrumented hot fire tests at the Marshall Space Flight Center. Infrared thermography was instrumental in defining nozzle hot wall surface temperatures. In-depth and outboard thermocouple data was used to correlate the kinetic decomposition routine used to predict char and pyrolysis depths. These depths were anchored with measured char and pyrolysis depths from cross-sectioned hot-fire nozzles. For the X-34 flight analysis, the model includes the ablative Thermal Protection System (TPS) material that protects the overwrap from the recirculating plume. Results from model correlation, hot-fire testing, and flight predictions will be discussed.

  1. Reverse flow through a large scale multichannel nozzle

    International Nuclear Information System (INIS)

    Duignan, M.R.; Nash, C.A.

    1992-01-01

    A database was developed for the flow of water through a scaled nozzle of a Savannah River Site reactor inlet plenum. The water flow in the nozzle was such that it ranged from stratified to water solid conditions. Data on the entry of air into the nozzle and plenum as a function of water flow are of interest in loss-of-coolant studies. The scaled nozzle was 44 cm long, had an entrance diameter of 95 mm, an exit opening of 58 mm x 356 mm, and an exit hydraulic diameter approximately equal to that of the inlet. Within the nozzle were three flow-straightening vanes which divided the flow path into four channels. All data were taken at steady-state and isothermal (300 K ± 1.5 K) conditions. During the reverse flow of water through the nozzle the point at which air begins to enter was predicted within 90% by a critical weir-flow calculation. The point of air entry into the plenum itself was found to be a function of flow conditions

  2. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

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

  3. Ablative Material Testing at Lewis Rocket Lab

    Science.gov (United States)

    1997-01-01

    The increasing demand for a low-cost, reliable way to launch commercial payloads to low- Earth orbit has led to the need for inexpensive, expendable propulsion systems for new launch vehicles. This, in turn, has renewed interest in less complex, uncooled rocket engines that have combustion chambers and exhaust nozzles fabricated from ablative materials. A number of aerospace propulsion system manufacturers have utilized NASA Lewis Research Center's test facilities with a high degree of success to evaluate candidate materials for application to new propulsion devices.

  4. Impact and mitigation of stratospheric ozone depletion by chemical rockets

    International Nuclear Information System (INIS)

    Mcdonald, A.J.

    1992-03-01

    The American Institute of Aeronautics and Astronautics (AIAA) conducted a workshop in conjunction with the 1991 AIAA Joint Propulsion Conference in Sacramento, California, to assess the impact of chemical rocket propulsion on the environment. The workshop included recognized experts from the fields of atmospheric physics and chemistry, solid rocket propulsion, liquid rocket propulsion, government, and environmental agencies, and representatives from several responsible environmental organizations. The conclusion from this workshop relative to stratospheric ozone depletion was that neither solid nor liquid rocket launchers have a significant impact on stratospheric ozone depletion, and that there is no real significant difference between the two

  5. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) Systems Study; Volume 1 - Executive Summary

    National Research Council Canada - National Science Library

    Ware, Larry

    1989-01-01

    ...) solid rocket boosters (SRBs) with liquid rocket boosters (LRBs), Figure 1.0-1. The main objectives of a LRB substitution for the SRB were increased STS safety and reliability and increased payload performance...

  6. Thrust distribution for attitude control in a variable thrust propulsion system with four ACS nozzles

    Science.gov (United States)

    Lim, Yeerang; Lee, Wonsuk; Bang, Hyochoong; Lee, Hosung

    2017-04-01

    A thrust distribution approach is proposed in this paper for a variable thrust solid propulsion system with an attitude control system (ACS) that uses a reduced number of nozzles for a three-axis attitude maneuver. Although a conventional variable thrust solid propulsion system needs six ACS nozzles, this paper proposes a thrust system with four ACS nozzles to reduce the complexity and mass of the system. The performance of the new system was analyzed with numerical simulations, and the results show that the performance of the system with four ACS nozzles was similar to the original system while the mass of the whole system was simultaneously reduced. Moreover, a feasibility analysis was performed to determine whether a thrust system with three ACS nozzles is possible.

  7. Arcjet nozzle area ratio effects

    Science.gov (United States)

    Curran, Francis M.; Sarmiento, Charles J.; Birkner, Bjorn W.; Kwasny, James

    1990-01-01

    An experimental investigation was conducted to determine the effect of nozzle area ratio on the operating characteristics and performance of a low power dc arcjet thruster. Conical thoriated tungsten nozzle inserts were tested in a modular laboratory arcjet thruster run on hydrogen/nitrogen mixtures simulating the decomposition products of hydrazine. The converging and diverging sides of the inserts had half angles of 30 and 20 degrees, respectively, similar to a flight type unit currently under development. The length of the diverging side was varied to change the area ratio. The nozzle inserts were run over a wide range of specific power. Current, voltage, mass flow rate, and thrust were monitored to provide accurate comparisons between tests. While small differences in performance were observed between the two nozzle inserts, it was determined that for each nozzle insert, arcjet performance improved with increasing nozzle area ratio to the highest area ratio tested and that the losses become very pronounced for area ratios below 50. These trends are somewhat different than those obtained in previous experimental and analytical studies of low Re number nozzles. It appears that arcjet performance can be enhanced via area ratio optimization.

  8. Arcjet Nozzle Area Ratio Effects

    Science.gov (United States)

    Curran, Francis M.; Sarmiento, Charles J.; Birkner, Bjorn W.; Kwasny, James

    1990-01-01

    An experimental investigation was conducted to determine the effect of nozzle area ratio on the operating characteristics and performance of a low power dc arcjet thruster. Conical thoriated tungsten nozzle inserts were tested in a modular laboratory arcjet thruster run on hydrogen/nitrogen mixtures simulating the decomposition products of hydrazine. The converging and diverging sides of the inserts had half angles of 30 and 20 degrees, respectively, similar to a flight type unit currently under development. The length of the diverging side was varied to change the area ratio. The nozzle inserts were run over a wide range of specific power. Current, voltage, mass flow rate, and thrust were monitored to provide accurate comparisons between tests. While small differences in performance were observed between the two nozzle inserts, it was determined that for each nozzle insert, arcjet performance improved with increasing nozzle area ratio to the highest area ratio tested and that the losses become very pronounced for area ratios below 50. These trends are somewhat different than those obtained in previous experimental and analytical studies of low Re number nozzles. It appears that arcjet performance can be enhanced via area ratio optimization.

  9. Injection nozzle for a turbomachine

    Science.gov (United States)

    Uhm, Jong Ho; Johnson, Thomas Edward; Kim, Kwanwoo

    2012-09-11

    A turbomachine includes a compressor, a combustor operatively connected to the compressor, an end cover mounted to the combustor, and an injection nozzle assembly operatively connected to the combustor. The injection nozzle assembly includes a first end portion that extends to a second end portion, and a plurality of tube elements provided at the second end portion. Each of the plurality of tube elements defining a fluid passage includes a body having a first end section that extends to a second end section. The second end section projects beyond the second end portion of the injection nozzle assembly.

  10. Airfoil nozzle and shroud assembly

    Science.gov (United States)

    Shaffer, J.E.; Norton, P.F.

    1997-06-03

    An airfoil and nozzle assembly are disclosed including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached. 5 figs.

  11. Flow-throttling orifice nozzle

    International Nuclear Information System (INIS)

    Sletten, H.L.

    1975-01-01

    A series-parallel-flow type throttling apparatus to restrict coolant flow to certain fuel assemblies of a nuclear reactor is comprised of an axial extension nozzle of the fuel assembly. The nozzle has a series of concentric tubes with parallel-flow orifice holes in each tube. Flow passes from a high pressure plenum chamber outside the nozzle through the holes in each tube in series to the inside of the innermost tube where the coolant, having dissipated most of its pressure, flows axially to the fuel element. (U.S.)

  12. Multiple-Nozzle Spray Head Applies Foam Insulation

    Science.gov (United States)

    Walls, Joe T.

    1993-01-01

    Spray head equipped with four-nozzle turret mixes two reactive components of polyurethane and polyisocyanurate foam insulating material and sprays reacting mixture onto surface to be insulated. If nozzle in use becomes clogged, fresh one automatically rotated into position, with minimal interruption of spraying process. Incorporates features recirculating and controlling pressures of reactive components to maintain quality of foam by ensuring proper blend at outset. Also used to spray protective coats on or in ships, aircraft, and pipelines. Sprays such reactive adhesives as epoxy/polyurethane mixtures. Components of spray contain solid-particle fillers for strength, fire retardance, toughness, resistance to abrasion, or radar absorption.

  13. Rockets two classic papers

    CERN Document Server

    Goddard, Robert

    2002-01-01

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

  14. Modelling of hydrothermal characteristics of centrifugal nozzles

    International Nuclear Information System (INIS)

    Yarkho, A.A.; Omelchenko, M.P.; Borshchev, V.A.

    1990-01-01

    Presented for the first time is a method of recalculating the hydrothermal characteristics of centrifugal nozzles obtained in laboratory conditions for full-scale nozzles. From the experimental hydrothermal characteristics of nozzles observed in the laboratory it is allowed to calculate the hydrothermal characteristics of any other centrifugal nozzle whose diameter and dimensionless geometric characteristic are known

  15. Hybrid rocket engine, theoretical model and experiment

    Science.gov (United States)

    Chelaru, Teodor-Viorel; Mingireanu, Florin

    2011-06-01

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

  16. Axisymmetric thrust-vectoring nozzle performance prediction

    International Nuclear Information System (INIS)

    Wilson, E. A.; Adler, D.; Bar-Yoseph, P.Z

    1998-01-01

    Throat-hinged geometrically variable converging-diverging thrust-vectoring nozzles directly affect the jet flow geometry and rotation angle at the nozzle exit as a function of the nozzle geometry, the nozzle pressure ratio and flight velocity. The consideration of nozzle divergence in the effective-geometric nozzle relation is theoretically considered here for the first time. In this study, an explicit calculation procedure is presented as a function of nozzle geometry at constant nozzle pressure ratio, zero velocity and altitude, and compared with experimental results in a civil thrust-vectoring scenario. This procedure may be used in dynamic thrust-vectoring nozzle design performance predictions or analysis for civil and military nozzles as well as in the definition of initial jet flow conditions in future numerical VSTOL/TV jet performance studies

  17. Equivalent nozzle in thermomechanical problems

    International Nuclear Information System (INIS)

    Cesari, F.

    1977-01-01

    When analyzing nuclear vessels, it is most important to study the behavior of the nozzle cylinder-cylinder intersection. For the elastic field, this analysis in three dimensions is quite easy using the method of finite elements. The same analysis in the non-linear field becomes difficult for designs in 3-D. It is therefore necessary to resolve a nozzle in two dimensions equivalent to a 3-D nozzle. The purpose of the present work is to find an equivalent nozzle both with a mechanical and thermal load. This has been achieved by the analysis in three dimensions of a nozzle and a nozzle cylinder-sphere intersection, of a different radius. The equivalent nozzle will be a nozzle with a sphere radius in a given ratio to the radius of a cylinder; thus, the maximum equivalent stress is the same in both 2-D and 3-D. The nozzle examined derived from the intersection of a cylindrical vessel of radius R=191.4 mm and thickness T=6.7 mm with a cylindrical nozzle of radius r=24.675 mm and thickness t=1.350 mm, for which the experimental results for an internal pressure load are known. The structure was subdivided into 96 finite, three-dimensional and isoparametric elements with 60 degrees of freedom and 661 total nodes. Both the analysis with a mechanical load as well as the analysis with a thermal load were carried out on this structure according to the Bersafe system. The thermal load consisted of a transient typical of an accident occurring in a sodium-cooled fast reactor, with a peak of the temperature (540 0 C) for the sodium inside the vessel with an insulating argon temperature constant at 525 0 C. The maximum value of the equivalent tension was found in the internal area at the union towards the vessel side. The analysis of the nozzle in 2-D consists in schematizing the structure as a cylinder-sphere intersection, where the sphere has a given relation to the

  18. Numerical Simulation of Twin Nozzle Injectors

    OpenAIRE

    Milak, Dino

    2015-01-01

    Fuel injectors for marine applications have traditionally utilized nozzles with symmetric equispaced orifice configuration. But in light of the new marine emission legislations the twin nozzle concept has arisen. The twin nozzle differs from the conventional configuration by utilizing two closely spaced orifices to substitute each orifice in the conventional nozzle. Injector manufacturers regard twin nozzle injectors as a promising approach to facilitate stable spray patterns independent of t...

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

    Science.gov (United States)

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

    2014-06-01

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

  20. The use of solid phase microextraction as sample preparation technique for determination of n-nitrosodimethylamine in water polluted by hydrazine-based rocket fuel

    Directory of Open Access Journals (Sweden)

    Bulat Kenessov

    2009-05-01

    Full Text Available A paper describes a method for determination of N-nitrosodimethylamine in water, polluted by spills of 1,1-dimethylhydrazine, based on solid phase microextraction coupled to gas chromatography/mass spectrometry. A method detection limit was determined to be 1 ug/kg, relative error was below 20%. A method is very sensitive and selective as well as quite simple, relatively cheap and fully automated.

  1. Eddie Rocket's Franchise

    OpenAIRE

    Vahter, Jenni

    2008-01-01

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

  2. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

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

  3. The flight of uncontrolled rockets

    CERN Document Server

    Gantmakher, F R; Dryden, H L

    1964-01-01

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

  4. Porous solid backbone impregnation for electrochemical energy conversion systems

    KAUST Repository

    Boulfrad, Samir

    2013-09-19

    An apparatus and method for impregnating a porous solid backbone. The apparatus may include a platform for holding a porous solid backbone, an ink jet nozzle configured to dispense a liquid solution onto the porous solid backbone, a positioning mechanism configured to position the ink jet nozzle proximate to a plurality of locations of the porous solid backbone, and a control unit configured to control the positioning mechanism to position the ink jet nozzle proximate to the plurality of locations and cause the ink jet nozzle to dispense the liquid solution onto the porous solid backbone.

  5. Porous solid backbone impregnation for electrochemical energy conversion systems

    KAUST Repository

    Boulfrad, Samir; Jabbour, Ghassan

    2013-01-01

    An apparatus and method for impregnating a porous solid backbone. The apparatus may include a platform for holding a porous solid backbone, an ink jet nozzle configured to dispense a liquid solution onto the porous solid backbone, a positioning mechanism configured to position the ink jet nozzle proximate to a plurality of locations of the porous solid backbone, and a control unit configured to control the positioning mechanism to position the ink jet nozzle proximate to the plurality of locations and cause the ink jet nozzle to dispense the liquid solution onto the porous solid backbone.

  6. Nozzle geometry for organic vapor jet printing

    Science.gov (United States)

    Forrest, Stephen R.; McGraw, Gregory

    2017-10-25

    A first device is provided. The device includes a print head. The print head further includes a first nozzle hermetically sealed to a first source of gas. The first nozzle has an aperture having a smallest dimension of 0.5 to 500 microns in a direction perpendicular to a flow direction of the first nozzle. At a distance from the aperture into the first nozzle that is 5 times the smallest dimension of the aperture of the first nozzle, the smallest dimension perpendicular to the flow direction is at least twice the smallest dimension of the aperture of the first nozzle.

  7. Design criteria of launching rockets for burst aerial shells

    Energy Technology Data Exchange (ETDEWEB)

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

    2000-04-01

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

  8. Focusing liquid microjets with nozzles

    International Nuclear Information System (INIS)

    Acero, A J; Ferrera, C; Montanero, J M; Gañán-Calvo, A M

    2012-01-01

    The stability of flow focusing taking place in a converging–diverging nozzle, as well as the size of the resulting microjets, is examined experimentally in this paper. The results obtained in most aspects of the problem are similar to those of the classical plate-orifice configuration. There is, however, a notable difference between flow focusing in nozzles and in the plate-orifice configuration. In the former case, the liquid meniscus oscillates laterally (global whipping) for a significant area of the control parameter plane, a phenomenon never observed when focusing with the plate-orifice configuration. Global whipping may constitute an important drawback of flow focusing with nozzles because it reduces the robustness of the technique. (paper)

  9. Prototype Morphing Fan Nozzle Demonstrated

    Science.gov (United States)

    Lee, Ho-Jun; Song, Gang-Bing

    2004-01-01

    Ongoing research in NASA Glenn Research Center's Structural Mechanics and Dynamics Branch to develop smart materials technologies for aeropropulsion structural components has resulted in the design of the prototype morphing fan nozzle shown in the photograph. This prototype exploits the potential of smart materials to significantly improve the performance of existing aircraft engines by introducing new inherent capabilities for shape control, vibration damping, noise reduction, health monitoring, and flow manipulation. The novel design employs two different smart materials, a shape-memory alloy and magnetorheological fluids, to reduce the nozzle area by up to 30 percent. The prototype of the variable-area fan nozzle implements an overlapping spring leaf assembly to simplify the initial design and to provide ease of structural control. A single bundle of shape memory alloy wire actuators is used to reduce the nozzle geometry. The nozzle is subsequently held in the reduced-area configuration by using magnetorheological fluid brakes. This prototype uses the inherent advantages of shape memory alloys in providing large induced strains and of magnetorheological fluids in generating large resistive forces. In addition, the spring leaf design also functions as a return spring, once the magnetorheological fluid brakes are released, to help force the shape memory alloy wires to return to their original position. A computerized real-time control system uses the derivative-gain and proportional-gain algorithms to operate the system. This design represents a novel approach to the active control of high-bypass-ratio turbofan engines. Researchers have estimated that such engines will reduce thrust specific fuel consumption by 9 percent over that of fixed-geometry fan nozzles. This research was conducted under a cooperative agreement (NCC3-839) at the University of Akron.

