Sample records for solid rocket nozzle

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


    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...Paul A.;, "Gelcasting of Alumina," J. Am. Ceram . Soc. 74[3], pp. 612-618, 1991. [18] Blomquist , B. A.; Fink, J. K.; Leibowitz, L.;, "The

  2. Thermal Barriers Developed for Solid Rocket Motor Nozzle Joints

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.


    Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn's braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block.

  3. Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.


    Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

  4. Development of high temperature materials for solid propellant rocket nozzle applications

    Manning, C. R., Jr.; Lineback, L. D.


    Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

  5. New discrimination method for ablative control mechanism in solid-propellant rocket nozzle


    A reasonable discrimination method for ablative control mechanism in solid-propellant rocket nozzle can improve the calculation accuracy of ablation rate. Based on the different rate constants for reactions of C with H2O and CO2,a new discrimination method for ablative control mechanism,which comprehensively considers the influence of nozzle surface temperature and gas component concentration,is presented. Using this new discrimination method,calculations were performed to simulate the nozzle throat insert ablation. The numerical results showed that the calculated ablation rate,which was more close to the measured values,was less than the value calculated by diffusion control mechanisms or by double control mechanisms. And H2O was proved to be the most detrimental oxidizing species in nozzle ablation.

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

    Lin Sun


    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.

  7. Discrimination for ablative control mechanism in solid-propellant rocket nozzle


    The ablation in solid-propellant rocket nozzle is a coupling process resulted by chemistry, heat and mass transfer. Based on the heat and mass transfer theory, the aero-thermo-dynamic, and thermo-chemical kinetics, the thermal-chemical ablation model is established. Simulations are completed on the heat flow field and chemical ablation in the nozzle with different concentrations, frequency factors and activation energy of H2. The calculation results show that the concentration and the activation energy of H2 can provoke the transformation of control mechanism, whereas the influence brought by the frequency factor of H2 is feeble under a high-temperature and high-pressure combustion circumstance. The discrimination for ablative control mechanism is dependent on both concentration and activation energy of H2. This study will be useful in handling ablation and thermal protection problem in the design of solid-propellant rocket.

  8. Test and Analysis of Solid Rocket Motor Nozzle Ablative Materials

    Clayton, J. Louie


    Asbestos free solid motor internal insulation samples were tested at the MSFC Hyperthermal Facility. Objectives of the test were to gather data for analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Instrumentation included use of total calorimeters, thermocouples, and a surface pyrometer for surface temperature measurement. Post-test sample forensics involved measurement of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero-thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

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

    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

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

    Clayton, J. Louie


    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.

  11. Arc Jet Test and Analysis of Asbestos Free Solid Rocket Motor Nozzle Dome Ablative Materials

    Clayton, J. Louie


    Asbestos free solid motor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

  12. Integrated Composite Rocket Nozzle Extension Project

    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...

  13. Reducing Thrusts In Solid-Fuel Rockets

    Bement, Laurence J.


    Thrust-terminating system conceived to reduce thrust of solid-propellant rocket motor in controlled manner such that thrust loads not increased or decreased beyond predictable levels. Concept involves explosively cutting opposing venting pairs in case of rocket motor above nozzles to initiate venting of chamber and reduction of thrust. Vents sized and numbered to control amount and rate of reduction in thrust.

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

    Nath, Nayani Kishore


    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.

  15. Nuclear thermal rocket nozzle testing and evaluation program

    Davidian, Kenneth O.; Kacynski, Kenneth J.


    Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. The Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis 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 + or - 1.17 pct.

  16. Flow separation in rocket nozzles under high altitude condition

    Stark, R.; Génin, C.


    The knowledge of flow separation in rocket nozzles is crucial for rocket engine design and optimum performance. Typically, flow separation is studied under sea-level conditions. However, this disregards the change of the ambient density during ascent of a launcher. The ambient flow properties are an important factor concerning the design of altitude-adaptive rocket nozzles like the dual bell nozzle. For this reason an experimental study was carried out to study the influence of the ambient density on flow separation within conventional nozzles.

  17. Parametric study of solar thermal rocket nozzle performance

    Pearson, J. Boise; Landrum, D. Brian; Hawk, Clark W.


    This paper details a numerical investigation of performance losses in low-thrust solar thermal rocket nozzles. The effects of nozzle geometry on three types of losses were studied; finite rate dissociation-recombination kinetic losses, two dimensional axisymmetric divergence losses, and compressible viscous boundary layer losses. Short nozzle lengths and supersonic flow produce short residence times in the nozzle and a nearly frozen flow, resulting in large kinetic losses. Variations in geometry have a minimal effect on kinetic losses. Divergence losses are relatively small, and careful shaping of the nozzle can nearly eliminate them. The boundary layer in these small nozzles can grow to a major fraction of nozzle radius, and cause large losses. These losses are attributed to viscous drag on the nozzle walls and flow blockage by the boundary layer, especially in the throat region. Careful shaping of the nozzle can produce a significant reduction in viscous losses.

  18. Solid propellant rocket motor

    Dowler, W. L.; Shafer, J. I.; Behm, J. W.; Strand, L. D. (Inventor)


    The characteristics of a solid propellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solid propellant igniter and a progressive burning main solid propellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

  19. Rocket nozzle thermal shock tests in an arc heater facility

    Painter, James H.; Williamson, Ronald A.


    A rocket motor nozzle thermal structural test technique that utilizes arc heated nitrogen to simulate a motor burn was developed. The technique was used to test four heavily instrumented full-scale Star 48 rocket motor 2D carbon/carbon segments at conditions simulating the predicted thermal-structural environment. All four nozzles survived the tests without catastrophic or other structural failures. The test technique demonstrated promise as a low cost, controllable alternative to rocket motor firing. The technique includes the capability of rapid termination in the event of failure, allowing post-test analysis.

  20. Research on Integrative Bonding Process of Solid Rocket Motor Nozzle%固体火箭发动机喷管一体化粘接工艺研究

    王纪霞; 包乐; 胡大宁; 张崇耿; 张新航


    Influence factors on nozzle bonding property including temperature, adhesive, bonding process and gap between metal shell and insert were analyzed to satisfy integrative bonding process of some solid rocket nozzle. The result of the test indicated that the best bonding process was found through controlling the factors, and the reliability of solid rocket was protected.%为了满足固体火箭发动机喷管一体化的粘接要求,提高产品的粘接质量,分析了胶粘剂性能、粘接工艺、温度及内衬与壳体的配合间隙对喷管粘接质量的影响,结果表明,通过控制影响一体化喷管粘接的各种因素,得出了最佳的工艺生产条件,保证了发动机工作的可靠性。

  1. Analytical study of nozzle performance for nuclear thermal rockets

    Davidian, Kenneth O.; Kacynski, Kenneth J.


    Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively.

  2. Analytical study of nozzle performance for nuclear thermal rockets

    Davidian, Kenneth O.; Kacynski, Kenneth J.


    A parametric study has been conducted by the NASA-Lewis Rocket Engine Design Expert System for the convergent-divergent nozzle of the Nuclear Thermal Rocket system, which uses a nuclear reactor to heat hydrogen to high temperature and then expands it through the nozzle. It is established by the study that finite-rate chemical reactions lower performance levels from theoretical levels. Major parametric roles are played by chamber temperature and chamber pressure. A maximum performance of 930 sec is projected at 2700 K, and of 1030 at 3100 K.

  3. Flow Separation and Numerical Simulation of Solid Rocket Motor Nozzle%固体火箭发动机喷管分离流动及其数值模拟

    王晓辉; 于存贵


    Separation flow could be generated in a high-altitude nozzle, in order to study the flow separation on the nozzle performance. A three-dimensional numerical simulation about the separation flow in the high-altitude nozzle of a solid rocket motor (SRM) was conducted. The flow results with different inlet total pressure were obtained by utilizing CFX software. It shows that, due to the separation flow, there are some adverse effects for the stability and thermal protection of the nozzle. The simulation provides a reference of the design of the high-altitude nozzle, also a basis for further study.%大面积比喷管在火箭发动机工作过程中可能产生流动分离的问题,为研究喷管流动分离对喷管性能的影响,利用计算流体力学软件CFX对某固体火箭发动机大面积比喷管内燃气分离流动进行数值模拟.计算出喷管在几种不同入口总压情况下的流场参数分布,显示分离流动会改变燃气内流场流动参数分布,进而会对喷管推力稳定性和热防护性产生不利影响.该研究能为进一步研究大面积比喷管设计提供参考.

  4. Integrated Ceramic Matrix Composite and Carbon/Carbon Structures for Large Rocket Engine Nozzles and Nozzle Extensions Project

    National Aeronautics and Space Administration — Low-cost access to space demands durable, cost-effective, efficient, and low-weight propulsion systems. Key components include rocket engine nozzles and nozzle...

  5. NASA's Advanced solid rocket motor

    Mitchell, Royce E.

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  6. NASA's Advanced solid rocket motor

    Mitchell, Royce E.


    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  7. 双脉冲固体发动机喷管传热烧蚀特性%Characterization of nozzle thermal and ablation response in dual-pulse solid rocket motors

    张晓光; 刘宇; 王长辉


    In order to investigate the nozzle thermal and ablation characteristics in dual- pulse solid rocket motors, the transient value of the throat diameter was obtained from the pressure and thrust measurements. Furthermore, the in-depth thermal response, pyroly- sis/char profiles and surface recession of the nozzle assembly were predicted through fully coupled fluid-solid analysis using the commercial code FLUENT. Results show that during pulse operation, the insulation material pyrolysis/char profiles expand and the nozzle insert erosion rate increases. During pulse separation, heat conduction in the material leads to the decrease in the material temperature difference. The heat transfer and ablation processes of pulse 1 and pulse separation make the nozzle insert exhibit small heat sink, high surface temperature and large surface roughness, which would result in higher throat erosion rate when pulse 2 operates.%为了研究双脉冲固体发动机喷管的传热烧蚀特性,由燃烧室压强及发动机推力试验曲线得到了喷管喉径的瞬变值,由FLUENT流体计算软件进行流固耦合传热烧蚀计算,得到了喷管瞬态温度分布、绝热材料热解炭化情况及碳/碳(C/C)喉衬瞬态烧蚀率,分析了脉冲工作过程及脉冲间隔时间对喷管传热烧蚀的影响.计算结果表明,脉冲工作过程中,绝热材料热解线、炭化线向材料内部扩展,喉衬烧蚀率不断增大;脉冲间隔时间内,喷管材料内部的导热使各处温差减小,温度趋于一致;第一脉冲的传热烧蚀与脉冲间隔的材料导热使第二脉冲工作时喉衬整体热沉小、内壁初始温度高、表面粗糙度大,从而导致较高烧蚀率.

  8. DURACON - Variable Emissivity Broadband Coatings for Liquid Propellant Rocket Nozzles Project

    National Aeronautics and Space Administration — The need exists for a fast drying, robust, low gloss, black, high emissivity coating that can be applied easily on aircraft rocket nozzles and nozzle extensions....

  9. The development of space solid rocket motors in China

    Jianding, Huang; Dingyou, Ye


    China has undertaken to research and develop composite solid propellant rocket motors since 1958. At the request of the development of space technology, composite solid propellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

  10. Carbon-Based Nozzle Thermochemical Erosion Characteristics in Solid Rocket Motors%固体火箭发动机碳基材料喷管热化学烧蚀特性

    张晓光; 王长辉; 刘宇; 任军学


    Based on the thermochemical erosion theory, a two-dimensional axisymmetric, coupled gas-solid-thermal numerical framework was established to predict the carbon-based nozzle erosion in solid rocket motors. Numerical simulations were carried out using the Wall Surface Reaction model of the commercial code FLUENT and the assumption whether the erosion process was chemical kinetics or diffusion controlled was not needed. The method was introduced to simulate the 70-1b BATES motor nozzle erosion and examine the effects of propellant composition, oxidizing species and chamber pressure. The calculated results agree well with experimental data. The erosion rate follows the trend exhibited by the heat flux distribution, and peaks slightly upstream of the throat. The erosion rate decreases with increasing aluminum content and increases almost linearly with chamber pressure. H20 is the dominant oxidizing species in dictating nozzle erosion.%为了准确预示固体火箭发动机碳基材料喷管的烧蚀率,依据热化学烧蚀理论,建立了喷管传热烧蚀的二维轴对称气-固-热耦合计算模型,计算通过FLUENT壁面化学反应模型完成,无需事先假设烧蚀控制机制。针对70-lb BATES发动机喷管进行了烧蚀计算,研究了推进剂配方、氧化性组分、燃烧室压强对喷管烧蚀的影响。结果表明:烧蚀率计算值与试验测试值吻合较好;烧蚀率分布遵循喷管内壁热流密度分布规律,在喉部上游入口处达到峰值;烧蚀率随推进剂Al含量增加而降低,随燃烧室压强升高而近似正比例增大;H2O是决定烧蚀的主要氧化性组分。

  11. Overview of the manufacturing sequence of the Advanced Solid Rocket Motor

    Chapman, John S.; Nix, Michael B.


    The manufacturing sequence of NASA's new Advanced Solid Rocket Motor, developed as a replacement of the Space Shuttle's existing Redesigned Solid Rocket Motor, is overviewed. Special attention is given to the case preparation, the propellant mix/cast, the nondestructuve evaluation, the motor finishing, and the refurbishment. The fabrication sequences of the case, the nozzle, and the igniter are described.

  12. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  13. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

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

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


    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.

  15. Nanoparticles for solid rocket propulsion

    Galfetti, L [Politecnico di Milano, SPLab, Milan (Italy); De Luca, L T [Politecnico di Milano, SPLab, Milan (Italy); Severini, F [Politecnico di Milano, SPLab, Milan (Italy); Meda, L [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marra, G [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marchetti, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Regi, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Bellucci, S [INFN, Laboratori Nazionali di Frascati, Frascati (Italy)


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  16. Nanoparticles for solid rocket propulsion

    Galfetti, L.; DeLuca, L. T.; Severini, F.; Meda, L.; Marra, G.; Marchetti, M.; Regi, M.; Bellucci, S.


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  17. Effects of gas temperature on nozzle damping experiments on cold-flow rocket motors

    Sun, Bing-bing; Li, Shi-peng; Su, Wan-xing; Li, Jun-wei; Wang, Ning-fei


    In order to explore the impact of gas temperature on the nozzle damping characteristics of solid rocket motor, numerical simulations were carried out by an experimental motor in Naval Ordnance Test Station of China Lake in California. Using the pulse decay method, different cases were numerically studied via Fluent along with UDF (User Defined Functions). Firstly, mesh sensitivity analysis and monitor position-independent analysis were carried out for the computer code validation. Then, the numerical method was further validated by comparing the calculated results and experimental data. Finally, the effects of gas temperature on the nozzle damping characteristics were studied in this paper. The results indicated that the gas temperature had cooperative effects on the nozzle damping and there had great differences between cold flow and hot fire test. By discussion and analysis, it was found that the changing of mainstream velocity and the natural acoustic frequency resulted from gas temperature were the key factors that affected the nozzle damping, while the alteration of the mean pressure had little effect. Thus, the high pressure condition could be replaced by low pressure to reduce the difficulty of the test. Finally, the relation of the coefficients "alpha" between the cold flow and hot fire was got.

  18. Solid Rocket Booster-Illustration


    This illustration is a cutaway of the solid rocket booster (SRB) sections with callouts. The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment. The boosters are designed to survive water impact at almost 60 miles per hour, maintain flotation with minimal damage, and preclude corrosion of the hardware exposed to the harsh seawater environment. Under the project management of the Marshall Space Flight Center, the SRB's are assembled and refurbished by the United Space Boosters. The SRM's are provided by the Morton Thiokol Corporation.

  19. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    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.

  20. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    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.

  1. Slag Prediction in Submerged Rocket Nozzle Through Two-Phase CFD Simulations

    Amit Kumar Chaturvedi


    Full Text Available A computational procedure has been established to predict the slag in a practical solid rocket motor with submerged nozzle. Both single-phase and two-phase flow analyses have been performed in the rocket motor port. Three-dimensional Navier-Stokes equations along with SST turbulence model have been solved for gas-phase calculations. The effect of ejected alumina particles from the propellant geometry on the flow field has been simulated through Lagrangian tracking method. The computational methodology is firstly validated by comparing against other numerical results of rocket motors available in the literature before applying the same to predict the slag accumulation of a submerged rocket motor for strategic applications. Burn-back geometries at different instants have been simulated and parametric studies were performed to find out the effect of Al2O3 particle size. It was observed that the slag capture rate increases uniformly with A12O3 particle size. The predicted slag accumulation data match closely with the ground test data for the range of conditions simulated in the present work.Defence Science Journal, Vol. 65, No. 2, March 2015, pp.99-106, DOI: Normal 0 false false false EN-US X-NONE X-NONE

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

    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)


    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.

  3. Effect of injector configuration in rocket nozzle film cooling

    Kumar, A. Lakshya; Pisharady, J. C.; Shine, S. R.


    Experimental and numerical investigations are carried out to analyze the effect of coolant injector configuration on overall film cooling performance in a divergent section of a rocket nozzle. Two different injector orientations are investigated: (1) shaped slots with a divergence angle of 15° (semi-divergent injector) (2) fully divergent slot (fully divergent injector). A 2-dimensional, axis-symmetric, multispecies computational model using finite volume formulation has been developed and validated against the experimental data. The experiments provided a consistent set of measurements for cooling effectiveness for different blowing ratios ranging from 3.7 to 6. Results show that the semi divergent configuration leads to higher effectiveness compared to fully divergent slot at all blowing ratios. The spatially averaged effectiveness results show that the difference between the two configurations is significant at higher blowing ratios. The increase in effectiveness was around 2 % at BR = 3.7 whereas it was around 12 % in the case of BR = 6. Numerical results show the presence of secondary flow recirculation zones near the jet exit for both the injectors. An additional recirculation zone present in the case of fully divergent injector caused an increase in mixing of the coolant and mainstream, and a reduction in film cooling performance.

  4. Lightweight Nozzle Extension for Liquid Rocket Engines Project

    National Aeronautics and Space Administration — The ARES J-2X requires a large nozzle extension. Currently, a metallic nozzle extension is being considered with carbon-carbon composite as a backup. In Phase 1,...

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

    Bondan Sofyan


    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.

  6. Low Cost Carbon-Carbon Rocket Nozzle Development Project

    National Aeronautics and Space Administration — This development will provide an inexpensive vacuum nozzle manufacturing option for NOFBXTM monopropellant systems that are currently being developed under NASA SBIR...

  7. Rapid Fabrication Techniques for Liquid Rocket Channel Wall Nozzles

    Gradl, Paul R.