  10. South Pole rockets, (1)

    International Nuclear Information System (INIS)

    Kimura, Iwane

    1977-01-01

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

  11. Multielement suppressor nozzles for thrust augmentation systems.

    Science.gov (United States)

    Lawrence, R. L.; O'Keefe, J. V.; Tate, R. B.

    1972-01-01

    The noise reduction and nozzle performance characteristics of large-scale, high-aspect-ratio multielement nozzle arrays operated at low velocities were determined by test. The nozzles are selected for application to high-aspect-ratio augmentor suppressors to be used for augmentor wing airplanes. Significant improvements in noise characteristics for multielement nozzles over those of round or high-aspect-ratio slot nozzles are obtained. Elliptical noise patterns typical of slot nozzles are presented for high-aspect-ratio multielement nozzle arrays. Additional advantages are available in OASPL noise reduction from the element size and spacing. Augmentor-suppressor systems can be designed for maximum beam pattern directivity and frequency spectrum shaping advantages. Measurements of the nozzle wakes show a correlation with noise level data and frequency spectrum peaks. The noise and jet wake results are compared with existing prediction procedures based on empirical jet flow equations, Lighthill relationships, Strouhal number, and empirical shock-induced screech noise effects.

  12. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    Science.gov (United States)

    Morris, Christopher I.

    2005-01-01

    Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous

  13. Ambipolar ion acceleration in an expanding magnetic nozzle

    Energy Technology Data Exchange (ETDEWEB)

    Longmier, Benjamin W; Carter, Mark D; Cassady, Leonard D; Chancery, William J; Diaz, Franklin R Chang; Glover, Tim W; Ilin, Andrew V; McCaskill, Greg E; Olsen, Chris S; Squire, Jared P [Ad Astra Rocket Company, 141 W. Bay Area Blvd, Webster, TX (United States); Bering, Edgar A III [Department of Physics and Department of Electrical and Computer Engineering, University of Houston, 617 Science and Research Building 1, Houston, TX (United States); Hershkowitz, Noah [Department of Engineering Physics, University of Wisconsin, 1500 Engineering Dr., Madison, WI (United States)

    2011-02-15

    The helicon plasma stage in the Variable Specific Impulse Magnetoplasma Rocket (VASIMR (registered)) VX-200i device was used to characterize an axial plasma potential profile within an expanding magnetic nozzle region of the laboratory based device. The ion acceleration mechanism is identified as an ambipolar electric field produced by an electron pressure gradient, resulting in a local axial ion speed of Mach 4 downstream of the magnetic nozzle. A 20 eV argon ion kinetic energy was measured in the helicon source, which had a peak magnetic field strength of 0.17 T. The helicon plasma source was operated with 25 mg s{sup -1} argon propellant and 30 kW of RF power. The maximum measured values of plasma density and electron temperature within the exhaust plume were 1 x 10{sup 20} m{sup -3} and 9 eV, respectively. The measured plasma density is nearly an order of magnitude larger than previously reported steady-state helicon plasma sources. The exhaust plume also exhibits a 95% to 100% ionization fraction. The size scale and spatial location of the plasma potential structure in the expanding magnetic nozzle region appear to follow the size scale and spatial location of the expanding magnetic field. The thickness of the potential structure was found to be 10{sup 4} to 10{sup 5} {lambda}{sub De} depending on the local electron temperature in the magnetic nozzle, many orders of magnitude larger than typical laboratory double layer structures. The background plasma density and neutral argon pressure were 10{sup 15} m{sup -3} and 2 x 10{sup -5} Torr, respectively, in a 150 m{sup 3} vacuum chamber during operation of the helicon plasma source. The agreement between the measured plasma potential and plasma potential that was calculated from an ambipolar ion acceleration analysis over the bulk of the axial distance where the potential drop was located is a strong confirmation of the ambipolar acceleration process.

  14. Turbomachine combustor nozzle including a monolithic nozzle component and method of forming the same

    Science.gov (United States)

    Stoia, Lucas John; Melton, Patrick Benedict; Johnson, Thomas Edward; Stevenson, Christian Xavier; Vanselow, John Drake; Westmoreland, James Harold

    2016-02-23

    A turbomachine combustor nozzle includes a monolithic nozzle component having a plate element and a plurality of nozzle elements. Each of the plurality of nozzle elements includes a first end extending from the plate element to a second end. The plate element and plurality of nozzle elements are formed as a unitary component. A plate member is joined with the nozzle component. The plate member includes an outer edge that defines first and second surfaces and a plurality of openings extending between the first and second surfaces. The plurality of openings are configured and disposed to register with and receive the second end of corresponding ones of the plurality of nozzle elements.

  15. Simulation of a Downsized FDM Nozzle

    DEFF Research Database (Denmark)

    Hofstätter, Thomas; Pimentel, Rodrigo; Pedersen, David B.

    2015-01-01

    This document discusses the simulat-ion of a downsized nozzle for fused deposition modelling (FDM), namely the E3D HotEnd Extruder with manufactured diameters of 200-400 μm in the nozzle tip. The nozzle has been simulated in terms of heat transfer and fluid flow giving an insight into the physical...

  16. Parametric Study of Sealant Nozzle

    Science.gov (United States)

    Yamamoto, Yoshimi

    It has become apparent in recent years the advancement of manufacturing processes in the aerospace industry. Sealant nozzles are a critical device in the use of fuel tank applications for optimal bonds and for ground service support and repair. Sealants has always been a challenging area for optimizing and understanding the flow patterns. A parametric study was conducted to better understand geometric effects of sealant flow and to determine whether the sealant rheology can be numerically modeled. The Star-CCM+ software was used to successfully develop the parametric model, material model, physics continua, and simulate the fluid flow for the sealant nozzle. The simulation results of Semco sealant nozzles showed the geometric effects of fluid flow patterns and the influences from conical area reduction, tip length, inlet diameter, and tip angle parameters. A smaller outlet diameter induced maximum outlet velocity at the exit, and contributed to a high pressure drop. The conical area reduction, tip angle and inlet diameter contributed most to viscosity variation phenomenon. Developing and simulating 2 different flow models (Segregated Flow and Viscous Flow) proved that both can be used to obtain comparable velocity and pressure drop results, however; differences are seen visually in the non-uniformity of the velocity and viscosity fields for the Viscous Flow Model (VFM). A comprehensive simulation setup for sealant nozzles was developed so other analysts can utilize the data.

  17. Process for manufacturing separating nozzles

    International Nuclear Information System (INIS)

    Bier, W.; Linder, G.; Mayer, E.

    1979-01-01

    The final form of the basic body and the unit consisting of the nozzle and peeling orifice provides immovable fixing of these parts. Surfaces of various components can then be milled, using milling tools, in one operation. Assembly can be made automatic. (DG) [de

  18. Nozzle for electric dispersion reactor

    Science.gov (United States)

    Sisson, W.G.; Basaran, O.A.; Harris, M.T.

    1995-11-07

    A nozzle for an electric dispersion reactor includes two concentric electrodes, the inner one of the two delivering disperse phase fluid into a continuous phase fluid. A potential difference generated by a voltage source creates a dispersing electric field at the end of the inner electrode. 4 figs.

  19. Structure of Partially Premixed Flames and Advanced Solid Propellants

    National Research Council Canada - National Science Library

    Branch, Melvyn

    1998-01-01

    The combustion of solid rocket propellants of advanced energetic materials involves a complex process of decomposition and condensed phase reactions in the solid propellant, gaseous flame reactions...

  20. Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

    Science.gov (United States)

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

    2004-01-01

    Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

  1. Particle bed reactor nuclear rocket concept

    International Nuclear Information System (INIS)

    Ludewig, H.

    1991-01-01

    The particle bed reactor nuclear rocket concept consists of fuel particles (in this case (U,Zr)C with an outer coat of zirconium carbide). These particles are packed in an annular bed surrounded by two frits (porous tubes) forming a fuel element; the outer one being a cold frit, the inner one being a hot frit. The fuel element are cooled by hydrogen passing in through the moderator. These elements are assembled in a reactor assembly in a hexagonal pattern. The reactor can be either reflected or not, depending on the design, and either 19 or 37 elements, are used. Propellant enters in the top, passes through the moderator fuel element and out through the nozzle. Beryllium used for the moderator in this particular design to withstand the high radiation exposure implied by the long run times

  2. Another Look at Rocket Thrust

    Science.gov (United States)

    Hester, Brooke; Burris, Jennifer

    2012-01-01

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

  3. Nozzle geometry variations on the discharge coefficient

    Directory of Open Access Journals (Sweden)

    M.M.A. Alam

    2016-03-01

    Full Text Available Numerical works have been conducted to investigate the effect of nozzle geometries on the discharge coefficient. Several contoured converging nozzles with finite radius of curvatures, conically converging nozzles and conical divergent orifices have been employed in this investigation. Each nozzle and orifice has a nominal exit diameter of 12.7×10−3 m. A 3rd order MUSCL finite volume method of ANSYS Fluent 13.0 was used to solve the Reynolds-averaged Navier–Stokes equations in simulating turbulent flows through various nozzle inlet geometries. The numerical model was validated through comparison between the numerical results and experimental data. The results obtained show that the nozzle geometry has pronounced effect on the sonic lines and discharge coefficients. The coefficient of discharge was found differ from unity due to the non-uniformity of flow parameters at the nozzle exit and the presence of boundary layer as well.

  4. The History of Rockets.

    Science.gov (United States)

    Newby, J. C.

    1988-01-01

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

  5. Development of an Aeroelastic Modeling Capability for Transient Nozzle Side Load Analysis

    Science.gov (United States)

    Wang, Ten-See; Zhao, Xiang; Zhang, Sijun; Chen, Yen-Sen

    2013-01-01

    Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development during test. While three-dimensional, transient, turbulent, chemically reacting computational fluid dynamics methodology has been demonstrated to capture major side load physics with rigid nozzles, hot-fire tests often show nozzle structure deformation during major side load events, leading to structural damages if structural strengthening measures were not taken. The modeling picture is incomplete without the capability to address the two-way responses between the structure and fluid. The objective of this study is to develop a coupled aeroelastic modeling capability by implementing the necessary structural dynamics component into an anchored computational fluid dynamics methodology. The computational fluid dynamics component is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, while the computational structural dynamics component is developed in the framework of modal analysis. Transient aeroelastic nozzle startup analyses of the Block I Space Shuttle Main Engine at sea level were performed. The computed results from the aeroelastic nozzle modeling are presented.

  6. Scaled Rocket Testing in Hypersonic Flow

    Science.gov (United States)

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

    2015-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Kalomba Mboyi

    2015-04-01

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

  8. Nuclear thermal rocket plume interactions with spacecraft. Final report

    International Nuclear Information System (INIS)

    Mauk, B.H.; Gatsonis, N.A.; Buzby, J.; Yin, X.

    1997-01-01

    This is the first study that has treated the Nuclear Thermal Rocket (NTR) effluent problem in its entirety, beginning with the reactor core, through the nozzle flow, to the plume backflow. The summary of major accomplishments is given below: (1) Determined the NTR effluents that include neutral, ionized and radioactive species, under typical NTR chamber conditions. Applied an NTR chamber chemistry model that includes conditions and used nozzle geometries and chamber conditions typical of NTR configurations. (2) Performed NTR nozzle flow simulations using a Navier-Stokes solver. We assumed frozen chemistry at the chamber conditions and used nozzle geometries and chamber conditions typical of NTR configurations. (3) Performed plume simulations using a Direct Simulation Monte Carlo (DSMC) code with chemistry. In order to account for radioactive trace species that may be important for contamination purposes we developed a multi-weighted DSMC methodology. The domain in our simulations included large regions downstream and upstream of the exit. Inputs were taken from the Navier-Stokes solutions

  9. Two-way coupled simulation of a flow laden with metallic particulates in overexpanded TIC nozzle

    International Nuclear Information System (INIS)

    Moshfegh, Abouzar; Shams, Mehrzad; Ebrahimi, Reza; Farnia, Mohammad Ali

    2009-01-01

    A simulation of non-reacting dilute gas-solid flow in a truncated ideal contour nozzle with consideration of external stream interactions is performed. The Eulerian-Lagrangian approach involving a two-way momentum and thermal coupling between gas and particles phases is also adopted. Of interests are to investigate the effects of particles diameter and mass flow fraction on the flow pattern, Mach number, pressure and temperature contours and their distributions along the nozzle centerline and wall. The main goal is to determine the separation point quantitatively when the particles characteristics change. Particles sample trajectories are illustrated throughout the flow field and a qualitative discussion on the way that physical properties of the nozzle exit flow and particles trajectories oscillate is prepared. The existence of solid particulates delays the separation prominently in the cases studied. The bigger particles and the higher particles mass flow fractions respectively advance and delay the separation occurrence. The particles trajectories oscillate when they expose to the crisscrossing (or diamond-shape) shock waves generated outside the nozzle to approach the exit jet conditions to the ambient. The simulation code is validated and verified, respectively, against a one-phase 2D convergent-divergent nozzle flow and a two-phase Jet Propulsion Laboratory nozzle flow, and acceptable agreements are achieved.

  10. Effect of Stagger on the Vibroacoustic Loads from Clustered Rockets

    Science.gov (United States)

    Rojo, Raymundo; Tinney, Charles E.; Ruf, Joseph H.

    2016-01-01

    The effect of stagger startup on the vibro-acoustic loads that form during the end- effects-regime of clustered rockets is studied using both full-scale (hot-gas) and laboratory scale (cold gas) data. Both configurations comprise three nozzles with thrust optimized parabolic contours that undergo free shock separated flow and restricted shock separated flow as well as an end-effects regime prior to flowing full. Acoustic pressure waveforms recorded at the base of the nozzle clusters are analyzed using various statistical metrics as well as time-frequency analysis. The findings reveal a significant reduction in end- effects-regime loads when engine ignition is staggered. However, regardless of stagger, both the skewness and kurtosis of the acoustic pressure time derivative elevate to the same levels during the end-effects-regime event thereby demonstrating the intermittence and impulsiveness of the acoustic waveforms that form during engine startup.

  11. US Rocket Propulsion Industrial Base Health Metrics

    Science.gov (United States)

    Doreswamy, Rajiv

    2013-01-01

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

  12. Optimization of Construction of the rocket-assisted projectile

    Directory of Open Access Journals (Sweden)

    Arkhipov Vladimir

    2017-01-01

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

  13. Performance modelling of plasma microthruster nozzles in vacuum

    Science.gov (United States)

    Ho, Teck Seng; Charles, Christine; Boswell, Rod

    2018-05-01

    Computational fluid dynamics and plasma simulations of three geometrical variations of the Pocket Rocket radiofrequency plasma electrothermal microthruster are conducted, comparing pulsed plasma to steady state cold gas operation. While numerical limitations prevent plasma modelling in a vacuum environment, results may be obtained by extrapolating from plasma simulations performed in a pressurised environment, using the performance delta from cold gas simulations performed in both environments. Slip regime boundary layer effects are significant at these operating conditions. The present investigation targets a power budget of ˜10 W for applications on CubeSats. During plasma operation, the thrust force increases by ˜30% with a power efficiency of ˜30 μNW-1. These performance metrics represent instantaneous or pulsed operation and will increase over time as the discharge chamber attains thermal equilibrium with the heated propellant. Additionally, the sculpted nozzle geometry achieves plasma confinement facilitated by the formation of a plasma sheath at the nozzle throat, and fast recombination ensures a neutral exhaust plume that avoids the contamination of solar panels and interference with externally mounted instruments.

  14. Variable volume combustor with pre-nozzle fuel injection system

    Science.gov (United States)

    Keener, Christopher Paul; Johnson, Thomas Edward; McConnaughhay, Johnie Franklin; Ostebee, Heath Michael

    2016-09-06

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system.

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

    Science.gov (United States)

    Gordon, Sanford; Kastner, Michael E

    1958-01-01

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

  16. Fractal analysis of agricultural nozzles spray

    Directory of Open Access Journals (Sweden)

    Francisco Agüera

    2012-02-01

    Full Text Available Fractal scaling of the exponential type is used to establish the cumulative volume (V distribution applied through agricultural spray nozzles in size x droplets, smaller than the characteristic size X. From exponent d, we deduced the fractal dimension (Df which measures the degree of irregularity of the medium. This property is known as 'self-similarity'. Assuming that the droplet set from a spray nozzle is self-similar, the objectives of this study were to develop a methodology for calculating a Df factor associated with a given nozzle and to determine regression coefficients in order to predict droplet spectra factors from a nozzle, taking into account its own Df and pressure operating. Based on the iterated function system, we developed an algorithm to relate nozzle types to a particular value of Df. Four nozzles and five operating pressure droplet size characteristics were measured using a Phase Doppler Particle Analyser (PDPA. The data input consisted of droplet size spectra factors derived from these measurements. Estimated Df values showed dependence on nozzle type and independence of operating pressure. We developed an exponential model based on the Df to enable us to predict droplet size spectra factors. Significant coefficients of determination were found for the fitted model. This model could prove useful as a means of comparing the behavior of nozzles which only differ in not measurable geometric parameters and it can predict droplet spectra factors of a nozzle operating under different pressures from data measured only in extreme work pressures.

  17. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

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

  18. Cryogenic rocket engine development at Delft aerospace rocket engineering

    NARCIS (Netherlands)

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

    2016-01-01

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

  19. Effects of axial gap and nozzle distribution on aerodynamic forces of a supersonic partial-admission turbine

    Directory of Open Access Journals (Sweden)

    Jinpeng JIANG

    2017-12-01

    Full Text Available The turbine in an LH2/LOX rocket engine is designed as a two-stage supersonic partial-admission turbine. Three-dimensional steady and unsteady simulations were conducted to analyze turbine performance and aerodynamic forces on rotor blades. Different configurations were employed to investigate the effects of the axial gap and nozzle distribution on the predicted performance and aerodynamic forces. Rotor blades experience unsteady aerodynamic forces because of the partial admission. Aerodynamic forces show periodicity in the admission region, and are close to zero after leaving the admission region. The unsteady forces in frequency domain indicate that components exist in a wide frequency region, and the admission passing frequency is dominant. Those multiples of the rotational frequency which are multiples of the nozzle number in a full-admission turbine are notable components. Results show that the turbine efficiency decreases as the axial gap between nozzles and the 1st stage rotor (rotor 1 increases. Fluctuation of the circumferential aerodynamic force on rotor 1 blades decreases with the axial gap increasing. The turbine efficiency decreases as the circumferential spacing between nozzles increases. Fluctuations of the circumferential and axial aerodynamic forces increase as the circumferential spacing increases. As for the non-equidistant nozzle distribution, it produces similar turbine performance and amplitude-frequency characteristics of forces to those of the normal configuration, when the mean spacing is equal to that of the normal case. Keywords: Aerodynamic force, Axial gap, Computational fluid dynamics (CFD, Nozzle distribution, Partial admission, Turbine

  20. Through an Annular Turbine Nozzle

    Directory of Open Access Journals (Sweden)

    Rainer Kurz

    1995-01-01

    is located in the gas turbine. The experiments were performed using total pressure probes and wall static pressure taps. The pitch variation modifies the flow field both upstream and downstream of the nozzle, although the experiments show that the effect is localized to the immediate neighborhood of the involved blades. The effects on the wakes and on the inviscid flow are discussed separately. The mean velocities show a strong sensitivity to the changes of the pitch, which is due to a potential flow effect rather than a viscous effect.

  1. Axisymmetric nozzles with chamfered contraction

    Czech Academy of Sciences Publication Activity Database

    Tesař, Václav

    2017-01-01

    Roč. 263, August (2017), s. 147-158 ISSN 0924-4247 Institutional support: RVO:61388998 Keywords : nozzles * chamfering * invariant Subject RIV: BK - Fluid Dynamics OBOR OECD: Fluids and plasma physics (including surface physics) Impact factor: 2.499, year: 2016 http://ac.els-cdn.com/S0924424716310329/1-s2.0-S0924424716310329-main.pdf?_tid=f953dc4c-873c-11e7-b8d0-00000aacb35d&acdnat=1503408341_51527a384c272a3c4e8f43e6046d789d

  2. Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

    Science.gov (United States)

    Sathiyanathan, Kartheephan

    This thesis describes the design, fabrication, and testing of a solid propellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

  3. Stress intensity factors of corner cracks in two nozzle-cylinder intersections

    International Nuclear Information System (INIS)

    Kobayashi, A.S.; Polvanich, N.; Emery, A.F.; Love, W.J.

    1977-01-01

    In a recent paper, the authors presented the stress-intensity-magnification factors of a quarter-elliptical surface crack in a quarter-infinite solid and a circular crack approaching a reentry corner in a three-quarter infinite solid. These stress-intensity-magnification factors were used together with a curvature-correction factor to estimate the stress-intensity factor of a corner crack at a nozzle-cylinder intersection. Through appropriate superposition of the above stress-intensity-magnification factors, stress-intensity factors for hypothetical corner cracks at a nozzle-cylinder intersection subjected to internal pressure and transient thermal-stress loadings can be obtained. A description of a computer code based on this procedure as well as its applications in analyzing two corner-crack problems at a nozzle-cylinder intersection are discussed in this paper

  4. Stress intensity factors of corner cracks in two nozzle-cylinder interactions

    International Nuclear Information System (INIS)

    Kobayashi, A.S.; Polvanich, N.; Emery, A.F.; Love, W.J.

    1977-01-01

    In a recent paper, the authors presented the stress-intensity-magnification factors of a quarter-elliptical surface crack in a quarter-infinite solid and a circular crack approaching a reentry corner in a three-quarter infinite solid. These stress-intensity-magnification factors were used together with a curvature-correction factor to estimate the stress-intensity factor of a corner crack at a nozzle-cylinder interaction. Through appropriate superposition of the above stress-intensity-magnification factors, stress-intensity factors for hypothetical corner cracks at a nozzle-cylinder intersection subjected to internal pressure and transient thermal-stress loadings can be obtained. A description of a computer code based on this procedure as well as its applications in analyzing two corner-crack probems at a nozzle-cylinder intersection are discussed in this paper. (Auth.)

  5. Large Liquid Rocket Testing: Strategies and Challenges

    Science.gov (United States)

    Rahman, Shamim A.; Hebert, Bartt J.

    2005-01-01

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

  6. This Is Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

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

  7. The Relativistic Rocket

    Science.gov (United States)

    Antippa, Adel F.

    2009-01-01

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

  8. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

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

  9. ROCKETS: Soar to Success

    Science.gov (United States)

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

    2017-01-01

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

  10. Liquid Rocket Engine Testing

    Science.gov (United States)

    Rahman, Shamim

    2005-01-01

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

  11. Baking Soda and Vinegar Rockets

    Science.gov (United States)

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

    2009-01-01

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

  12. Pengaruh Jarak dan Posisi Nozzle terhadap Daya Turbin Pelton

    OpenAIRE

    Kurniawan, Yani; Pane, Erlanda Augupta; Ismail, Ismail

    2017-01-01

    Pelton Turbine is a turbine which use nozzle as officers the direction of a stream water in order to move around of blade turbine. The rotating of turbine blade efected by some parameters such as the distance of the nozzle, position of nozzle, diameter of nozzle, number of nozzle, and the geometry shape of the blade turbine. An experimental study to analyze the affect of distance and position nozzle to Pelton Turbine of performance. The research method used experiment parameter was position o...

  13. Pengaruh Jarak dan Posisi Nozzle Terhadap Daya Turbin Pelton

    OpenAIRE

    Yani Kurniawan; Erlanda Augupta Pane; Ismail

    2017-01-01

    Pelton Turbine is a turbine which use nozzle as officers the direction of a stream water in order to move around of blade turbine. The rotating of turbine blade efected by some parameters such as the distance of the nozzle, position of nozzle, diameter of nozzle, number of nozzle, and the geometry shape of the blade turbine. An experimental study to analyze the affect of distance and position nozzle to Pelton Turbine of performance. The research method used experiment parameter was position o...

  14. Multivariable optimization of liquid rocket engines using particle swarm algorithms

    Science.gov (United States)

    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.

  15. Retrofit design of remotely removable decontamination spray nozzles for the new waste calcining facility at the Idaho National Engineering Laboratory

    International Nuclear Information System (INIS)

    Gay, J.A.

    1988-01-01

    High level radioactive liquid waste is converted to a solid form at the Idaho Chemical Processing Plant (ICPP). The conversion is done by a fluidized bed combustion process in the calciner vessel. The interior decontamination system for the calciner vessel uses a common header bolted to four decontamination nozzles around the upper head. The retrofit was required to eliminate hands-on maintenance and difficulty in nozzle removal because of nozzle plugging. The retrofit design for this project demonstrates the solution of problems associated with thermal phenomena, structural supports, seismic requirements, remote handling and installations into extremely restricted spaces

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

    Science.gov (United States)

    Ferragut, N. J.

    1976-01-01

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

  17. Palo Verde Unit 3 BMI nozzle modification

    International Nuclear Information System (INIS)

    Waskey, D.

    2015-01-01

    The 61 BMI (Bottom Mount Instrumentation) nozzles of the unit 3 of the Palo Verde plant have been examined through ASME Code Case N722. The nozzle 3 was the only one with leakage noted. The ultrasound testing results are characteristic of PWSCC (Primary Water Stress Corrosion Cracking). The initiation likely occurred at a weld defect which was exposed to the primary water environment resulting in PWSCC. All other nozzles (60) showed no unacceptable indications. Concerning nozzle 3 one crack in J-groove weld connected large defect to primary water. An environmental model has been used to simulate and optimize the repair. The AREVA crew was on site 18 days after contract award and the job was completed in 12 days, 30 hours ahead of baseline schedule. This series of slides describes the examination of the BMI nozzles, the repair steps, and alternative design concepts

  18. Isentropic Gas Flow for the Compressible Euler Equation in a Nozzle

    Science.gov (United States)

    Tsuge, Naoki

    2013-08-01

    We study the motion of isentropic gas in a nozzle. Nozzles are used to increase the thrust of engines or to accelerate a flow from subsonic to supersonic. Nozzles are essential parts for jet engines, rocket engines and supersonicwind tunnels. In the present paper, we consider unsteady flow, which is governed by the compressible Euler equation, and prove the existence of global solutions for the Cauchy problem. For this problem, the existence theorem has already been obtained for initial data away from the sonic state, (Liu in Commun Math Phys 68:141-172, 1979). Here, we are interested in the transonic flow, which is essential for engineering and physics. Although the transonic flow has recently been studied (Tsuge in J Math Kyoto Univ 46:457-524, 2006; Lu in Nonlinear Anal Real World Appl 12:2802-2810, 2011), these papers assume monotonicity of the cross section area. Here, we consider the transonic flow in a nozzle with a general cross section area. When we prove global existence, the most difficult point is obtaining a bounded estimate for approximate solutions. To overcome this, we employ a new invariant region that depends on the space variable. Moreover, we introduce a modified Godunov scheme. The corresponding approximate solutions consist of piecewise steady-state solutions of an auxiliary equation, which yield a desired bounded estimate. In order to prove their convergence, we use the compensated compactness framework.

  19. External Cylindrical Nozzle with Controlled Vacuum

    Directory of Open Access Journals (Sweden)

    V. N. Pil'gunov

    2015-01-01

    Full Text Available There is a developed design of the external cylindrical nozzle with a vacuum camera. The paper studies the nozzle controllability of flow rate via regulated connection of the evacuated chamber to the atmosphere through an air throttle. Working capacity of the nozzle with inlet round or triangular orifice are researched. The gap is provided in the nozzle design between the external wall of the inlet orifice and the end face of the straight case in the nozzle case. The presented mathematical model of the nozzle with the evacuated chamber allows us to estimate the expected vacuum amount in the compressed section of a stream and maximum permissible absolute pressure at the inlet orifice. The paper gives experimental characteristics of the fluid flow process through the nozzle for different values of internal diameter of a straight case and an extent of its end face remoteness from an external wall of the inlet orifice. It estimates how geometry of nozzle constructive elements influences on the volume flow rate. It is established that the nozzle capacity significantly depends on the shape of inlet orifice. Triangular orifice nozzles steadily work in the mode of completely filled flow area of the straight case at much more amounts of the limit pressure of the flow. Vacuum depth in the evacuated chamber also depends on the shape of inlet orifice: the greatest vacuum is reached in a nozzle with the triangular orifice which 1.5 times exceeds the greatest vacuum with the round orifice. Possibility to control nozzle capacity through the regulated connection of the evacuated chamber to the atmosphere was experimentally estimated, thus depth of flow rate regulation of the nozzle with a triangular orifice was 45% in comparison with 10% regulation depth of the nozzle with a round orifice. Depth of regulation calculated by a mathematical model appeared to be much more. The paper presents experimental dependences of the flow coefficients of nozzle input orifice

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

    Science.gov (United States)

    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.

  1. 2005 40th Annual Armament Systems: Guns - Ammunition - Rockets - Missiles Conference and Exhibition. Volume 2: Wednesday

    Science.gov (United States)

    2005-04-28

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

  2. The relativistic rocket

    Energy Technology Data Exchange (ETDEWEB)

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

    2009-05-15

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful method that can be applied to a wide range of special relativistic problems of linear acceleration.

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

    Directory of Open Access Journals (Sweden)

    Jichao Hu

    2014-01-01

    Full Text Available 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.

  4. Supersonic flaw detection device for nozzle

    International Nuclear Information System (INIS)

    Hata, Moriki.

    1996-01-01

    In a supersonic flaw detection device to be attached to a body surface of a reactor pressure vessel for automatically detecting flaws of a welded portion of a horizontally connected nozzle by using supersonic waves, a running vehicle automatically running along a circumferential direction of the nozzle comprises a supersonic flaw detection means for detecting flaws of the welded portion of the nozzle by using supersonic waves, and an inclination angle sensor for detecting the inclination angle of the running vehicle relative to the central axis of the nozzle. The running distance of the vehicle running along the circumference of the nozzle, namely, the position of the running vehicle from a reference point of the nozzle can be detected accurately by dividing the distance around the nozzle by the inclination angle detected by the inclination angle sensor. Accordingly, disadvantages in the prior art, for example, that the detected values obtained by using an encoder are changed by slipping or idle running of the magnet wheels are eliminated, and accurate flaw detection can be conducted. In addition, an operation of visually adjusting the reference point for the device can be eliminated. An operator's exposure dose can be reduced. (N.H.)

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

    Science.gov (United States)

    Williams, R. W. (Compiler)

    1996-01-01

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

  6. Design considerations for a pressure-driven multi-stage rocket

    Science.gov (United States)

    Sauerwein, Steven Craig

    2002-01-01

    The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

  7. Liquid Rocket Engine Testing Overview

    Science.gov (United States)

    Rahman, Shamim

    2005-01-01

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

  8. Aerospike Nozzle for Rotating Detonation Engine Application

    Data.gov (United States)

    National Aeronautics and Space Administration — This proposal presents a graduate MS research thesis on improving the efficiency of rotating detonation engines by using aerospike nozzle technologies. A rotating...

  9. Space Storable Hybrid Rockets for Orbit Insertion or In Situ Resource Utilization Applications

    Data.gov (United States)

    National Aeronautics and Space Administration — This research effort will pave the way towards a Mars Sample Return (MSR) campaign and potentially, future human exploration of Mars. Hybrid rockets utilize a solid...

  10. Rocket + Science = Dialogue

    Science.gov (United States)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

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

  11. Rocket Assembly and Checkout Facility

    Data.gov (United States)

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

  12. Transient Three-Dimensional Side Load Analysis of Out-of-Round Film Cooled Nozzles

    Science.gov (United States)

    Wang, Ten-See; Lin, Jeff; Ruf, Joe; Guidos, Mike

    2010-01-01

    The objective of this study is to investigate the effect of nozzle out-of-roundness on the transient startup side loads at a high altitude, with an anchored computational methodology. The out-of-roundness could be the result of asymmetric loads induced by hardware attached to the nozzle, asymmetric internal stresses induced by previous tests, and deformation, such as creep, from previous tests. The rocket engine studied encompasses a regeneratively cooled thrust chamber and a film cooled nozzle extension with film coolant distributed from a turbine exhaust manifold. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Transient startup computations were performed with the out-of-roundness achieved by four different degrees of ovalization: one perfectly round, one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation-line-jump is the peak side load physics for the round, slightly our-of-round, and more out-of-round cases, and the peak side load increases as the degree of out-of-roundness increases. For the significantly out-of-round nozzle, however, the peak side load reduces to comparable to that of the round nozzle and the separation line jump is not the peak side load physics. The counter-intuitive result of the significantly out-of-round case is found to be related to a side force reduction mechanism that splits the effect of the separation-line-jump into two parts, not only in the circumferential direction and most importantly in time.

  13. Transient Three-Dimensional Analysis of Nozzle Side Load in Regeneratively Cooled Engines

    Science.gov (United States)

    Wang, Ten-See

    2005-01-01

    Three-dimensional numerical investigations on the start-up side load physics for a regeneratively cooled, high-aspect-ratio nozzle were performed. The objectives of this study are to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system simulation. Computations were performed for both the adiabatic and cooled walls in order to understand the effect of boundary conditions. Finite-rate chemistry was used throughout the study so that combustion effect is always included. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with oscillation of shocks across the lip. Wall boundary conditions drastically affect the computed side load physics: the adiabatic nozzle prefers free-shock separation while the cooled nozzle favors restricted-shock separation, resulting in higher peak side load for the cooled nozzle than that of the adiabatic nozzle. By comparing the computed physics with those of test observations, it is concluded that cooled wall is a more realistic boundary condition, and the oscillation of the restricted-shock separation flow pattern across the lip along with its associated tangential shock motion are the dominant side load physics for a regeneratively cooled, high aspect-ratio rocket engine.