    The functions of a regeneratively-cooled nozzle are to (1) expand combustion gases to increase exhaust gas velocity while, (2) maintaining adequate wall temperatures to prevent structural failure, and (3) transfer heat from the hot gases to the coolant fluid to promote injector performance and stability. Regeneratively-cooled nozzles are grouped into two categories: tube-wall nozzles and channel wall nozzles. A channel wall nozzle is designed with an internal liner containing a series of integral coolant channels that are closed out with an external jacket. Manifolds are attached at each end of the nozzle to distribute coolant to and away from the channels. A variety of manufacturing techniques have been explored for channel wall nozzles, including state of the art laser-welded closeouts and pressure-assisted braze closeouts. This paper discusses techniques that NASA MSFC is evaluating for rapid fabrication of channel wall nozzles that address liner fabrication, slotting techniques and liner closeout techniques. Techniques being evaluated for liner fabrication include large-scale additive manufacturing of freeform-deposition structures to create the liner blanks. Abrasive water jet milling is being evaluated for cutting the complex coolant channel geometries. Techniques being considered for rapid closeout of the slotted liners include freeform deposition, explosive bonding and Cold Spray. Each of these techniques, development work and results are discussed in further detail in this paper.

  8. Environmentally compatible solid rocket propellants

    Jacox, James L.; Bradford, Daniel J.


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

  9. Facility for cold flow testing of solid rocket motor models

    Bacchus, D. L.; Hill, O. E.; Whitesides, R. Harold


    A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solid propellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The

  10. Cathodic Protection Deployment on Space Shuttle Solid Rocket Boosters

    Zook, Lee M.


    Corrosion protection of the space shuttle solid rocket boosters incorporates the use of cathodic protection(anodes) in concert with several coatings systems. The SRB design has large carbon/carbon composites(motor nozzle) electrically connected to an aluminum alloy structure. Early in the STS program, the aluminum structures incurred tremendous corrosive attack due primarily to the galvanic couple to the carbon/carbon nozzle at coating damage locations. Also contributing to the galvanic corrosion problem were stainless steel and titanium alloy components housed within the aluminum structures and electrically connected to the aluminum structures. This paper will highlight the evolution in the protection of the aluminum structures, providing historical information and summary data from the operation of the corrosion protection systems. Also, data and information will be included regarding the evaluation and deployment of inorganic zinc rich primers as anode area on the aluminum structures.

  11. Design and Experimental Study on Spinning Solid Rocket Motor

    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.

  12. Advanced Solid Rocket Launcher and Its Evolution

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

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

  13. Heat transfer in rocket engine combustion chambers and nozzles

    Anderson, P. G.; Cheng, G. C.; Farmer, R. C.


    Complexities of liquid rocket engine heat transfer which involve the injector faceplate and regeneratively and film cooled walls are being investigated by computational analysis. A conjugate heat transfer analysis will be used to describe localized heating phenomena associated with particular injector configurations and coolant channels and film coolant dumps. These components are being analyzed, and the analyses verified with appropriate test data. Finally, the component analyses will be synthesized into an overall flowfield/heat transfer model. The FDNS code is being used to make the component analyses. Particular attention is being given to the representation of the thermodynamic properties of the fluid streams and to the method of combining the detailed models to represent overall heating. Unit flow models of specific coaxial injector elements have been developed and will be described. Since test data from the NLS development program are not available, new validation heat transfer data have been sought. Suitable data were obtained from a Rocketdyne test program on a model hydrocarbon/oxygen engine. Simulations of these test data will be presented. Recent interest in the hybrid motor have established the need for analyses of ablating solid fuels in the combustion chamber. Analysis of a simplified hybrid motor will also be presented.

  14. Wind Tunnel Tests on Aerodynamic Characteristics of Advanced Solid Rocket

    Kitamura, Keiichi; Fujimoto, Keiichiro; Nonaka, Satoshi; Irikado, Tomoko; Fukuzoe, Moriyasu; Shima, Eiji

    The Advanced Solid Rocket is being developed by JAXA (Japan Aerospace Exploration Agency). Since its configuration has been changed very recently, its aerodynamic characteristics are of great interest of the JAXA Advanced Solid Rocket Team. In this study, we carried out wind tunnel tests on the aerodynamic characteristics of the present configuration for Mach 1.5. Six test cases were conducted with different body configurations, attack angles, and roll angles. A six component balance, oilflow visualization, Schlieren images were used throughout the experiments. It was found that, at zero angle-of-attack, the flow around the body were perturbed and its drag (axial force) characteristics were significantly influenced by protruding body components such as flanges, cable ducts, and attitude control units of SMSJ (Solid Motor Side Jet), while the nozzle had a minor role. With angle-of-attack of five degree, normal force of CNα = 3.50±0.03 was measured along with complex flow features observed in the full-component model; whereas no crossflow separations were induced around the no-protuberance model with CNα = 2.58±0.10. These values were almost constant with respect to the angle-of-attack in both of the cases. Furthermore, presence of roll angle made the flow more complicated, involving interactions of separation vortices. These data provide us with fundamental and important aerodynamic insights of the Advanced Solid Rocket, and they will be utilized as reference data for the corresponding numerical analysis.

  15. Aeroelastic stability analysis of flexible overexpanded rocket nozzle

    Bekka, N.; Sellam, M.; Chpoun, A.


    The aim of this paper is to present a new aeroelastic stability model taking into account the viscous effects for a supersonic nozzle flow in overexpanded regimes. This model is inspired by the Pekkari model which was developed initially for perfect fluid flow. The new model called the "Modified Pekkari Model" (MPM) considers a more realistic wall pressure profile for the case of a free shock separation inside the supersonic nozzle using the free interaction theory of Chapman. To reach this objective, a code for structure computation coupled with aerodynamic excitation effects is developed that allows the analysis of aeroelastic stability for the overexpanded nozzles. The main results are presented in a comparative manner using existing models (Pekkari model and its extended version) and the modified Pekkari model developed in this work.