  14. Rocket-Powered Parachutes Rescue Entire Planes

    Science.gov (United States)

    2010-01-01

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

  15. Nuclear rocket propulsion

    International Nuclear Information System (INIS)

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

    1991-01-01

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

  16. Cross-talk effect in electrostatic based capillary array nozzles

    International Nuclear Information System (INIS)

    Choi, Kyung Hyun; Rahman, Khalid; Khan, Arshad; Kim, Dong Soo

    2011-01-01

    Electrohydrodynamic printing is a promising technique for printed electronics application. Most researchers working in this field are using a single nozzle configuration. However, for large area printing a multi-nozzle setup will be required for time and cost effective process. In this paper the influence of electric field and flow-rate on jetting angle on multi-nozzle array has been investigated experimentally. A three nozzle setup has been used in a linear array by using glass capillary as a nozzle with independent voltage applied on each nozzle and independent ink supply. The experiments are performed by changing the nozzle to nozzle gap and the effect on the jetting angle has been investigated. It has been observed that by increasing the applied voltage the jetting angle also increases at fixed flow-rate. In case of increasing the flow-rate, the jetting angle first increases with increase in flow-rate, but as the flow-rate increases at certain level the jetting angle decreases; moreover, at a high flow-rate the cone-jet length starts increasing. Numerical simulation has been performed to have a better understanding of the electric-field with respect to jetting angles. The influence of one nozzle on another nozzle is also investigated by operating the nozzle independently by using different operating cases. The cross-talk effect is also minimized by reducing the nozzle diameter. At 250 μm nozzle diameter the cross-talk effect was negligible for 5 mm nozzle-to-nozzle gap. This study will help in better understanding of the interaction between different nozzles in multi-nozzle cases and better design of the multi-nozzle system by minimizing the effects of adjacent nozzles for multi-nozzle electrohydrodynamic printing system

  17. Two-Dimensional Motions of Rockets

    Science.gov (United States)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

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

    Science.gov (United States)

    Ingemar Skoog, A.

    2010-02-01

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

  19. Rhenium Rocket Manufacturing Technology

    Science.gov (United States)

    1997-01-01

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

  20. Cycle Trades for Nuclear Thermal Rocket Propulsion Systems

    Science.gov (United States)

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

    2003-01-01

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

  1. Micro-Rockets for the Classroom.

    Science.gov (United States)

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

    1999-01-01

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

  2. Single nozzle spray drift measurements of drift reducing nozzles at two forward speeds

    NARCIS (Netherlands)

    Stallinga, H.; Zande, van de J.C.; Michielsen, J.G.P.; Velde, van P.

    2016-01-01

    In 2011‒2012 single nozzle field experiments were carried out to determine the effect of different flat fan spray nozzles of the spray drift reduction classes 50, 75, 90 and 95% on spray drift at two different forward speeds (7.2 km h-1 and 14.4 km h-1). Experiments were performed with a single

  3. Performance Enhancement of the Space Shuttle RSRM Nozzle-to-Case Joint Using a Carbon Rope Barrier

    Science.gov (United States)

    Ewing, M. E.; McGuire, J. R.; McWhorter, B. B.; Frost, D. L.

    1999-01-01

    A carbon rope "thermal barrier" is being considered as a component to enhance performance of the Reusable Solid Rocket Motor (RSRM) nozzle-to-case joint. Fundamental performance characteristics of the rope have been considered in this paper. In particular, resistance to erosion, ability to filter particulate matter, thermal capacitance, and flow resistance have been considered. Testing results have shown the rope to be resistant to the corrosive internal environment of the RSRM. The rope has also been shown to be an effective "slag barrier." A desirable feature of the rope would be the ability to act as a heat sink. However, analyses have indicated that the thermal capacitance of the rope is not large enough to reduce the temperature of an impinging gas stream below the ablation temperature of the 0-ring for significant time periods, The real value of the rope is its ability to act as a flow diffuser. Flow resistance test, were performed on the rope In the course of testing the rope between parallel plates, an undesirable "blow-by" phenomenon was observed when the compressive stress in the rope was smaller than the upstream gas pressure. It was found, however, that in the converging passage of the actual design, the rope would consistently "Self-seat" and thereby prevent blow-by, even in the absence of any precompression. Flow resistance values have been quantified for use in future analyses. The work presented here provides an initial thermal-fluid assessment of the rope for this application, and lays the groundwork for future development.

  4. Nuclear Rocket Engine Reactor

    CERN Document Server

    Lanin, Anatoly

    2013-01-01

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

  5. Robotic cleaning of radwaste tank nozzles

    International Nuclear Information System (INIS)

    Boughman, G.; Jones, S.L.

    1992-01-01

    The Susquehanna radwaste processing system includes two reactor water cleanup phase separator tanks and one waste sludge phase separator tank. A system of educator nozzles and associated piping is used to provide mixing in the tanks. The mixture pumped through the nozzles is a dense resin-and-water slurry, and the nozzles tend to plug up during processing. The previous method for clearing the nozzles had been for a worker to enter the tanks and manually insert a hydrolaser into each nozzle, one at a time. The significant radiation exposure and concern for worker safety in the tank led the utility to investigate alternate means for completing this task. The typical tank configuration is shown in a figure. The initial approach investigated was to insert a manipulator arm in the tank. This arm would be installed by workers and then teleoperated from a remote control station. This approach was abandoned because of several considerations including educator location and orientation, excessive installation time, and cost. The next approach was to use a mobile platform that would operate on the tank floor. This approach was selected as being the most feasible solution. After a competitive selection process, REMOTEC was selected to provide the mobile platform. Their proposal was based on the commercial ANDROS Mark 5 platform

  6. Lower nozzle of PWR fuel assembly

    International Nuclear Information System (INIS)

    Furutani, Nobuo.

    1994-01-01

    A lower nozzle comprises a regular square plate and legs. The plate has a plurality of holes for securing thimble tubes and a great number of water flowing ports. Ridges each having a lower end surface inclined toward inner side of the plate are disposed at the outer circumference of the plate. The legs suspend downwardly from the corners of the plate and support the plate at a predetermined gap between a lower reactor core plate and the plate. The inclined surfaces of the ridges disposed at the outer circumference of the plate retain coolants, that were caused to flow to the outside passing between the legs of the nozzle, while dividing them to the inside of the nozzle and circulate the coolants upwardly passing through the water flowing ports of the plate. Further, since obstacles abut against the inclined surfaces of the ridges and flow to the inner side of the lower nozzle, obstacles in the coolants can be captured substantially entirely by the lower nozzle. (I.N.)

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

    Science.gov (United States)

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

    2010-01-01

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

  8. Two-Rockets Thought Experiment

    Science.gov (United States)

    Smarandache, Florentin

    2014-03-01

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

  9. Li/Li2 supersonic nozzle beam

    International Nuclear Information System (INIS)

    Wu, C.Y.R.; Crooks, J.B.; Yang, S.C.; Way, K.R.; Stwalley, W.C.

    1977-01-01

    The characterization of a lithium supersonic nozzle beam was made using spectroscopic techniques. It is found that at a stagnation pressure of 5.3 kPa (40 torr) and a nozzle throat diameter of 0.4 mm the ground state vibrational population of Li 2 can be described by a Boltzmann distribution with T/sub v/ = 195 +- 30 0 K. The rotational temperature is found to be T/sub r/ = 70 +- 20 0 K by band shape analysis. Measurements by quadrupole mass spectrometer indicates that approximately 10 mole per cent Li 2 dimers are formed at an oven body temperature of 1370 0 K n the supersonic nozzle expansion. This measured mole fraction is in good agreement with the existing dimerization theory

  10. Dual-nozzle microfluidic droplet generator

    Science.gov (United States)

    Choi, Ji Wook; Lee, Jong Min; Kim, Tae Hyun; Ha, Jang Ho; Ahrberg, Christian D.; Chung, Bong Geun

    2018-05-01

    The droplet-generating microfluidics has become an important technique for a variety of applications ranging from single cell analysis to nanoparticle synthesis. Although there are a large number of methods for generating and experimenting with droplets on microfluidic devices, the dispensing of droplets from these microfluidic devices is a challenge due to aggregation and merging of droplets at the interface of microfluidic devices. Here, we present a microfluidic dual-nozzle device for the generation and dispensing of uniform-sized droplets. The first nozzle of the microfluidic device is used for the generation of the droplets, while the second nozzle can accelerate the droplets and increase the spacing between them, allowing for facile dispensing of droplets. Computational fluid dynamic simulations were conducted to optimize the design parameters of the microfluidic device.

  11. Biannular Airbreathing Nozzle Rig (BANR) facility checkout and plug nozzle performance test data

    Science.gov (United States)

    Cummings, Chase B.

    2010-09-01

    The motivation for development of a supersonic business jet (SSBJ) platform lies in its ability to create a paradigm shift in the speed and reach of commercial, private, and government travel. A full understanding of the performance capabilities of exhaust nozzle configurations intended for use in potential SSBJ propulsion systems is critical to the design of an aircraft of this type. Purdue University's newly operational Biannular Airbreathing Nozzle Rig (BANR) is a highly capable facility devoted to the testing of subscale nozzles of this type. The high accuracy, six-axis force measurement system and complementary mass flowrate measurement capabilities of the BANR facility make it rather ideally suited for exhaust nozzle performance appraisal. Detailed accounts pertaining to methods utilized in the proper checkout of these diagnostic capabilities are contained herein. Efforts to quantify uncertainties associated with critical BANR test measurements are recounted, as well. Results of a second hot-fire test campaign of a subscale Gulfstream Aerospace Corporation (GAC) axisymmetric, shrouded plug nozzle are presented. Determined test article performance parameters (nozzle thrust efficiencies and discharge coefficients) are compared to those of a previous test campaign and numerical simulations of the experimental set-up. Recently acquired data is compared to published findings pertaining to plug nozzle experiments of similar scale and operating range. Suggestions relating to the future advancement and improvement of the BANR facility are provided. Lessons learned with regards to test operations and calibration procedures are divulged in an attempt to aid future facility users, as well.

  12. The Swedish sounding rocket programme

    International Nuclear Information System (INIS)

    Bostroem, R.

    1980-01-01

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

  13. Combustor nozzles in gas turbine engines

    Science.gov (United States)

    Johnson, Thomas Edward; Keener, Christopher Paul; Stewart, Jason Thurman; Ostebee, Heath Michael

    2017-09-12

    A micro-mixer nozzle for use in a combustor of a combustion turbine engine, the micro-mixer nozzle including: a fuel plenum defined by a shroud wall connecting a periphery of a forward tube sheet to a periphery of an aft tubesheet; a plurality of mixing tubes extending across the fuel plenum for mixing a supply of compressed air and fuel, each of the mixing tubes forming a passageway between an inlet formed through the forward tubesheet and an outlet formed through the aft tubesheet; and a wall mixing tube formed in the shroud wall.

  14. Antithermal shield for rockets with heat evacuation by infrared radiation reflection

    Directory of Open Access Journals (Sweden)

    Ioan RUSU

    2010-12-01

    Full Text Available At high speed, the friction between the air mass and the rocket surface causes a localheating of over 1000 Celsius degrees. For the heat protection of the rocket, on its outside surfacethermal shields are installed.Studying the Coanda effect, the fluid flow on solids surface, respectively, the author Ioan Rusuhas discovered by simply researches that the Coanda effect could be /extended also to the fluid flowon discontinuous solids, namely, on solids provided with orifices. This phenomenon was named by theauthor, the expanded Coanda effect. Starting with this discovery, the author has invented a thermalshield, registered at The State Office for inventions and Trademarks OSIM, deposit F 2010 0153This thermal shield:- is built as a covering rocket sheet with many orifices installed with a minimum space fromthe rocket body- takes over the heat fluid generated by the frontal part of the rocket and avoids the directcontact between the heat fluid and the rocket body- ensures the evacuation of the infrared radiation, generated by the heat fluid flowing overthe shield because of the extended Coanda effect by reflection from the rocket bodysurface.

  15. Turbocharger with variable nozzle having vane sealing surfaces

    Science.gov (United States)

    Arnold, Philippe [Hennecourt, FR; Petitjean, Dominique [Julienrupt, FR; Ruquart, Anthony [Thaon les Vosges, FR; Dupont, Guillaume [Thaon les Vosges, FR; Jeckel, Denis [Thaon les Vosges, FR

    2011-11-15

    A variable nozzle for a turbocharger includes a plurality of vanes rotatably mounted on a nozzle ring and disposed in a nozzle flow path defined between the nozzle ring and an opposite nozzle wall. Either or both of the faces of the nozzle ring and nozzle wall include(s) at least one step that defines sealing surfaces positioned to be substantially abutted by airfoil surfaces of the vanes in the closed position of the vanes and to be spaced from the airfoil surfaces in positions other than the closed position. This substantial abutment between the airfoil surfaces and the sealing surfaces serves to substantially prevent exhaust gas from leaking past the ends of the airfoil portions. At the same time, clearances between the nozzle ring face and the end faces of the airfoil portions can be sufficiently large to prevent binding of the vanes under all operating conditions.

  16. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Science.gov (United States)

    Greene, William D. (Inventor)

    2009-01-01

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

  17. Gas dynamic virtual nozzle for generation of microscopic droplet streams

    Energy Technology Data Exchange (ETDEWEB)

    DePonte, D P; Weierstall, U; Schmidt, K; Warner, J; Starodub, D; Spence, J C H; Doak, R B [Department of Physics, Arizona State University, Tempe, AZ 85287-1504 (United States)], E-mail: dandeponte@gmail.com

    2008-10-07

    As shown by Ganan-Calvo (1998 Phys. Rev. Lett. 80 285-8), a free liquid jet can be compressed in diameter through gas dynamic forces exerted by a coaxially co-flowing gas, obviating the need for a solid nozzle to form a microscopic liquid jet and thereby alleviating the clogging problems that plague conventional droplet sources of small diameter. We describe in this paper a novel form of droplet beam source based on this principle. The source is miniature, robust, dependable, easily fabricated, essentially immune to clogging and eminently suitable for delivery of microscopic liquid droplets, including hydrated biological samples, into vacuum for analysis using vacuum instrumentation. Monodisperse, single-file droplet streams are generated by triggering the device with a piezoelectric actuator.

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

    Science.gov (United States)

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

    2004-08-01

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

  19. Russian Meteorological and Geophysical Rockets of New Generation

    Science.gov (United States)

    Yushkov, V.; Gvozdev, Yu.; Lykov, A.; Shershakov, V.; Ivanov, V.; Pozin, A.; Afanasenkov, A.; Savenkov, Yu.; Kuznetsov, V.

    2015-09-01

    To study the process in the middle and upper atmosphere, ionosphere and near-Earth space, as well as to monitor the geophysical environment in Russian Federal Service for Hydrology and Environmental Monitoring (ROSHYDROMET) the development of new generation of meteorological and geophysical rockets has been completed. The modern geophysical research rocket system MR-30 was created in Research and Production Association RPA "Typhoon". The basis of the complex MR-30 is a new geophysical sounding rocket MN-300 with solid propellant, Rocket launch takes place at an angle of 70º to 90º from the launcher, which is a farm with a guide rail type required for imparting initial rotation rocket. The Rocket is spin stabilized with a spin rate between 5 and 7 Hz. Launch weight is 1564 kg, and the mass of the payload of 50 to 150 kg. MR-300 is capable of lifting up to 300 km, while the area of dispersion points for booster falling is an ellipse with parameters 37x 60 km. The payload of the rocket MN-300 consists of two sections: a sealed, located below the instrument compartment, and not sealed, under the fairing. Block of scientific equipment is formed on the platform in a modular layout. This makes it possible to solve a wide range of tasks and conduct research and testing technologies using a unique environment of space, as well as to conduct technological experiments testing and research systems and spacecraft equipment. New Russian rocket system MERA (MEteorological Rocket for Atmospheric Research) belongs to so called "dart" technique that provide lifting of small scientific payload up to altitude 100 km and descending with parachute. It was developed at Central Aerological Observatory jointly with State Unitary Enterprise Instrument Design Bureau. The booster provides a very rapid acceleration to about Mach 5. After the burning phase of the buster the dart is separated and continues ballistic flight for about 2 minutes. The dart carries the instrument payload+ parachute

  20. Numerical Investigation of Twin-Nozzle Rocket Plume Phenomenology, Part 2

    National Research Council Canada - National Science Library

    Ebrahimi, Houshang

    1998-01-01

    .... The Van Leer Flux Splitting option has been successfully implemented into the existing GIFS model and provides a more robust solution scheme, making application of the model more reasonable for engineering applications...

  1. A Computer Code for Fully-Coupled Rocket Nozzle Flows (FULLNOZ)

    Science.gov (United States)

    1975-04-01

    surface (i.e. each integration It would be useful to incorporate an "initializing" scheme which utilizes comb tstion chamber properties as initial...density is greater than the critical electron density. (During the initial stages of the expansion process , where particle tempera- tures are very high it...34iW to19Cs*4909too xs *d99$900 wool ?* 0. SeFC16, .t) .6?900 1, 3x *,30?%I0 to 41,171 0I. 9"CI ,."v *?’o.9 A3 qhbs99r.oo, v.U118 0.1 ,t It Od Cs Sol-C

  2. Effect of Injector Nozzle Holes on Diesel Engine Performance

    OpenAIRE

    Semin,; Yusof, Mohd Yuzri Mohd; Arof, Aminuddin Md; Shaharudin, Daneil Tomo; Ismail, Abdul Rahim

    2010-01-01

    All of the injector nozzle holes have examined and the results are shown that the seven holes nozzle have provided the best burning result for the fuel in-cylinder burned in any different engine speeds and the best burning is in low speed engine. In engine performance effect, all of the nozzles have examined and the five holes nozzle provided the best result in indicted power, indicated torque and ISFC in any different engine speeds.

  3. EUVS Sounding Rocket Payload

    Science.gov (United States)

    Stern, Alan S.