  16. Introduction to rocket science and engineering

    Taylor, Travis S


    What Are Rockets? The History of RocketsRockets of the Modern EraRocket Anatomy and NomenclatureWhy Are Rockets Needed? Missions and PayloadsTrajectoriesOrbitsOrbit Changes and ManeuversBallistic Missile TrajectoriesHow Do Rockets Work? ThrustSpecific ImpulseWeight Flow RateTsiolkovsky's Rocket EquationStagingRocket Dynamics, Guidance, and ControlHow Do Rocket Engines Work? The Basic Rocket EngineThermodynamic Expansion and the Rocket NozzleExit VelocityRocket Engine Area Ratio and LengthsRocket Engine Design ExampleAre All Rockets the Same? Solid Rocket EnginesLiquid Propellant Rocket Engines

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

    Santoro, Robert J.; Pal, Sibtosh


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

  18. Characterisation of Materials used in Flex Bearings of Large Solid Rocket Motors

    CH.V. Ram Mohan


    Full Text Available Solid rocket motors are propulsion devices for both satellite launchers and missiles, which require guidance and steering to fly along a programmed trajectory and to compensate for flight disturbances. A typical solid rocket motor consists of motor case, solid propellant grain, motor insulation, igniter and nozzle. In most solid rocket motors, thrust vector control (TVC is required. One of the most efficient methods of TVC is by flex nozzle system. The flex nozzle consists of a flexible bearing made of an elastomeric material alternating with reinforcement rings of metallic or composite material. The material characterisation of AFNOR 15CDV6 steel and the natural rubber-based elastomer developed for use in flex nozzle are discussed. This includes testing, modelling of the material, selection of a material model suitable for analysis, and the validation of material model.Defence Science Journal, 2011, 61(3, pp.264-269, DOI:

  19. Specific Impulses Losses in Solid Propellant Rockets


    to use the collision function form proposed by Golovin to simplify this production term: 4C><=) <P- .: Accordingly: m hence, by integration: Now, we...November 21, 1940 in Paris, Seine. VFirst Thesis. "Contribution to the Study of Specific i Impulse Loss in Solid Propellant Rockets." Second Thesis


    XU Xiao-qiang; DENG Jian; REN An-lu; LU Chuan-jing


    In this paper, a study of the high-speed gas jet of a rocket nozzle underwater was carried out using commercially available CFD software FLUENT with it's user-defined-function.The volume of fluid technique based on finite volume method was adopted to solve the time-dependent multiphase flow including a compressible phase, and the PISO algorithm was included.The computed results show that this problem was calculated successfully.The gas bubble behind the nozzle, and the wave structure existing in highly compressed gas in water were captured accurately.

  1. 微型固体姿控发动机微喷管内气粒两相流动规律CFD-DSMC研究%Research on the gas-particle two-phase flow in the micro nozzle of attitude control micro solid rocket motor

    夏广庆; 张斌; 孙得川; 陈茂林


    微型固体姿控发动机在航天领域具有广泛的应用前景.以基于MEMS技术的微喷管为研究对象,首先通过计算微喷管中的克努森数,得到了微喷管中的气相流动状态;然后,采用CFD-DSMC方法,模拟了微喷管中的气粒两相流动,并研究了颗粒相质量分数和粒径对气相流动的影响.结果表明,在所研究的来流条件下,微喷管中的连续介质假设是成立的;气相与颗粒相间的动量和能量交换,导致气相马赫数降低、温度升高,同时也导致颗粒相速度增加、温度降低;颗粒相质量分数和粒径均能显著影响气相的马赫数和温度.%Attitude control micro solid rocket motor has wide application potential in the aerospace field. The gas-particle two-phase flow in the micro nozzle based on the MEMS technology was investigated. Firstly, through calculating the Knudsen number of the micro nozzle, the gas phase flow state in the micro nozzle was obtained. Then the gas-particle two-phase flow in the micro nozzle was simulated by using the method of CFD-DSMC. The influence of particle mass fraction and particle diameter on the gas phase flow was studied. The result shows that the continuum assumption in the micro nozzle is established under the conditions of the defined flow in the study. The exchange of momentum and energy between the gas phase and the particle can not noly reduce the gas phase Mach number and raise the temperature, but also increase the particle phase velocity and decrease the temperature. The particle phase mass fraction and particle diameter can significantly influence the Mach number and temperature of gas phase.

  2. Computational and experimental study on supersonic film cooling for liquid rocket nozzle applications

    Vijayakumar Vishnu


    Full Text Available An experimental and computational investigation of supersonic film cooling (SFC was conducted on a subscale model of a rocket engine nozzle. A computational model of a convergent-divergent nozzle was generated, incorporating a secondary injection module for film cooling in the divergent section. Computational Fluid Dynamic (CFD simulations were run on the model and different injection configurations were analyzed. The CFD simulations also analyzed the parameters that influence film cooling effectiveness. Subsequent to the CFD analysis and literature survey an angled injection configuration was found to be more effective, therefore the hardware was fabricated for the same. The fabricated nozzle was later fixed to an Air-Kerosene combustor and numerous sets of experiments were conducted in order to ascertain the effect on film cooling on the nozzle wall. The film coolant employed was gaseous Nitrogen. The results showed substantial cooling along the walls and a considerable reduction in heat transfer from the combustion gas to the wall of the nozzle. Finally the computational model was validated using the experimental results. There was fairly good agreement between the predicted nozzle wall temperature and the value obtained through experiments.

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

    Gradl, Paul; Brandsmeier, Will


    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.

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

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


    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.

  5. Ignition transient analysis of solid rocket motor

    Han, Samuel S.


    Measurement data on the performance of Space Shuttle Solid Rocket Motor show wide variations in the head-end pressure changes and the total thrust build-up during the ignition transient periods. To analyze the flow and thermal behavior in the tested solid rocket motors, a 1-dimensional, ideal gas flow model via the SIMPLE algorithm was developed. Numerical results showed that burning patterns in the star-shaped head-end segment of the propellant and the erosive burning rate are two important factors controlling the ignition transients. The objective of this study is to extend the model to include the effects of aluminum particle commonly used in solid propellants. To treat the effects of aluminum-oxide particles in the combustion gas, conservation of mass, momentum, and energy equations for the particles are added in the numerical formulation and integrated by an inter-phase-slip algorithm.

  6. Atmospheric scavenging of solid rocket exhaust effluents

    Fenton, D. L.; Purcell, R. Y.


    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

  7. The 260: The Largest Solid Rocket Motor Ever Tested

    Crimmins, P.; Cousineau, M.; Rogers, C.; Shell, V.


    Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

  8. End-effects-regime in full scale and lab scale rocket nozzles

    Rojo, Raymundo; Tinney, Charles; Baars, Woutijn; Ruf, Joseph


    Modern rockets utilize a thrust-optimized parabolic-contour design for their nozzles for its high performance and reliability. However, the evolving internal flow structures within these high area ratio rocket nozzles during start up generate a powerful amount of vibro-acoustic loads that act on the launch vehicle. Modern rockets must be designed to accommodate for these heavy loads or else risk a catastrophic failure. This study quantifies a particular moment referred to as the ``end-effects regime,'' or the largest source of vibro-acoustic loading during start-up [Nave & Coffey, AIAA Paper 1973-1284]. Measurements from full scale ignitions are compared with aerodynamically scaled representations in a fully anechoic chamber. Laboratory scale data is then matched with both static and dynamic wall pressure measurements to capture the associating shock structures within the nozzle. The event generated during the ``end-effects regime'' was successfully reproduced in the both the lab-scale models, and was characterized in terms of its mean, variance and skewness, as well as the spectral properties of the signal obtained by way of time-frequency analyses.

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

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


    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.

  10. 固体火箭发动机喉衬流场及热结构耦合分析%Coupled fluid, thermal and structural analysis of nozzle inserts in solid rocket motors

    张晓光; 王长辉; 刘宇; 任军学


    In order to accurately predict the thermal and structural behavior of throat inserts in solid rocket-motor environments, a fluid-thermal-structural coupling model was established based on CFD code FLUENT and FEM code ANSYS. Fully coupled fluid-solid analysis was performed first to simulate the heat transfer and material erosion process, using the method of whole-field discretization and solution in FLUENT. Subsequently, the structural analysis was carried out in ANSYS based on the resulting pressure and thermal loading. In such a procedure, two-way coupling was considered between flow and heat transfer while one-way coupling method was employed in structural analysis. Numerical results show that the convective heat transfer coefficient gradually decreases with increase of wall temperature during the motor firing. The wall temperature and erosion rate follow the trend exhibited by the heat flux distribution, and attain peak at upstream of the throat. The stress is most severe in the throat region due to the steep temperature gradient.%为了准确预示固体火箭发动机喉衬在燃气环境中的烧蚀传热及热结构行为,建立了基于FLUENT流体计算软件和ANSYS结构分析软件的流场及热结构耦合分析模型.由FLUENT采用整场离散、整场求解的方法进行流固耦合烧蚀传热模拟,得到的压强及温度分布导入ANSYS进行结构分析,实现了流场与烧蚀传热的双向耦合以及流场、热到结构的单向耦合.算例结果表明,在发动机工作过程中,喉衬内壁对流换热系数因壁温升高而逐渐降低;内壁温度及烧蚀率遵循内壁热流密度的分布规律,在喉部上游达到峰值;喉部区域对流换热严重,固相材料温度梯度高,是应力集中区.

  11. Application of Optimization Techniques to Design of Unconventional Rocket Nozzle Configurations

    Follett, W.; Ketchum, A.; Darian, A.; Hsu, Y.


    Several current rocket engine concepts such as the bell-annular tri-propellant engine, and the linear aerospike being proposed for the X-33 require unconventional three dimensional rocket nozzles which must conform to rectangular or sector shaped envelopes to meet integration constraints. These types of nozzles exist outside the current experience database, therefore, the application of efficient design methods for these propulsion concepts is critical to the success of launch vehicle programs. The objective of this work is to optimize several different nozzle configurations, including two- and three-dimensional geometries. Methodology includes coupling computational fluid dynamic (CFD) analysis to genetic algorithms and Taguchi methods as well as implementation of a streamline tracing technique. Results of applications are shown for several geometeries including: three dimensional thruster nozzles with round or super elliptic throats and rectangualar exits, two- and three-dimensional thrusters installed within a bell nozzle, and three dimensional thrusters with round throats and sector shaped exits. Due to the novel designs considered for this study, there is little experience which can be used to guide the effort and limit the design space. With a nearly infinite parameter space to explore, simple parametric design studies cannot possibly search the entire design space within the time frame required to impact the design cycle. For this reason, robust and efficient optimization methods are required to explore and exploit the design space to achieve high performance engine designs. Five case studies which examine the application of various techniques in the engineering environment are presented in this paper.

  12. Flow processes in overexpanded chemical rocket nozzles. Part 3: Methods for the aimed flow separation and side load reduction

    Schmucker, R. H.


    Methods aimed at reduction of overexpansion and side load resulting from asymmetric flow separation for rocket nozzles with a high opening ratio are described. The methods employ additional measures for nozzles with a fixed opening ratio. The flow separation can be controlled by several types of nozzle inserts, the properties of which are discussed. Side loads and overexpansion can be reduced by adapting the shape of the nozzle and taking other additional measures for controlled separation of the boundary layer, such as trip wires.

  13. Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

    Shi, John J.


    At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.

  14. C/C-SiC Composites for Nozzle of Solid Propellant Ramjet

    WANG Lingling


    Full Text Available Carbon fiber reinforced carbon and silicon carbide matrix composites for nozzle inner of solid propellant ramjet were prepared by using the hybrid process of "chemical vapor infiltration + precursor impregnation pyrolysis (CVI+PIP". The microstructure, flexural and anti-ablation properties of the C/C-SiC composites and hydraulic test and rocket motor hot firing test for nozzle inner of solid propellant ramjet were comprehensively investigated. The results show that when the flexural strength of the composite reachs 197 MPa, the fracture damage behavior of the composites presents typical toughness mode.Also the composites has excellent anti-ablative property, i.e., linear ablation rate is only 0.0063 mm·s-1 after 200 s ablation. The C/C-SiC component have excellent integral bearing performance with the hydraulic bursting pressure of 6.5 MPa, and the high temperature combination property of the C/C-SiC composite nozzle inner is verified through motor hot firing of solid propellant ramjet.

  15. Analysis of Flame Deflector Spray Nozzles in Rocket Engine Test Stands

    Sachdev, Jai S.; Ahuja, Vineet; Hosangadi, Ashvin; Allgood, Daniel C.


    The development of a unified tightly coupled multi-phase computational framework is described for the analysis and design of cooling spray nozzle configurations on the flame deflector in rocket engine test stands. An Eulerian formulation is used to model the disperse phase and is coupled to the gas-phase equations through momentum and heat transfer as well as phase change. The phase change formulation is modeled according to a modified form of the Hertz-Knudsen equation. Various simple test cases are presented to verify the validity of the numerical framework. The ability of the methodology to accurately predict the temperature load on the flame deflector is demonstrated though application to an actual sub-scale test facility. The CFD simulation was able to reproduce the result of the test-firing, showing that the spray nozzle configuration provided insufficient amount of cooling.