    1996-01-01

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

  4. Not just rocket science

    Energy Technology Data Exchange (ETDEWEB)

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

    2007-10-15

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

  5. Nuclear rocket engine reactor

    Energy Technology Data Exchange (ETDEWEB)

    Lanin, Anatoly

    2013-07-01

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

  6. Engineering thermal engine rocket adventurer for space nuclear application

    International Nuclear Information System (INIS)

    Nam, Seung H.; Suh, Kune Y.; Kang, Seong G.

    2008-01-01

    The conceptual design for the first-of-a-kind engineering of Thermal Engine Rocket Adventure (TERA) is described. TERA comprising the Battery Omnibus Reactor Integral System (BORIS) as the heat resource and the Space Propulsion Reactor Integral System (SPRIS) as the propulsion system, is one of the advanced Nuclear Thermal Rocket (NTR) engine utilizing hydrogen (H 2 ) propellant being developed at present time. BORIS in this application is an open cycle high temperature gas cooled reactor that has eighteen fuel elements for propulsion and one fuel element for electricity generation and propellant pumping. Each fuel element for propulsion has its own small nozzle. The nineteen fuel elements are arranged into hexagonal prism shape in the core and surrounded by outer Be reflector. The TERA maximum power is 1,000 MW th , specific impulse 1,000 s, thrust 250,000 N, and the total mass is 550 kg including the reactor, turbo pump and auxiliaries. Each fuel element comprises the fuel assembly, moderators, pressure tube and small nozzle. The TERA fuel assembly is fabricated of 93% enriched 1.5 mm (U, Zr, Nb)C wafers in 25.3% voided Square Lattice Honeycomb (SLHC). The H 2 propellant passes through these flow channels. This study is concerned with thermohydrodynamic analysis of the fuel element for propulsion with hypothetical axial power distribution because nuclear analysis of TERA has not been performed yet. As a result, when the power distribution of INSPI's M-SLHC is applied to the fuel assembly, the local heat concentration of fuel is more serious and the pressure of the initial inlet H 2 is higher than those of constant average power distribution applied. This means the fuel assembly geometry of 1.5 mm fuel wafers and 25.3% voided SLHC needs to be changed in order to reduce thermal and mechanical shocks. (author)

  7. Easier Analysis With Rocket Science

    Science.gov (United States)

    2003-01-01

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

  8. SAFE testing nuclear rockets economically

    International Nuclear Information System (INIS)

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

    2003-01-01

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

  9. CFD Analysis On The Performance Of Wind Turbine With Nozzles

    Directory of Open Access Journals (Sweden)

    Chunkyraj Kh

    2015-08-01

    Full Text Available In this paper an effort has been made in dealing with fluid characteristic that enters a converging nozzle and analysis of the nozzle is carried out using Computational Fluid Dynamics package ANSYS WORKBENCH 14.5. The paper is the continuation of earlier work Analytical and Experimental performance evaluation of Wind turbine with Nozzles. First the CFD analysis will be carried out on nozzle in-front of wind turbine where streamline velocity at the exit volume flow rate in the nozzle and pressure distribution across the nozzle will be studied. Experiments were conducted on the Wind turbine with nozzles and the corresponding power output at different air speed and different size of nozzles were calculated. Different shapes and dimensions with special contours and profiles of nozzles were studied. It was observed that the special contour nozzles have superior outlet velocity and low pressure at nozzle exit the design has maximum Kinetic energy. These indicators conclude that the contraction designed with the new profile is a good enhancing of the nozzle performance.

  10. A fundamental study of a variable critical nozzle flow

    International Nuclear Information System (INIS)

    Kim, Jea Hyung; Kim, Heuy Dong; Park, Kyung Am

    2003-01-01

    The mass flow rate of gas flow through critical nozzle depends on the nozzle supply conditions and the cross-sectional area at the nozzle throat. In order that the critical nozzle can be operated at a wide range of supply conditions, the nozzle throat diameter should be controlled to change the flow passage area. This can be achieved by means of a variable critical nozzle. In the present study, both experimental and computational works are performed to develop variable critical nozzle. A cone-cylinder with a diameter of d is inserted into conventional critical nozzle. It can move both upstream and downstream, thereby changing the cross-sectional area of the nozzle throat. Computational work using the axisymmetric, compressible Navier-Stokes equations is carried out to simulate the variable critical nozzle flow. An experiment is performed to measure the mass flow rate through variable critical nozzle. The present computational results are in close agreement with measured ones. The boundary layer displacement and momentum thickness are given as a function of Reynolds number. An empirical equation is obtained to predict the discharge coefficient of variable critical nozzle

  11. Some Problems of Rocket-Space Vehicles' Characteristics co- ordination

    Science.gov (United States)

    Sergienko, Alexander A.

    2002-01-01

    of the XX century suffered a reverse. The designers of the United States' firms and enterprises of aviation and rocket-space industry (Boeing, Rocketdyne, Lockheed Martin, McDonnell Douglas, Rockwell, etc.) and NASA (Marshall Space Flight Center, Johnson Space Center, Langley Research Center and Lewis Research Center and others) could not correctly co-ordinate the characteristics of a propulsion system and a space vehicle for elaboration of the "Single-Stage-To-Orbit" reusable vehicle (SSTO) as an integral whole system, which is would able to inject a payload into an orbit and to return back on the Earth. jet nozzle design as well as the choice of propulsion system characteristics, ensuring the high ballistic efficiency, are considered in the present report. The efficiency criterions for the engine and launch system parameters optimization are discussed. The new methods of the nozzle block optimal parameters' choice for the satisfaction of the object task of flight are suggested. The family of SSTO with a payload mass from 5 to 20 ton and initial weight under 800 ton is considered.

  12. New atomization nozzle for spray drying

    NARCIS (Netherlands)

    Deventer, H.C. van; Houben, R.J.; Koldeweij, R.B.J.

    2013-01-01

    A new atomization nozzle based on ink jet technology is introduced for spray drying. Application areas are the food and dairy industry, in the first instance, because in these industries the quality demands on the final powders are high with respect to heat load, powder shape, and size distribution.

  13. Clamp and Gas Nozzle for TIG Welding

    Science.gov (United States)

    Gue, G. B.; Goller, H. L.

    1982-01-01

    Tool that combines clamp with gas nozzle is aid to tungsten/inert-gas (TIG) welding in hard-to-reach spots. Tool holds work to be welded while directing a stream of argon gas at weld joint, providing an oxygen-free environment for tungsten-arc welding.

  14. Fabrication of Microglass Nozzle for Microdroplet Jetting

    Directory of Open Access Journals (Sweden)

    Dan Xie

    2015-02-01

    Full Text Available An ejection aperture nozzle is the essential part for all microdrop generation techniques. The diameter size, the flow channel geometry, and fluid impedance are the key factors affecting the ejection capacity. A novel low-cost fabrication method of microglass nozzle involving four steps is developed in this work. In the first heating step, the glass pipette is melted and pulled. Then, the second heating step is to determine the tip cone angle and modify the flow channel geometry. The desired included angle is usually of 30~45 degrees. Fine grind can determine the exact diameter of the hole. Postheating step is the final process and it can reduce the sharpness of the edges of the hole. Micronozzles with hole diameters varying from 30 to 100 µm are fabricated by the homemade inexpensive and easy-to-operate setup. Hydrophobic treating method of microglass nozzle to ensure stable and accurate injection is also introduced in this work. According to the jetting results of aqueous solution, UV curing adhesive, and solder, the fabricated microglass nozzle can satisfy the need of microdroplet jetting of multimaterials.

  15. Microalgal cell disruption via ultrasonic nozzle spraying.

    Science.gov (United States)

    Wang, M; Yuan, W

    2015-01-01

    The objective of this study was to understand the effect of operating parameters, including ultrasound amplitude, spraying pressure, nozzle orifice diameter, and initial cell concentration on microalgal cell disruption and lipid extraction in an ultrasonic nozzle spraying system (UNSS). Two algal species including Scenedesmus dimorphus and Nannochloropsis oculata were evaluated. Experimental results demonstrated that the UNSS was effective in the disruption of microalgal cells indicated by significant changes in cell concentration and Nile red-stained lipid fluorescence density between all treatments and the control. It was found that increasing ultrasound amplitude generally enhanced cell disruption and lipid recovery although excessive input energy was not necessary for best results. The effect of spraying pressure and nozzle orifice diameter on cell disruption and lipid recovery was believed to be dependent on the competition between ultrasound-induced cavitation and spraying-generated shear forces. Optimal cell disruption was not always achieved at the highest spraying pressure or biggest nozzle orifice diameter; instead, they appeared at moderate levels depending on the algal strain and specific settings. Increasing initial algal cell concentration significantly reduced cell disruption efficiency. In all UNSS treatments, the effectiveness of cell disruption and lipid recovery was found to be dependent on the algal species treated.

  16. Design criteria for piping and nozzles program

    International Nuclear Information System (INIS)

    Moore, S.E.; Bryson, J.W.

    1977-01-01

    This report reviews the activities and accomplishments of the Design Criteria for Piping and Nozzles program being conducted by the Oak Ridge National Laboratory for the period July 1, 1975, to September 30, 1976. The objectives of the program are to conduct integrated experimental and analytical stress analysis studies of piping system components and isolated and closely-spaced pressure vessel nozzles in order to confirm and/or improve the adequacy of structural design criteria and analytical methods used to assure the safe design of nuclear power plants. Activities this year included the development of a finite-element program for analyzing two closely spaced nozzles in a cylindrical pressure vessel; a limited-parameter study of vessels with isolated nozzles, finite-element studies of piping elbows, a fatigue test of an out-of-round elbow, summary and evaluation of experimental studies on the elastic-response and fatigue failure of tees, parameter studies on the behavior of flanged joints, publication of fifteen topical reports and papers on various experimental and analytical studies; and the development and acceptance of a number of design rules changes to the ASME Code. 2 figures, 2 tables

  17. Load calculation on the nozzle in a flue gas desulphurization system

    Science.gov (United States)

    Róbert, Olšiak; Zoltán, Fuszko; Zoltán, Csuka

    2017-09-01

    The desulphurization system is used to remove sulfur oxides from exhaust, so-called flue gases through absorbing them via the sprayed suspension. The suspension delivered from the pump system to the atmospheric bi-directional double hollow cone nozzle has the prescribed working pressure. The unknown mechanical load on the solid body of the nozzle is present through the change of moment due to the flow of the suspension through the bi-directional outflow areas [1], [4]. The calculation of the acting forces and torques in the 3 directions was carried out with the methods of computational fluid dynamics (CFD) in the software ANSYS Fluent. The geometric model of the flow areas of the nozzle were created with the methods of reverse engineering. The computational mesh required by the CFD solver was created, and its quality verified with the standard criteria. The used boundary conditions were defined by the hydraulic parameters of the pump system, the properties of the suspension present in the hydraulic system were specified by sample analysis. The post-processed and analyzed results of the CFD calculation, the pressure-field and the velocity magnitudes in particular directions were further used as input parameters at the mechanical analysis of the load on the bi-directional nozzle.

  18. A unique nuclear thermal rocket engine using a particle bed reactor

    Science.gov (United States)

    Culver, Donald W.; Dahl, Wayne B.; McIlwain, Melvin C.

    1992-01-01

    Aerojet Propulsion Division (APD) studied 75-klb thrust Nuclear Thermal Rocket Engines (NTRE) with particle bed reactors (PBR) for application to NASA's manned Mars mission and prepared a conceptual design description of a unique engine that best satisfied mission-defined propulsion requirements and customer criteria. This paper describes the selection of a sprint-type Mars transfer mission and its impact on propulsion system design and operation. It shows how our NTRE concept was developed from this information. The resulting, unusual engine design is short, lightweight, and capable of high specific impulse operation, all factors that decrease Earth to orbit launch costs. Many unusual features of the NTRE are discussed, including nozzle area ratio variation and nozzle closure for closed loop after cooling. Mission performance calculations reveal that other well known engine options do not support this mission.

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

    International Nuclear Information System (INIS)

    Houseman, John; Young, W.S.

    1974-01-01

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

  20. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    Science.gov (United States)

    Bradley, David E.; Mireles, Omar R.; Hickman, Robert R.

    2011-01-01

    Deep space missions with large payloads require high specific impulse (Isp) and relatively high thrust in order to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average Isp. Nuclear thermal rockets (NTR) capable of high Isp thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high temperature hydrogen exposure on fuel elements is limited. The primary concern is the mechanical failure of fuel elements which employ high-melting-point metals, ceramics or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via non-contact RF heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  1. Rocket Engine Innovations Advance Clean Energy

    Science.gov (United States)

    2012-01-01

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

  2. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

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

    2016-01-01

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

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

    National Research Council Canada - National Science Library

    Shelley, J

    2001-01-01

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

  4. Wake effect in rocket observation

    International Nuclear Information System (INIS)

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

    1975-01-01

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

  5. Multi-Rocket Thought Experiment

    Science.gov (United States)

    Smarandache, Florentin

    2014-03-01

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

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

    Science.gov (United States)

    Williams, R. W. (Compiler)

    1996-01-01

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

  7. Karl Poggensee - A widely unknown German rocket pioneer - The early years 1930-1934 - A chronology

    Science.gov (United States)

    Rohrwild, Karlheinz

    2017-09-01

    The rediscovered estate of Karl Poggensee allows to reproduce chronologically his rocket tests of the period 1930-1934 almost completely for the first time. Thrilled by the movie ;The Woman in the Moon; for the idea of space travel, he started as a student of Hinderburg-Polytechnikum (IAO), Oldenburg, to build his first solid-fuel rocket, producing his own propellant charges. Being a coming electrical engineer his main goal was not set up new record heights, but to provide his rockets with automatic measuring instruments, camera and parachute release systems. The optimization of this sequence was his main focus.

  8. Jet-Surface Interaction: High Aspect Ratio Nozzle Test, Nozzle Design and Preliminary Data

    Science.gov (United States)

    Brown, Clifford; Dippold, Vance

    2015-01-01

    The Jet-Surface Interaction High Aspect Ratio (JSI-HAR) nozzle test is part of an ongoing effort to measure and predict the noise created when an aircraft engine exhausts close to an airframe surface. The JSI-HAR test is focused on parameters derived from the Turbo-electric Distributed Propulsion (TeDP) concept aircraft which include a high-aspect ratio mailslot exhaust nozzle, internal septa, and an aft deck. The size and mass flow rate limits of the test rig also limited the test nozzle to a 16:1 aspect ratio, half the approximately 32:1 on the TeDP concept. Also, unlike the aircraft, the test nozzle must transition from a single round duct on the High Flow Jet Exit Rig, located in the AeroAcoustic Propulsion Laboratory at the NASA Glenn Research Center, to the rectangular shape at the nozzle exit. A parametric nozzle design method was developed to design three low noise round-to-rectangular transitions, with 8:1, 12:1, and 16: aspect ratios, that minimizes flow separations and shocks while providing a flat flow profile at the nozzle exit. These designs validated using the WIND-US CFD code. A preliminary analysis of the test data shows that the actual flow profile is close to that predicted and that the noise results appear consistent with data from previous, smaller scale, tests. The JSI-HAR test is ongoing through October 2015. The results shown in the presentation are intended to provide an overview of the test and a first look at the preliminary results.

  9. Rocket Science 101 Interactive Educational Program

    Science.gov (United States)

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

    2007-01-01

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

  10. Identification of Noise Sources During Rocket Engine Test Firings and a Rocket Launch Using a Microphone Phased-Array

    Science.gov (United States)

    Panda, Jayanta; Mosher, Robert N.; Porter, Barry J.

    2013-01-01

    A 70 microphone, 10-foot by 10-foot, microphone phased array was built for use in the harsh environment of rocket launches. The array was setup at NASA Wallops launch pad 0A during a static test firing of Orbital Sciences' Antares engines, and again during the first launch of the Antares vehicle. It was placed 400 feet away from the pad, and was hoisted on a scissor lift 40 feet above ground. The data sets provided unprecedented insight into rocket noise sources. The duct exit was found to be the primary source during the static test firing; the large amount of water injected beneath the nozzle exit and inside the plume duct quenched all other sources. The maps of the noise sources during launch were found to be time-dependent. As the engines came to full power and became louder, the primary source switched from the duct inlet to the duct exit. Further elevation of the vehicle caused spilling of the hot plume, resulting in a distributed noise map covering most of the pad. As the entire plume emerged from the duct, and the ondeck water system came to full power, the plume itself became the loudest noise source. These maps of the noise sources provide vital insight for optimization of sound suppression systems for future Antares launches.

  11. Rocket Science at the Nanoscale.

    Science.gov (United States)

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

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

  12. Reactor vessel nozzle cracks: a photoelastic study

    International Nuclear Information System (INIS)

    Smith, C.W.

    1979-01-01

    A method consisting of a marriage between the ''frozen stress'' photoelastic approach and the local stress field equations of linear elastic fracture mechanics for estimating stress intensity factor distributions in three dimensional, finite cracked body problems is reviewed and extensions of the method are indicated. The method is then applied to the nuclear reactor vessel nozzle corner crack problem for both Intermediate Test Vessel and Boiling Water Reactor geometries. Results are compared with those of other investigators. 35 refs

  13. Combustion Performance of a Staged Hybrid Rocket with Boron addition

    Science.gov (United States)

    Lee, D.; Lee, C.

    2018-04-01

    In this paper, the effect of boron on overall system specific impulse was investigated. Additionally, a series of combustion tests was carried out to analyze and evaluate the effect of boron addition on O/F variation and radial temperature profiles. To maintain the hybrid rocket engine advantages, upper limit of boron contents in solid fuel was set to be 10 wt%. The results also suggested that, when adding boron to solid fuel, it helped to provide more uniform radial temperature distribution and also to increase specific impulse by 3.2%.

  14. Lymphocytes on sounding rocket flights.

    Science.gov (United States)

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

    1994-05-01

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

  15. Consort 1 sounding rocket flight

    Science.gov (United States)

    Wessling, Francis C.; Maybee, George W.

    1989-01-01

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

  16. Pegasus Rocket Model

    Science.gov (United States)

    1996-01-01

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

  17. Nozzle flow calculation for real gases

    International Nuclear Information System (INIS)

    Bier, K.; Ehrler, F.; Hartz, U.; Kissau, G.

    1977-01-01

    The flow of CHF 2 Cl vapor (refrigerant R 22) through a Laval nozzle of annular geometry has been investigated in the region near the saturation line with stagnation pressures up to 85 per cent of the critical pressure. Static pressure profiles measured along the nozzle axis were found in good agreement with profiles calculated for one-dimensional isentropic flow of the real gas the thermal properties of which were derived from an equation of state proposed previously by Rombusch. Minor deviations between measured and calculated static pressure curves occur in the supersonic part of the mozzle, especially when supersaturated states of the vapour are passed. These deviations can be attributed to uncertainties in the calculation of the enthalpy and to a small influence of the static pressure probe. An additional investigation was concerned with an approximate calculation of the nozzle flow of real gases. In this approximation the well known relations of ideal gas dynamics are applied, the ratio of specific heats for the ideal gas being replaced, however, by a suitably adapted isentropic exponent, which can be determined e.g. from measured values of the Laval pressure or of the mass flow. For pressure ratios p/po between 1 and approximately 0.1, corresponding to Mach numbers up to approximately 2.2, all the interesting properties of the investigated flow of CHF 2 Cl vapour are approximated within a few per cent. (orig.) [de

  18. Coherent structures in a supersonic complex nozzle

    Science.gov (United States)

    Magstadt, Andrew; Berry, Matthew; Glauser, Mark

    2016-11-01

    The jet flow from a complex supersonic nozzle is studied through experimental measurements. The nozzle's geometry is motivated by future engine designs for high-performance civilian and military aircraft. This rectangular jet has a single plane of symmetry, an additional shear layer (referred to as a wall jet), and an aft deck representative of airframe integration. The core flow operates at a Mach number of Mj , c = 1 . 6 , and the wall jet is choked (Mj , w = 1 . 0). This high Reynolds number jet flow is comprised of intense turbulence levels, an intricate shock structure, shear and boundary layers, and powerful corner vortices. In the present study, stereo PIV measurements are simultaneously sampled with high-speed pressure measurements, which are embedded in the aft deck, and far-field acoustics in the anechoic chamber at Syracuse University. Time-resolved schlieren measurements have indicated the existence of strong flow events at high frequencies, at a Strouhal number of St = 3 . 4 . These appear to result from von Kàrmàn vortex shedding within the nozzle and pervade the entire flow and acoustic domain. Proper orthogonal decomposition is applied on the current data to identify coherent structures in the jet and study the influence of this vortex street. AFOSR Turbulence and Transition Program (Grant No. FA9550-15-1-0435) with program managers Dr. I. Leyva and Dr. R. Ponnappan.

  19. Head spray nozzle in reactor pressure vessel

    International Nuclear Information System (INIS)

    Hatano, Shun-ichi.

    1990-01-01

    In a reactor pressure vessel of a BWR type reactor, a head spray nozzle is used for cooling the head of the pressure vessel and, in view of the thermal stresses, it is desirable that cooling is applied as uniformly as possible. A conventional head spray is constituted by combining full cone type nozzles. Since the sprayed water is flown down upon water spraying and the sprayed water in the vertical direction is overlapped, the flow rate distribution has a high sharpness to form a shape as having a maximum value near the center and it is difficult to obtain a uniform flow rate distribution in the circumferential direction. Then, in the present invention, flat nozzles each having a spray water cross section of laterally long shape, having less sharpness in the circumferential distribution upon spraying water to the inner wall of the pressure vessel and having a wide angle of water spray are combined, to make the flow rate distribution of spray water uniform in the inner wall of the pressure vessel. Accordingly, the pressure vessel can be cooled uniformly and thermal stresses upon cooling can be decreased. (N.H.)

  20. Stress analysis of PCV nozzle junction

    International Nuclear Information System (INIS)

    Uchiyama, Shoichi; Oikawa, Tsuneo; Hoshino, Seizo

    1976-01-01

    Most of various pressure vessels comprise each one cylindrical shell and one or more nozzles. In this study, in order to analyze the stress in the structures of this type as minutely and exactly as possible, the program for stress analysis by the finite element method was made, which is required for the strength analysis for three-dimensional structures. Especially, the problem of the stress distribution around nozzle junctions was solved theoretically with the program. The program for the analysis developed in this study is provided with various functions, such as the input generator for cylindrical, conical and spherical shells, and plotter, and is very covenient. The accuracy of analysis is very good. The method of analysis and the calculation of the rigidity matrices for the deformation in plane and bending are explained. The result of the stress analysis around the nozzle junctions of a containment vessel with this program was in good agreement with experimental data and the result with SAP-4 code, therefore the propriety of the calculated result with this program was proved. Also calculations were carried out on three cases, namely a flat plate fixed at one end with distributed load, a cylinder fixed at one end with internal pressure, and an I-beam fixed at one end with concentrated load. The calculated results agreed well with theoretical solutions in all cases. (Kako, I.)

  1. Flow energy piezoelectric bimorph nozzle harvester

    Science.gov (United States)

    Sherrit, Stewart; Lee, Hyeong Jae; Walkemeyer, Phillip; Hasenoehrl, Jennifer; Hall, Jeffrey L.; Colonius, Tim; Tosi, Luis Phillipe; Arrazola, Alvaro; Kim, Namhyo; Sun, Kai; Corbett, Gary

    2014-04-01

    There is a need for a long-life power generation scheme that could be used downhole in an oil well to produce 1 Watt average power. There are a variety of existing or proposed energy harvesting schemes that could be used in this environment but each of these has its own limitations. The vibrating piezoelectric structure is in principle capable of operating for very long lifetimes (decades) thereby possibly overcoming a principle limitation of existing technology based on rotating turbo-machinery. In order to determine the feasibility of using piezoelectrics to produce suitable flow energy harvesting, we surveyed experimentally a variety of nozzle configurations that could be used to excite a vibrating piezoelectric structure in such a way as to enable conversion of flow energy into useful amounts of electrical power. These included reed structures, spring mass-structures, drag and lift bluff bodies and a variety of nozzles with varying flow profiles. Although not an exhaustive survey we identified a spline nozzle/piezoelectric bimorph system that experimentally produced up to 3.4 mW per bimorph. This paper will discuss these results and present our initial analyses of the device using dimensional analysis and constitutive electromechanical modeling. The analysis suggests that an order-of-magnitude improvement in power generation from the current design is possible.

  2. Pengaruh Jarak dan Posisi Nozzle Terhadap Daya Turbin Pelton

    Directory of Open Access Journals (Sweden)

    Yani Kurniawan

    2017-12-01

    Full Text Available Pelton Turbine is a turbine which use nozzle as officers the direction of a stream water in order to move around of blade turbine. The rotating of turbine blade efected by some parameters such as the distance of the nozzle, position of nozzle, diameter of nozzle, number of nozzle, and the geometry shape of the blade turbine. An experimental study to analyze the affect of distance and position nozzle to Pelton Turbine of performance. The research method used experiment parameter was position of nozzle with three variations, first position is the right side horizontal of bottom shaft turbine, second position is vertical to down direction, and third position is the left side horizontal of upper shaft turbine. The parameter of nozzle distance used five variations was 24 cm, 23 cm, 22 cm, 21 cm, dan 20 cm, which measured from the end of position nozzle to blade turbine. The result shows that the right side horizontal of bottom shaft turbine with distance of nozzle 23 cm had the maximum performance to produce a power 125 Watt with the rotation of shaft turbine 263 rpm.

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2010-03-01

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

  4. Design study of laser fusion rocket

    International Nuclear Information System (INIS)

    Nakashima, Hideki; Shoyama, Hidetoshi; Kanda, Yukinori

    1991-01-01

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

  5. Launch Excitement with Water Rockets

    Science.gov (United States)

    Sanchez, Juan Carlos; Penick, John

    2007-01-01

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

  6. Measuring Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  7. Development of top nozzle for Korean standard LWR fuel

    Energy Technology Data Exchange (ETDEWEB)

    Lee, S. K.; Kim, I. K.; Choi, K. S.; Kim, Y. H.; Lee, J. N.; Kim, H. K. [KNFC, Taejon (Korea, Republic of)

    2001-10-01

    Performance evaluation was executed for each component and its assembly for the deduced Top Nozzles to develop the new Top Nozzle for LWR. This new Top Nozzle is composed of the optimum components among the derived Top Nozzles that have been evaluated in the viewpoint of structural integrity, simpleness of dismantle and assembly, manufacturability etc. In this study, the developed Top Nozzle satisfied all the related design criteria. In special, it makes fuel repair time reduced by assembling and disassembling itself as one body, and improves Fuel Assembly holddown ability by revising the design parameters of its spring and the structural integrity through the betterment of its geometrical shpae of Flange and Holddown Plate as compared with the existing LWR Top Nozzles.

  8. Infrasound and Seismic Recordings of Rocket Launches from Kennedy Space Center, 2016-2017

    Science.gov (United States)

    McNutt, S. R.; Thompson, G.; Brown, R. G.; Braunmiller, J.; Farrell, A. K.; Mehta, C.

    2017-12-01

    We installed a temporary 3-station seismic-infrasound network at Kennedy Space Center (KSC) in February 2016 to test sensor calibrations and train students in field deployment and data acquisitions techniques. Each station featured a single broadband 3-component seismometer and a 3-element infrasound array. In May 2016 the network was scaled back to a single station due to other projects competing for equipment. To date 8 rocket launches have been recorded by the infrasound array, as well as 2 static tests, 1 aborted launch and 1 rocket explosion (see next abstract). Of the rocket launches recorded 4 were SpaceX Falcon-9, 2 were ULA Atlas-5 and 2 were ULA Delta-IV. A question we attempt to answer is whether the rocket engine type and launch trajectory can be estimated with appropriate travel-time, amplitude-ratio and spectral techniques. For example, there is a clear Doppler shift in seismic and infrasound spectrograms from all launches, with lower frequencies occurring later in the recorded signal as the rocket accelerates away from the array. Another question of interest is whether there are relationships between jet noise frequency, thrust and/or nozzle velocity. Infrasound data may help answer these questions. We are now in the process of deploying a permanent seismic and infrasound array at the Astronaut Beach House. 10 more rocket launches are schedule before AGU. NASA is also conducting a series of 33 sonic booms over KSC beginning on Aug 21st. Launches and other events at KSC have provided rich sources of signals that are useful to characterize and gain insight into physical processes and wave generation from man-made sources.

  9. Mounting apparatus for a nozzle guide vane assembly

    Science.gov (United States)

    Boyd, Gary L.; Shaffer, James E.

    1995-01-01

    The present invention provides a ceramic nozzle guide assembly with an apparatus for mounting it to a metal nozzle case that includes an intermediate ceramic mounting ring. The mounting ring includes a plurality of projections that are received within a plurality of receptacles formed in the nozzle case. The projections of the mounting ring are secured within the receptacles by a ceramic retainer that allows contact between the two components only along arcuate surfaces thus eliminating sliding contact between the components.

  10. Fluidized-bed calciner with combustion nozzle and shroud

    International Nuclear Information System (INIS)

    Wielang, J.A.; Palmer, W.B.; Kerr, W.B.

    1977-01-01

    A nozzle employed as a burner within a fluidized bed is coaxially enclosed within a tubular shroud that extends beyond the nozzle length into the fluidized bed. The open-ended shroud portion beyond the nozzle end provides an antechamber for mixture and combustion of atomized fuel with an oxygen-containing gas. The arrangement provides improved combustion efficiency and excludes bed particles from the high-velocity, high-temperature portions of the flame to reduce particle attrition. 4 claims, 2 figures

  11. Variable volume combustor with aerodynamic fuel flanges for nozzle mounting

    Science.gov (United States)

    McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward; Ostebee, Heath Michael

    2016-09-20

    The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and for providing the flow of fuel therethrough. The fuel injection system also may include a number of aerodynamic fuel flanges connecting the micro-mixer fuel nozzles and the support struts.

  12. Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

    Science.gov (United States)

    Bui, Trong

    2010-01-01

    An axisymmetric full Navier-Stokes computational fluid dynamics (CFD) study was conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock (LaNCETS) research airplane, was considered. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature, similar to what was observed in the LaNCETS flight data, is observed for the highly underexpanded nozzle flow. The CFD results provide a detailed description of the nozzle flow physics involved in the LaNCETS nozzle at different nozzle expansion conditions and help in interpreting LaNCETS flight data as well as in the eventual CFD analysis of a full LaNCETS aircraft. The current study also provided important information on proper modeling of the LaNCETS aircraft nozzle. The primary objective of the current CFD research effort was to support the LaNCETS flight research data analysis effort by studying the detailed nozzle exhaust jet plume s imperfect expansion effects on the sonic boom signature of a supersonic aircraft. Figure 1 illustrates the primary flow physics present in the interaction between the exhaust jet plume shock and the sonic boom coming off of an axisymmetric body in supersonic flight. The steeper tail shock from highly expanded jet plume reduces the dip of the sonic boom N-wave signature. A structured finite-volume compressible full Navier-Stokes CFD code was used in the current study. This approach is not limited by the simplifying assumptions inherent in previous sonic boom analysis efforts. Also, this study was the first known jet plume sonic boom CFD study in which the full viscous nozzle flow field was modeled, without

  13. Experimental study of subsonic microjet escaping from a rectangular nozzle

    Science.gov (United States)

    Aniskin, V. M.; Maslov, A. A.; Mukhin, K. A.

    2016-10-01

    The first experiments on the subsonic laminar microjets escaping from the nozzles of rectangular shape are carried out. The nozzle size is 83.3x3823 microns. Reynolds number calculated by the nozzle height and the average flow velocity at the nozzle exit ranged from 58 to 154. The working gas was air at room temperature. The velocity decay and velocity fluctuations along the center line of the jet are determined. The fundamental difference between the laminar microjets characteristics and subsonic turbulent jets of macro size is shown. Based on measurements of velocity fluctuations it is shown the presence of laminar-turbulent transition in microjets and its location is determined.

  14. Heat and fluid flow properties of circular impinging jet with a low nozzle to plate spacing. Improvement by nothched nozzle; Nozzle heibankan kyori ga chiisai baai no enkei shototsu funryu no ryudo dennetsu tokusei. Kirikaki nozzle ni yoru kaizen kojo

    Energy Technology Data Exchange (ETDEWEB)

    Shakouchih, T. [Mie University, Mie (Japan). Faculty of Engineering; Matsumoto, A.; Watanabe, A.

    2000-10-25

    It is well known that as decreasing the nozzle to plate spacing considerably the heat transfer coefficient of circular impinging jet, which impinges to the plate normally, increases remarkably. At that time, the flow resistance of nozzle-plate system also increases rapidly. In this study, in order to reduce the flow resistance and to enhance the heat transfer coefficient of the circular impinging jet with a considerably low nozzle to plate spacing, a special nozzle with notches is proposed, and considerable improvement of the flow and heat transfer properties are shown. The mechanism of enhancement of the heat transfer properties is also discussed. (author)

  15. A research on polyether glycol replaced APCP rocket propellant

    Science.gov (United States)

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

    2017-08-01

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

  16. Development of high performance hybrid rocket fuels

    Science.gov (United States)

    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

  17. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Science.gov (United States)

    2010-01-01

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

  18. BWR feedwater nozzle and control-rod-drive return line nozzle cracking

    International Nuclear Information System (INIS)

    Anon.

    1981-01-01

    In its 1978 Annual Report to Congress, the Nuclear Regulatory Commission identified as an unresolved safety issue the appearance of cracks in feedwater nozzles at boiling-water reactors (BWRs). Later similar cracking, detected in return water lines for control-rod-drive systems at BWRs, was designated Part II of the issue. This article outlines the resolution of these cracking problems

  19. Nozzle evaluation for Project W-314

    International Nuclear Information System (INIS)

    Galbraith, J.D.

    1998-01-01

    Revisions to the waste transfer system piping to be implemented by Project W-314 will eliminate the need to access a majority of interfarm jumper connections associated with specific process pits. Additionally, connections that formerly facilitated waste transfers from the Plutonium-Uranium Extraction (PUREX) Plant are no longer required. This document identified unneeded process pit jumper connections, describes former designated routing, denotes current status (i.e., open or blanked), and recommends appropriate disposition for all. Blanking of identified nozzles should be accomplished by Project W-314 upon installation of jumpers and acceptance by Tank Waste Remediation System (TWRS) Tank Farm Operations

  20. Bottom nozzle of a LWR fuel assembly

    International Nuclear Information System (INIS)

    Leroux, J.C.