  16. Low NOx nozzle tip for a pulverized solid fuel furnace

    Donais, Richard E; Hellewell, Todd D; Lewis, Robert D; Richards, Galen H; Towle, David P


    A nozzle tip [100] for a pulverized solid fuel pipe nozzle [200] of a pulverized solid fuel-fired furnace includes: a primary air shroud [120] having an inlet [102] and an outlet [104], wherein the inlet [102] receives a fuel flow [230]; and a flow splitter [180] disposed within the primary air shroud [120], wherein the flow splitter disperses particles in the fuel flow [230] to the outlet [104] to provide a fuel flow jet which reduces NOx in the pulverized solid fuel-fired furnace. In alternative embodiments, the flow splitter [180] may be wedge shaped and extend partially or entirely across the outlet [104]. In another alternative embodiment, flow splitter [180] may be moved forward toward the inlet [102] to create a recessed design.

  17. Development of a miniature solid propellant rocket motor for use in plume simulation studies

    Baran, W. J.


    A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

  18. Solid Rocket Launch Vehicle Explosion Environments

    Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.


    Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

  19. Slip-model Performance for Underexpanded Micro-scale Rocket Nozzle Flows

    José A. Morí(n)igo; José Hermida Quesada; Francisco Caballero Requena


    In aerospace Micro-ElectroMechanical Systems (MEMS), the characteristic length scale of the flow approaches the molecular mean free path, thus invalidating the continuum description and enforcing the use of particle methods, like the Direct Simulation Monte Carlo (DSMC), to deal with the non-equilibrium regions. Within the slip-regime (0.01<Kn<~0.1) both approaches, continuum and particle-based, seem to behave well in terms of accuracy. The present study summarizes the implementation and results obtained with a 2nd-order slip boundary condition in a Navier-Stokes solver to address the rarefaction near the nozzle walls. Its assessment and application to a cold-gas micro-scale conical nozzle of 300μm throat diameter, discharging into the low-pressure freestream,constitutes the major aim of the work. The slip-model incorporates the velocity slip with thermal creep and temperature jump, thus permitting to deal with non-isothermal flows as well. Results show that the gas experiences an intense rarefaction in the lip vicinity, pointing to the limits of model validity. Furthermore, a strong Mach deceleration is observed, attributed to the rather thick subsonic boundary layer and supersonic bulk heating caused by the viscous dissipation, in contrast with the expansion to occur in large rocket nozzles during underexpanded operation.

  20. The Chameleon Solid Rocket Propulsion Model

    Robertson, Glen A.


    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).

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

    Keanini, R G; Tkacik, Peter T; Weggel, David C; Knight, P Douglas


    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. Comp...

  2. Study of the Crack Found in Mould Pressing of Tailpipe Nozzle Insert about Some Solid Rocket Motor%发动机尾管内衬模压裂纹的研究

    邹敏; 张飞; 李晓奋; 严伟兴; 李聪; 张新航


    针对发动机尾管内衬经模压成型后经常出现不同程度的裂纹缺陷,通过对模压预混料的质量指标、成型过程中的温度和升温速率等主要因素的分析与控制,发现裂纹的产生是多种因素综合作用的结果。通过对成型工艺过程中各因素的加强控制,可以避免裂纹缺陷的产生。%Aiming at the crack usually found at the surface of the tailpipe nozzle insert of some SRM during moulding pressing, the forming reasons are analyzed and controlled by target of quality, temperature and heating rate. Crack is the result of comprehensive shaping of many factors. Strengthening various factors in the process of molding process control can avoid the generation of crack defects.

  3. Innovative Metallized Formulations for Solid Rocket Propulsion

    Luigi T DeLUCA; Luciano GALFETTI; Filippo MAGGI; Giovanni COLOMBO; Alice REINA; Stefano DOSSI; Daniele CONSONNI; Melissa BRAMBILLA


    Several metallized solid rocket propellants,AP/Metal/HTPB in the ratio 68/18/1 4,were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano.Effects of the metals (micrometric and nanometric Al,B,Mg,and a variety of dual metals) on the performance of the propellant were studied and contrasted to a conventional micrometric aluminum (30 μm average grain size) taken as reference.It is shown that the propellant microstructure plays a fundamental role in controlling the critical aggregation/agglomeration phenomena occurring below and near the burning surface.Two specific effects of microstructure in terms of steady burning rate and average agglomerate size are illustrated.

  4. Erosion and deposition of carbon-carbon nozzle inserts in low-thrust long-duration solid rocket motors%小推力长时间工作固体火箭发动机C/C喉衬的烧蚀与沉积

    张晓光; 王长辉; 刘宇; 熊文波; 任军学


    针对C/C喉衬喷管小推力长时间工作固体火箭发动机,分别开展了含铝、不含铝两种推进剂状态的地面试验。根据燃烧室压强及发动机推力测试曲线计算了喷管喉径的瞬变值,对比研究了喉衬的烧蚀、沉积过程,指出含铝推进剂发动机C/C喉衬先后经历初始沉积、沉积消融、持续烧蚀、烧蚀与沉积交替四个阶段,而推进剂不含铝时则没有明显的初始沉积与沉积消融。讨论了推进剂配方、燃烧室压强、喷管结构等因素对喉衬烧蚀、沉积的影响,并提出了相应的改善措施。%In order to study the erosion and deposition process of the carbon-carbon nozzle inserts in low-thrust,long-duration solid rocket applications,two motors using metallized and non-metallized propellants,respectively,were tested.The transient value of the throat diameter was obtained from the pressure and thrust measurements.It was indicated that for metallized propellant the throat experienced four stages as initial deposition,deposition melt,continuous erosion,alternately erosion and deposition,while no initial deposition and melt phenomena were observed in the non-metallized case.Some approaches to minimize throat variation were developed based upon the investigation of the effect of chamber pressure,propellant composition and nozzle configurations on the throat erosion and deposition.

  5. Nozzle

    Chen, Alexander G.; Cohen, Jeffrey M.


    A fuel injector has a number of groups of nozzles. The groups are generally concentric with an injector axis. Each nozzle defines a gas flowpath having an outlet for discharging a fuel/air mixture jet. There are means for introducing the fuel to the air. One or more groups of the nozzles are oriented to direct the associated jets skew to the injector axis.

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

    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.

  7. Solid Rocket Motor Design Using Hybrid Optimization

    Kevin Albarado


    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.