    1991-01-01

    The bottom nozzle consists of a transverse element in form of box having a bending resistant grid structure which has an outer peripheral frame of cross-section corresponding to that of the fuel assembly and which has walls defining large cells. The transverse element has a retainer plate with a regular array of openings. The retainer plate is fixed above and parallel to the grid structure with a spacing in order to form, between the grid structure and the retainer plate a free space for tranquil flow of cooling water and for debris collection [fr

  1. Airfoil shape for a turbine nozzle

    Science.gov (United States)

    Burdgick, Steven Sebastian; Patik, Joseph Francis; Itzel, Gary Michael

    2002-01-01

    A first-stage nozzle vane includes an airfoil having a profile according to Table I. The annulus profile of the hot gas path is defined in conjunction with the airfoil profile and the profile of the inner and outer walls by the Cartesian coordinate values given in Tables I and II, respectively. The airfoil is a three-dimensional bowed design, both in the airfoil body and in the trailing edge. The airfoil is steam and air-cooled by flowing cooling mediums through cavities extending in the vane between inner and outer walls.

  2. Study of nozzle deposit formation mechanism for direct injection gasoline engines; Chokufun gasoline engine yo nozzle no deposit seisei kaiseki

    Energy Technology Data Exchange (ETDEWEB)

    Kinoshita, M; Saito, A [Toyota Central Research and Development Labs., Inc., Aichi (Japan); Matsushita, S [Toyota Motor Corp., Aichi (Japan); Shibata, H [Nippon Soken, Inc., Tokyo (Japan); Niwa, Y [Denso Corp., Aichi (Japan)

    1997-10-01

    Nozzles in fuel injectors for direct injection gasoline engines are exposed to high temperature combustion gases and soot. In such a rigorous environment, it is a fear that fuel flow rate changes in injectors by deposit formation on nozzles. Fundamental factors of nozzle deposit formation were investigated through injector bench tests and engine dynamometer tests. Deposit formation processes were observed by SEM through engine dynamometer tests. The investigation results reveal nozzle deposit formation mechanism and how to suppress the deposit. 4 refs., 8 figs., 3 tabs.

  3. UEDGE Simulations for Power and Particle Flow Analysis of FRC Rocket

    Science.gov (United States)

    Zheng, Fred; Evans, Eugene S.; McGreivy, Nick; Kaptanoglu, Alan; Izacard, Olivier; Cohen, Samuel A.

    2017-10-01

    The field-reversed configuration (FRC) is under consideration for use in a direct fusion drive (DFD) rocket propulsion system for future space missions. To achieve a rocket configuration, the FRC is embedded within an asymmetric magnetic mirror, in which one end is closed and contains a gas box, and the other end is open and incorporates a magnetic nozzle. Neutral deuterium is injected into the gas box, and flows through the scrape-off layer (SOL) around the core plasma and out the magnetic nozzle, both cooling the core and serving as propellant. Previous studies have examined a range of operating conditions for the SOL of a DFD using UEDGE, a 2D fluid code; discrepancies on the order of 5% were found during the analysis of overall power balance. This work extends the analysis of the previously-studied SOL geometry by updating boundary conditions and conducting a detailed study of power and particle flows within the simulation with the goals of modeling electrical power generation instead of thrust and achieving higher specific impulse. This work was supported, in part, by DOE Contract Number DE-AC02-09CH11466 and Princeton Environmental Institute.

  4. EURCYL. A program to generate finite element meshes for pressure vessel nozzles

    International Nuclear Information System (INIS)

    De Windt, P.; Reynen, J.

    1974-12-01

    EURCYL is a program dealing with the automatic generation of finite element meshes for pressure vessel nozzles, using isoparametric elements with 8, 20 or 32 nodes. Options exist to generate BWR nozzles as well as PWR nozzles

  5. Passive Rocket Diffuser Theory: A Re-Examination of Minimum Second Throat Size

    Science.gov (United States)

    Jones, Daniel R.

    2016-01-01

    Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure during testing without using active control systems. Among the most critical design parameters is the relative area of the diffuser throat to that of the nozzle throat. A smaller second throat is generally desirable because it decreases the stagnation-to-ambient pressure ratio the diffuser requires for nominal operation. There is a limit, however. Below a certain size, the second throat can cause pressure buildup within the diffuser and prevent it from reaching the start condition that protects the nozzle from side-load damage. This paper presents a method for improved estimation of the minimum second throat area which enables diffuser start. The new 3-zone model uses traditional quasi-one-dimensional compressible flow theory to approximate the structure of two distinct diffuser flow fields observed in Computational Fluid Dynamics (CFD) simulations and combines them to provide a less-conservative estimate of the second throat size limit. It is unique among second throat sizing methods in that it accounts for all major conical nozzle and second throat diffuser design parameters within its limits of application. The performance of the 3-zone method is compared to the historical normal shock and force balance methods, and verified against a large number of CFD simulations at specific heat ratios of 1.4 and 1.25. Validation is left as future work, and the model is currently intended to function only as a first-order design tool.

  6. Grit blasting nozzle fabricated from mild tool steel proves satisfactory

    Science.gov (United States)

    Mc Farland, J. E.; Turbitt, B.

    1966-01-01

    Dry blasting with glass beads through a nozzle assembly descales both the outside and inside surfaces of tubes of Inconel 718 used for the distribution of gaseous oxygen. The inside of the nozzle is coated with polyurethane and the deflector with a commercially available liquid urethane rubber.

  7. Numerical analysis of choked converging nozzle flows with surface ...

    Indian Academy of Sciences (India)

    Choked converging nozzle flow and heat transfer characteristics are numerically investigated by means of a recent computational model that integrates the axisymmetric continuity, state, momentum and energy equations. To predict the combined effects of nozzle geometry, friction and heat transfer rates, analyses are ...

  8. Multi-orifice deposition nozzle for additive manufacturing

    Science.gov (United States)

    Lind, Randall F.; Post, Brian K.; Cini, Colin L.

    2017-11-21

    An additive manufacturing extrusion head includes a nozzle for accepting and depositing a heated material onto a work surface and/or part. The nozzle includes a valve body and an internal poppet body moveable between positions to permit deposition of at least two bead sizes of heated material onto a work surface and/or part.

  9. Noise from Aft Deck Exhaust Nozzles: Differences in Experimental Embodiments

    Science.gov (United States)

    Bridges, James E.

    2014-01-01

    Two embodiments of a rectangular nozzle on an aft deck are compared. In one embodiment the lower lip of the nozzle was extended with the sidewalls becoming triangles. In a second embodiment a rectangular nozzle was fitted with a surface that fit flush to the lower lip and extended outward from the sides of the nozzle, approximating a semi-infinite plane. For the purpose of scale-model testing, making the aft deck an integral part of the nozzle is possible for relatively short deck lengths, but a separate plate model is more flexible, accounts for the expanse of deck to the sides of the nozzle, and allows the nozzle to stand off from the deck. Both embodiments were tested and acoustic far-field results were compared. In both embodiments the extended deck introduces a new noise source, but the amplitude of the new source was dependent upon the span (cross-stream dimension) of the aft deck. The noise increased with deck length (streamwise dimension), and in the case of the beveled nozzle it increased with increasing aspect ratio. In previous studies of slot jets in wings it was noted that the increased noise from the extended aft deck appears as a dipole at the aft deck trailing edge, an acoustic source type with different dependence on velocity than jet mixing noise. The extraneous noise produced by the aft deck in the present studies also shows this behavior both in directivity and in velocity scaling.

  10. Combustor nozzle for a fuel-flexible combustion system

    Science.gov (United States)

    Haynes, Joel Meier [Niskayuna, NY; Mosbacher, David Matthew [Cohoes, NY; Janssen, Jonathan Sebastian [Troy, NY; Iyer, Venkatraman Ananthakrishnan [Mason, OH

    2011-03-22

    A combustor nozzle is provided. The combustor nozzle includes a first fuel system configured to introduce a syngas fuel into a combustion chamber to enable lean premixed combustion within the combustion chamber and a second fuel system configured to introduce the syngas fuel, or a hydrocarbon fuel, or diluents, or combinations thereof into the combustion chamber to enable diffusion combustion within the combustion chamber.

  11. Ultrasonic pattern recognition study of feedwater nozzle inner radius indication

    International Nuclear Information System (INIS)

    Yoneyama, H.; Takama, S.; Kishigami, M.; Sasahara, T.; Ando, H.

    1983-01-01

    A study was made to distinguish defects on feed-water nozzle inner radius from noise echo caused by stainless steel cladding by using ultrasonic pattern recognition method with frequency analysis technique. Experiment has been successfully performed on flat clad plates and nozzle mock-up containing fatigue cracks and the following results which shows the high capability of frequency analysis technique are obtained

  12. Finite element analysis of inclined nozzle-plate junctions

    International Nuclear Information System (INIS)

    Dixit, K.B.; Seth, V.K.; Krishnan, A.; Ramamurthy, T.S.; Dattaguru, B.; Rao, A.K.

    1979-01-01

    Estimation of stress concentration at nozzle to plate or shell junctions is a significant problem in the stress analysis of nuclear reactors. The topic is a subject matter of extensive investigations and earlier considerable success has been reported on analysis for the cases when the nozzle is perpendicular to the plate or is radial to the shell. Analytical methods for the estimation of stress concentrations for the practical situations when the intersecting nozzle is inclined to the plate or is non-radial to the shell is rather scanty. Specific complications arise in dealing with the junction region when the nozzle with circular cross-section meets the non-circular cut-out on the plate or shell. In this paper a finite element analysis is developed for inclined nozzles and results are presented for nozzle-plate junctions. A method of analysis is developed with a view to achieving simultaneously accuracy of results and simplicity in the choice of elements and their connectivity. The circular nozzle is treated by axisymmetric conical shell elements. The nozzle portion in the region around the junction and the flat plate is dealt with by triangular flat shell elements. Special transition elements are developed for joining the flat shell elements with the axisymmetric elements under non-axisymmetric loading. A substructure method of analysis is adopted which achieves considerable economy in handling the structure and also conveniently combines the different types of elements in the structure. (orig.)

  13. Volatiles from solids

    Energy Technology Data Exchange (ETDEWEB)

    Loughrey, C T

    1939-08-24

    To remove volatiles from solids, such as oil shale, gases, and/or vapours are passed through a mass of the materials, the vapours and gases separated, and the vapours condensed. The volatile-containing solid materials are fed to a retort, and a shaft is driven to rotate an impeller so as to displace the liquid and create a vortex tube, which draws in gas from the atmosphere through an intake, twyer, interstices in the material in the retort, a conduit, chamber, tubes, another chamber and cylinder. This gas is carried outwardly and upwardly by the vortices in the liquid and is carried to discharge through three conduits. The vapours entrained by the gas are part condensed in the liquid and the remainder directed to a condenser. Steam may be delivered to the twyer through a nozzle of a pipe, with or without air, and combustible hydrocarbon fuel may be fed through the burner nozzle or solid fuel may be directed from feeder and combusted in the twyer.

  14. Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

    Science.gov (United States)

    de León, Pablo

    2009-12-01

    One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

  15. Study on steam pressure characteristics in various types of nozzles

    Science.gov (United States)

    Firman; Anshar, Muhammad

    2018-03-01

    Steam Jet Refrigeration (SJR) is one of the most widely applied technologies in the industry. The SJR system was utilizes residual steam from the steam generator and then flowed through the nozzle to a tank that was containing liquid. The nozzle converts the pressure energy into kinetic energy. Thus, it can evaporate the liquid briefly and release it to the condenser. The chilled water, was produced from the condenser, can be used to cool the product through a heat transfer process. This research aims to study the characteristics of vapor pressure in different types of nozzles using a simulation. The Simulation was performed using ANSYS FLUENT software for nozzle types such as convergent, convrgent-parallel, and convergent-divergent. The results of this study was presented the visualization of pressure in nozzles and was been validated with experiment data.

  16. TMI-2 instrument nozzle examinations at Argonne National Laboratory

    International Nuclear Information System (INIS)

    Neimark, L.A.; Shearer, T.L.; Purohit, A.; Hins, A.G.

    1993-09-01

    Six of the 14 instrument-penetration-tube nozzles removed from the lower head of TMI-2 were examined to identify damage mechanisms, provide insight to the fuel relocation scenario, and provide input data to the margin-to-failure analysis. Visual inspection, gamma scanning, metallography, microhardness measurements, and scanning electron microscopy were used to obtain the desired information. The results showed varying degrees of damage to the lower head nozzles, from ∼50% melt-off to no damage at all to near-neighbor nozzles. The elevations of nozzle damage suggested that the lower elevations (near the lower head) were protected from molten fuel, apparently by an insulating layer of fuel debris. The pattern of nozzle damage was consistent with fuel movement toward the hot-spot location identified in the vessel wall. Evidence was found for the existence of a significant quantity of control assembly debris on the lower head before the massive relocation of fuel occurred

  17. Novel design for transparent high-pressure fuel injector nozzles

    Science.gov (United States)

    Falgout, Z.; Linne, M.

    2016-08-01

    The efficiency and emissions of internal combustion (IC) engines are closely tied to the formation of the combustible air-fuel mixture. Direct-injection engines have become more common due to their increased practical flexibility and efficiency, and sprays dominate mixture formation in these engines. Spray formation, or rather the transition from a cylindrical liquid jet to a field of isolated droplets, is not completely understood. However, it is known that nozzle orifice flow and cavitation have an important effect on the formation of fuel injector sprays, even if the exact details of this effect remain unknown. A number of studies in recent years have used injectors with optically transparent nozzles (OTN) to allow observation of the nozzle orifice flow. Our goal in this work is to design various OTN concepts that mimic the flow inside commercial injector nozzles, at realistic fuel pressures, and yet still allow access to the very near nozzle region of the spray so that interior flow structure can be correlated with primary breakup dynamics. This goal has not been achieved until now because interior structures can be very complex, and the most appropriate optical materials are brittle and easily fractured by realistic fuel pressures. An OTN design that achieves realistic injection pressures and grants visual access to the interior flow and spray formation will be explained in detail. The design uses an acrylic nozzle, which is ideal for imaging the interior flow. This nozzle is supported from the outside with sapphire clamps, which reduces tensile stresses in the nozzle and increases the nozzle's injection pressure capacity. An ensemble of nozzles were mechanically tested to prove this design concept.

  18. Numerical simulation of a liquid propellant rocket motor

    Science.gov (United States)

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

    2001-03-01

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

  19. Two-dimensional motions of rockets

    International Nuclear Information System (INIS)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

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

    Science.gov (United States)

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

    2002-01-01

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

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

    Science.gov (United States)

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

    2016-12-01

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

  2. A Basic Study on the Ejection of ICI Nozzle under Severe Accidents

    Energy Technology Data Exchange (ETDEWEB)

    Cho, Jong Rae; Bae, Ji Hoon; Bang, Kwang Hyun [Korea Maritime and Ocean University, Busan (Korea, Republic of); Park, Jong Woong [Dongguk University, Gyeongju (Korea, Republic of)

    2016-05-15

    Nozzle injection should be blocked because it affect to the environment if its melting core exposes outside. The purpose of this study is to carry out the thermos mechanical analysis due to debris relocation under severe accidents and to predict the nozzle ejection calculated considering the contact between the nozzle and lower head, and the supports of pipe cables. As a result of analyzing process of severe accidents, there was melting reaction between nozzle and the lower head. In this situation, we might predict the non-uniform contact region of nozzle hole of lower head and nozzle outside, delaying ejection of nozzles. But after melting, the average remaining length of the nozzle was 120mm and the maximum vertical displacement of lower nozzle near the weld is 3.3mm so there would be no nozzle this model, because the cable supports restrains the vertical displacement of nozzle.

  3. PULSED MOLECULAR BEAM PRODUCTION WITH NOZZLES

    Energy Technology Data Exchange (ETDEWEB)

    Hagena, Otto-Friedrich

    1963-05-15

    Molecular beam experiments that can be carried out in pulsed operation may be performed at considerably reduced expense for apparatus if, for pulse generation, the gas supply to the beam production system is interrupted as opposed to the usual steady molecular beam. This technique is studied by measuring intensity vs time of molecular beam impulses of varying length, how fast and through which intermediate states the initial intensity of the impulse attains equilibrium, and in which way the intensity of the molecular-beam impulse is affected by the pulse length and by increasing pressure in the first pressure stage. For production of pulses, a magnetically actuated, quick shutting, valve is used whose scaling area is the inlet cone of the nozzle used for the beam generation. The shortest pulses produced had a pulse length of 1.6 ms. (auth)

  4. Specific decontamination methods: water nozzle, cavitation erosion

    International Nuclear Information System (INIS)

    Boulitrop, D.; Gauchon, J.P.; Lecoffre, Y.

    1984-05-01

    The erosion and decontamination tests carried out in the framework of this study, allowed to specify the fields favourable to the use of the high pressure jet taking into account the determinant parameters that are the pressure and the target-nozzle distance. The previous spraying of gels with chemical reagents (sulfuric acid anf hydrazine) allows to get better decontamination factors. Then, the feasibility study of a decontamination method by cavitation erosion is presented. Gelled compounds for decontamination have been developed; their decontamination quality has been evaluated by comparative contamination tests in laboratory and decontamination tests of samples of materials used in nuclear industry; this last method is adapted to remote handling devices and produces a low quantity of secondary effluents, so it allows to clean high contaminated installation on the site without additional exposure of the personnel [fr

  5. Feedback mechanism for smart nozzles and nebulizers

    Science.gov (United States)

    Montaser, Akbar [Potomac, MD; Jorabchi, Kaveh [Arlington, VA; Kahen, Kaveh [Kleinburg, CA

    2009-01-27

    Nozzles and nebulizers able to produce aerosol with optimum and reproducible quality based on feedback information obtained using laser imaging techniques. Two laser-based imaging techniques based on particle image velocimetry (PTV) and optical patternation map and contrast size and velocity distributions for indirect and direct pneumatic nebulizations in plasma spectrometry. Two pulses from thin laser sheet with known time difference illuminate droplets flow field. Charge coupled device (CCL)) captures scattering of laser light from droplets, providing two instantaneous particle images. Pointwise cross-correlation of corresponding images yields two-dimensional velocity map of aerosol velocity field. For droplet size distribution studies, solution is doped with fluorescent dye and both laser induced florescence (LIF) and Mie scattering images are captured simultaneously by two CCDs with the same field of view. Ratio of LIF/Mie images provides relative droplet size information, then scaled by point calibration method via phase Doppler particle analyzer.

  6. Noise Prediction Module for Offset Stream Nozzles

    Science.gov (United States)

    Henderson, Brenda S.

    2011-01-01

    A Modern Design of Experiments (MDOE) analysis of data acquired for an offset stream technology was presented. The data acquisition and concept development were funded under a Supersonics NRA NNX07AC62A awarded to Dimitri Papamoschou at University of California, Irvine. The technology involved the introduction of airfoils in the fan stream of a bypass ratio (BPR) two nozzle system operated at transonic exhaust speeds. The vanes deflected the fan stream relative to the core stream and resulted in reduced sideline noise for polar angles in the peak jet noise direction. Noise prediction models were developed for a range of vane configurations. The models interface with an existing ANOPP module and can be used or future system level studies.

  7. Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion

    Science.gov (United States)

    Rahman, Shamim A.

    2010-01-01

    Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.

  8. Star-grain rocket motor - nonsteady internal ballistics

    Energy Technology Data Exchange (ETDEWEB)

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

    2004-01-01

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

  9. Rotational flow in tapered slab rocket motors

    Science.gov (United States)

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

    2006-10-01

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

  10. Nuclear rockets: High-performance propulsion for Mars

    International Nuclear Information System (INIS)

    Watson, C.W.

    1994-05-01

    A new impetus to manned Mars exploration was introduced by President Bush in his Space Exploration Initiative. This has led, in turn, to a renewed interest in high-thrust nuclear thermal rocket propulsion (NTP). The purpose of this report is to give a brief tutorial introduction to NTP and provide a basic understanding of some of the technical issues in the realization of an operational NTP engine. Fundamental physical principles are outlined from which a variety of qualitative advantages of NTP over chemical propulsion systems derive, and quantitative performance comparisons are presented for illustrative Mars missions. Key technologies are described for a representative solid-core heat-exchanger class of engine, based on the extensive development work in the Rover and NERVA nuclear rocket programs (1955 to 1973). The most driving technology, fuel development, is discussed in some detail for these systems. Essential highlights are presented for the 19 full-scale reactor and engine tests performed in these programs. On the basis of these tests, the practicality of graphite-based nuclear rocket engines was established. Finally, several higher-performance advanced concepts are discussed. These have received considerable attention, but have not, as yet, developed enough credibility to receive large-scale development

  11. Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

    Science.gov (United States)

    Westra, Douglas G.; West, Jeffrey S.

    2014-01-01

    NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

  12. Flame Interactions and Thermoacoustics in Multiple-Nozzle Combustors

    Science.gov (United States)

    Dolan, Brian

    The first major chapter of original research (Chapter 3) examines thermoacoustic oscillations in a low-emission staged multiple-nozzle lean direct injection (MLDI) combustor. This experimental program investigated a relatively practical combustor sector that was designed and built as part of a commercial development program. The research questions are both practical, such as under what conditions the combustor can be safely operated, and fundamental, including what is most significant to driving the combustion oscillations in this system. A comprehensive survey of operating conditions finds that the low-emission (and low-stability) intermediate and outer stages are necessary to drive significant thermoacoustics. Phase-averaged and time-resolved OH* imaging show that dramatic periodic strengthening and weakening of the reaction zone downstream of the low-emission combustion stages. An acoustic modal analysis shows the pressure wave shapes and identifies the dominant thermoacoustic behavior as the first longitudinal mode for this combustor geometry. Finally, a discussion of the likely significant coupling mechanisms is given. Periodic reaction zone behavior in the low-emission fuel stages is the primary contributor to unsteady heat release. Differences between the fuel stages in the air swirler design, the fuel number of the injectors, the lean blowout point, and the nominal operating conditions all likely contribute to the limit cycle behavior of the low-emission stages. Chapter 4 investigates the effects of interaction between two adjacent swirl-stabilized nozzles using experimental and numerical tools. These studies are more fundamental; while the nozzle hardware is the same as the lean direct injection nozzles used in the MLDI combustion concept, the findings are generally applicable to other swirl-stabilized combustion systems as well. Much of the work utilizes a new experiment where the distance between nozzles was varied to change the level of interaction

  13. Controllable Solid Propulsion Combustion and Acoustic Knowledge Base Improvements

    Science.gov (United States)

    McCauley, Rachel; Fischbach, Sean; Fredrick, Robert

    2012-01-01

    Controllable solid propulsion systems have distinctive combustion and acoustic environments that require enhanced testing and analysis techniques to progress this new technology from development to production. In a hot gas valve actuating system, the movement of the pintle through the hot gas exhibits complex acoustic disturbances and flow characteristics that can amplify induced pressure loads that can damage or detonate the rocket motor. The geometry of a controllable solid propulsion gas chamber can set up unique unsteady flow which can feed acoustic oscillations patterns that require characterization. Research in this area aids in the understanding of how best to design, test, and analyze future controllable solid rocket motors using the lessons learned from past government programs as well as university research and testing. This survey paper will give the reader a better understanding of the potentially amplifying affects propagated by a controllable solid rocket motor system and the knowledge of the tools current available to address these acoustic disturbances in a preliminary design. Finally the paper will supply lessons learned from past experiences which will allow the reader to come away with understanding of what steps need to be taken when developing a controllable solid rocket propulsion system. The focus of this survey will be on testing and analysis work published by solid rocket programs and from combustion and acoustic books, conference papers, journal articles, and additionally from subject matter experts dealing currently with controllable solid rocket acoustic analysis.

  14. Application of LBB to a nozzle-pipe interface

    Energy Technology Data Exchange (ETDEWEB)

    Yu, Y.J.; Sohn, G.H.; Kim, Y.J. [and others

    1997-04-01

    Typical LBB (Leak-Before-Break) analysis is performed for the highest stress location for each different type of material in the high energy pipe line. In most cases, the highest stress occurs at the nozzle and pipe interface location at the terminal end. The standard finite element analysis approach to calculate J-Integral values at the crack tip utilizes symmetry conditions when modeling near the nozzle as well as away from the nozzle region to minimize the model size and simplify the calculation of J-integral values at the crack tip. A factor of two is typically applied to the J-integral value to account for symmetric conditions. This simplified analysis can lead to conservative results especially for small diameter pipes where the asymmetry of the nozzle-pipe interface is ignored. The stiffness of the residual piping system and non-symmetries of geometry along with different material for the nozzle, safe end and pipe are usually omitted in current LBB methodology. In this paper, the effects of non-symmetries due to geometry and material at the pipe-nozzle interface are presented. Various LBB analyses are performed for a small diameter piping system to evaluate the effect a nozzle has on the J-integral calculation, crack opening area and crack stability. In addition, material differences between the nozzle and pipe are evaluated. Comparison is made between a pipe model and a nozzle-pipe interface model, and a LBB PED (Piping Evaluation Diagram) curve is developed to summarize the results for use by piping designers.

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

    International Nuclear Information System (INIS)

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

    2016-01-01

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

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

    Science.gov (United States)

    Cai, Guobiao; Li, Chengen; Tian, Hui

    2016-11-01

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

  17. SSTO rockets. A practical possibility

    Science.gov (United States)

    Bekey, Ivan

    1994-07-01

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

  18. Fuel injector nozzle for an internal combustion engine

    Science.gov (United States)

    Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

    2008-11-04

    A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

  19. The fabrication of nozzles for nuclear components by welding

    International Nuclear Information System (INIS)

    Moraes, M.M.; Krausser, P.; Echeverria, J.A.V.

    1986-01-01

    A nozzle with medium outside diameter of 1000 mm and medium thickness of 150 mm composed integrally by deposited metal by submerged-arc (wire S3NiMo1, 0.5mm) was fabricated in NUCLEP. The nondestructive, mechanical, metallographic and chemical testing carried out in a test sample made by the same procedure and welding parameters, showed results according to specifications established for primary components for nuclear power plants, and the tests presented mechanical properties and tenacity better than similar nozzle samples. This nozzle is cheapest concerning to importations, in respecting to its forged similar. (M.C.K.) [pt

  20. Water Impact Prediction Tool for Recoverable Rockets

    Science.gov (United States)

    Rooker, William; Glaese, John; Clayton, Joe

    2011-01-01

    Reusing components from a rocket launch can be cost saving. NASA's space shuttle system has reusable components that return to the Earth and impact the ocean. A primary example is the Space Shuttle Solid Rocket Booster (SRB) that descends on parachutes to the Earth after separation and impacts the ocean. Water impact generates significant structural loads that can damage the booster, so it is important to study this event in detail in the design of the recovery system. Some recent examples of damage due to water impact include the Ares I-X First Stage deformation as seen in Figure 1 and the loss of the SpaceX Falcon 9 First Stage.To ensure that a component can be recovered or that the design of the recovery system is adequate, an adequate set of structural loads is necessary for use in failure assessments. However, this task is difficult since there are many conditions that affect how a component impacts the water and the resulting structural loading that a component sees. These conditions include the angle of impact with respect to the water, the horizontal and vertical velocities, the rotation rate, the wave height and speed, and many others. There have been attempts to simulate water impact. One approach is to analyze water impact using explicit finite element techniques such as those employed by the LS-Dyna tool [1]. Though very detailed, this approach is time consuming and would not be suitable for running Monte Carlo or optimization analyses. The purpose of this paper is to describe a multi-body simulation tool that runs quickly and that captures the environments a component might see. The simulation incorporates the air and water interaction with the component, the component dynamics (i.e. modes and mode shapes), any applicable parachutes and lines, the interaction of winds and gusts, and the wave height and speed. It is capable of quickly conducting Monte Carlo studies to better capture the environments and genetic algorithm optimizations to reproduce a

  1. Validation of a Solid Rocket Motor Internal Environment Model

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

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

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

    Science.gov (United States)

    1988-04-01

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

  3. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    Science.gov (United States)

    2007-08-30

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

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

    National Research Council Canada - National Science Library

    Buckmaster, John

    2007-01-01

    ..., the combustion of heterogeneous propellants containing aluminum, the use of a genetic algorithm to optimally define false-kinetics parameters in propellant combustion modeling, the calculation of fluctuations...

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

    Science.gov (United States)

    2007-10-01

    and ellipsoids, and the packing of pellets relevant to igniter modeling. Other topics are the instabilities of smolder waves, premixed flame...instabilities in narrow tubes, and flames supported by a spinning porous plug burner . Much of this work has been reported in the high-quality archival...perchlorate in fuel binder, the combustion of model propellant packs of ellipses and ellipsoids, and the packing of pellets relevant to igniter modeling

  6. Maneuver of Spinning Rocket in Flight

    OpenAIRE

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

    1980-01-01

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

  7. Ionospheric shock waves triggered by rockets

    Directory of Open Access Journals (Sweden)

    C. H. Lin

    2014-09-01

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

  8. Aerodynamics and flow characterisation of multistage rockets

    Science.gov (United States)

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

    2017-05-01

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

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

    Directory of Open Access Journals (Sweden)

    V. N. Zubov

    2015-01-01

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

  10. Design and Analyses of High Aspect Ratio Nozzles for Distributed Propulsion Acoustic Measurements

    Science.gov (United States)

    Dippold, Vance F., III

    2016-01-01

    A series of three convergent round-to-rectangular high-aspect ratio nozzles were designed for acoustics measurements. The nozzles have exit area aspect ratios of 8:1, 12:1, and 16:1. With septa inserts, these nozzles will mimic an array of distributed propulsion system nozzles, as found on hybrid wing-body aircraft concepts. Analyses were performed for the three nozzle designs and showed that the flow through the nozzles was free of separated flow and shocks. The exit flow was mostly uniform with the exception of a pair of vortices at each span-wise end of the nozzle.

  11. A Flight Demonstration of Plasma Rocket Propulsion

    Science.gov (United States)

    Petro, Andrew

    1999-01-01

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

  12. Subsonic Glideback Rocket Demonstrator Flight Testing

    Science.gov (United States)

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

    2001-01-01

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

  13. Performances Study of a Hybrid Rocket Engine

    Directory of Open Access Journals (Sweden)

    Adrian-Nicolae BUTURACHE

    2018-06-01

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

  14. Altitude Compensating Nozzle Transonic Performance Flight Demonstration, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Altitude compensating nozzles continue to be of interest for use on future launch vehicle boosters and upper stages because of their higher mission average Isp and...

  15. Design methods of Coanda effect nozzle with two streams

    Directory of Open Access Journals (Sweden)

    Michele TRANCOSSI

    2014-03-01

    Full Text Available This paper continues recent research of the authors about the ACHEON Coanda effect two streams nozzle. This nozzle aims to produce an effective deflection of a propulsive jet with a correspondent deviation of the thrust vector in a 2D plane. On the basis of a previously published mathematical model, based on integral equations, it tries to produce an effective design guideline, which can be adopted for design activities of the nozzle for aeronautic propulsion. The presented model allows defining a governing method for this innovative two stream synthetic jet nozzle. The uncertainness level of the model are discussed and novel aircraft architectures based on it are presented. A CFD validation campaign is produced focusing on validating the model and the designs produced.

  16. Characteristics of Multiplexed Grooved Nozzles for High Flow Rate Electrospray

    International Nuclear Information System (INIS)

    Kim, Kyoung Tae; Kim, Woo Jin; Kim, Sang Soo

    2007-01-01

    The electrospray operated in the cone-jet mode can generate highly charged micro droplets in an almost uniform size at flow rates. Therefore, the multiplexing system which can retain the characteristics of the cone-jet mode is inevitable for the electrospray application. This experiment reports the multiplexed grooved nozzle system with the extractor. The effects of the grooves and the extractor on the performance of the electrospray were evaluated through experiments. Using the grooved nozzle, the stable cone-jet mode can be achieved at the each groove in the grooved mode. Furthermore, the number of nozzles per unit area is increased by the extractor. The multiplexing density is 12 jets per cm 2 at 30 mm distance from the nozzle tip to the ground plate. The multiplexing system for the high flow rate electrospray is realized with the extractor which can diminish the space charge effect without sacrificing characteristics of the cone-jet mode

  17. Characterization of Plasmadynamics within a Small Magnetic Nozzle

    Data.gov (United States)

    National Aeronautics and Space Administration — This proposal presents an experimental and theoretical research project intended to develop a more refined model of the underlying physics of magnetic nozzles. The...

  18. Separation of a light additive gas by separation nozzle cascades

    International Nuclear Information System (INIS)

    Becker, E.; Bley, P.; Ehrfeld, W.; Fritz, W.; Steinhaus, H.

    1984-01-01

    Double-turn separation nozzles, in comparison with single-turn separation nozzles, offer much greater advantages in the separation of UF6 and H2 than in the separation of the U isotopes, for which the double-turn separation nozzles were conceived. By using a double-turn separation-nozzle stage as a preseparation stage in combination with a low-temperature separator, one can reduce the ratio of the buffer input stream to the product stream, in contrast with the solution used up to this time, with only a slight increase in cost of about an order of magnitude. The control program in the case of return feeding of the UF6 from the buffer and the danger of production losses connected with it are thereby correspondingly diminished. An example is given of the enrichment of 235U using the title facility with UF6. (orig./PW)

  19. Optimal Thrust Vectoring for an Annular Aerospike Nozzle, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Recent success of an annular aerospike flight test by NASA Dryden has prompted keen interest in providing thrust vector capability to the annular aerospike nozzle...

  20. Effects of injection nozzle exit width on rotating detonation engine

    Science.gov (United States)

    Sun, Jian; Zhou, Jin; Liu, Shijie; Lin, Zhiyong; Cai, Jianhua

    2017-11-01

    A series of numerical simulations of RDE modeling real injection nozzles with different exit widths are performed in this paper. The effects of nozzle exit width on chamber inlet state, plenum flowfield and detonation propagation are analyzed. The results are compared with that using an ideal injection model. Although the ideal injection model is a good approximation method to model RDE inlet, the two-dimensional effects of real nozzles are ignored in the ideal injection model so that some complicated phenomena such as the reflected waves caused by the nozzle walls and the reversed flow into the nozzles can not be modeled accurately. Additionally, the ideal injection model overpredicts the block ratio. In all the cases that stabilize at one-wave mode, the block ratio increases as the nozzle exit width gets smaller. The dual-wave mode case also has a relatively high block ratio. A pressure oscillation in the plenum with the same main frequency with the rotating detonation wave is observed. A parameter σ is applied to describe the non-uniformity in the plenum. σ increases as the nozzle exit width gets larger. Under some condition, the heat release on the interface of fresh premixed gas layer and detonation products can be strong enough to induce a new detonation wave. A spontaneous mode-transition process is observed for the smallest exit width case. Due to the detonation products existing in the premixed gas layer before the detonation wave, the detonation wave will propagate through reactants and products alternately, and therefore its strength will vary with time, especially near the chamber inlet. This tendency gets weaker as the injection nozzle exit width increases.