  8. Heat transfer in rocket engine combustion chambers and regeneratively cooled nozzles


    A conjugate heat transfer computational fluid dynamics (CFD) model to describe regenerative cooling in the main combustion chamber and nozzle and in the injector faceplate region for a launch vehicle class liquid rocket engine was developed. An injector model for sprays which treats the fluid as a variable density, single-phase media was formulated, incorporated into a version of the FDNS code, and used to simulate the injector flow typical of that in the Space Shuttle Main Engine (SSME). Various chamber related heat transfer analyses were made to verify the predictive capability of the conjugate heat transfer analysis provided by the FDNS code. The density based version of the FDNS code with the real fluid property models developed was successful in predicting the streamtube combustion of individual injector elements.

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

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


    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.

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

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


    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.

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

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


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

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

    Xiaodong He


    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.

  13. Magnetic resonance imaging (MRI) of solid rocket components

    Wallner, A.S. [Missouri Western State College, St. Joseph, MO (United States); Nissan, R.A.; Merwin, L.H. [Naval Air Warfare Center, China Lake, CA (United States)] [and others


    The evaluation of solid rocket components has become an area of great interest. Studying these materials with MRI offers a great advantage to observe knit lines, regions of inhomogeneity, voids, defects, plasticizer rich/poor areas and solids distribution because of the nondestructive nature of the technique. Aspects of sample preparation, spectroscopic relaxation studies, and MRI as a method of studying these systems will be discussed. Initial images show the ability to image propellant, liner, and explosive materials with an in-plane resolution of 70 {mu}m/pixel. These initial images show that MRI can be developed as a viable nondestructive evaluation method of solid rocket components.

  14. Application of C/C Composites in Rocket Engine Nozzles in Japan%日本火箭发动机喷管用C/C复合材料

    李崇俊; 崔红; 李瑞珍


    The application state of C/C composites in solid rocket motor nozzle in Japan.The components include a rosetta carbon fabric laminated 2D-C/C exit cone for satellite apogee boost motor,and 3D-C/C throat inserts for solid rocket booster and launch vehicle.The rayon based carbon fiber is adopted to make the 2D-C/C exit cone.The 3D-C/C throat insert,used as the 1st stage of M-V solid rocket launch vehicle,has a dimension of Фll00 mm in outer diameter and a density of 1.95 g/em3.The 3 D preform is orthogonally weaved in a cylinder structure,and then densified by a repeating heat isostatic pressure-graphitization cycles.Applications of C/C composites in both solid and liquid rocket motor nozzle extendable exit cones are future development trend in this area.%介绍了C/C复合材料在日本固体火箭发动机喷管的应用情况,主要包括卫星远地点助推发动机用螺旋形状碳布铺层的2D-C/C扩张段、固体助推器及固体运载用3D-C/C喉衬.2D-C/C扩张段采用黏胶丝基碳纤维成型,M-V固体运载一级发动机C/C喉衬采用碳纤维三向正交圆筒编织结构,热等静压-石墨化致密,外径Ф1100 mm,密度达1.95 g/cm3.C/C复合材料在固体及液体火箭发动机喷管延伸出口锥的应用是未来的发展方向.

  15. Measurement and Characterization of Space Shuttle Solid Rocket Motor Plume Acoustics

    Kenny, Jeremy; Hobbs, Chris; Plotkin, Ken; Pilkey, Debbie


    in Utah. The remaining RSRM static firings will take place on elevated terrain, with the nozzle exit plume being mostly undeflected and the landscape allowing placement of microphones within direct line of sight to the exhaust plume. These measurements will help assess the current extrapolation process by direct comparison between subscale and full scale solid rocket motor data.

  16. Performance of Several Conical Convergent-Divergent Rocket-Type Exhaust Nozzles

    Campbell, C. E.; Farley, J. M.


    An investigation was conducted to obtain nozzle performance data with relatively large-scale models at pressure ratios as high as 120. Conical convergent-divergent nozzles with divergence angles alpha of 15, 25, and 29 deg. were each tested at area ratios of approximately 10, 25, and 40. Heated air (1200 F) was supplied at the nozzle inlet at pressures up to 145 pounds per square inch absolute and was exhausted into quiescent air at pressures as low as 1.2 pounds per square inch absolute. Thrust ratios for all nozzle configurations are presented over the range of pressure ratios attainable and were extrapolated when possible to design pressure ratio and beyond. Design thrust ratios decreased with increasing nozzle divergence angle according to the trend predicted by the (1 + cos alpha)/2 parameter. Decreasing the nozzle divergence angle resulted in sizable increases in thrust ratio for a given surface-area ratio (nozzle weight), particularly at low nozzle pressure ratios. Correlations of the nozzle static pressure at separation and of the average static pressure downstream of separation with various nozzle parameters permitted the calculation of thrust in the separated-flow region from unseparated static-pressure distributions. Thrust ratios calculated by this method agreed with measured values within about 1 percent.

  17. On the history of the development of solid-propellant rockets in the Soviet Union

    Pobedonostsev, Y. A.


    Pre-World War II Soviet solid-propellant rocket technology is reviewed. Research and development regarding solid composite preparations of pyroxyline TNT powder is described, as well as early work on rocket loading calculations, problems of flight stability, and aircraft rocket launching and ground rocket launching capabilities.

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

    Ali, Aliyah N.; Borrer, Jerry L.


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

  19. Stratospheric aluminum oxide. [possibly from solid-fuel rocket exhausts

    Brownlee, D. E.; Tomandl, D.; Ferry, G. V.


    Balloons and U-2 aircraft were used to collect micrometer-sized stratospheric aerosols. It was discovered that for the past 6 years at least, aluminum oxide spheres have been the major stratospheric particulate in the size range from 3 to 8 micrometers. The most probable source of the spheres is the exhaust from solid-fuel rockets.

  20. Burn Rate Modelling of Solid Rocket Propellants (Short Communication

    A.R. Kulkarni


    Full Text Available A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

  1. STS-27 Atlantis, OV-104, solid rocket booster (SRB) inspection


    Engineers, kneeling inside a hollow solid rocket booster (SRB), closely inspect the SRB segments and seams in the Kennedy Space Center (KSC) rotation and processing facility. The SRB will be used on STS-27 Atlantis, Orbiter Vehicle (OV) 104. The booster segments were transported via rail car from Morton Thiokol's Utah manufacturing plant. View provided by KSC with alternate number KSC-88PC-492.

  2. Storable Hypergolic Solid Fuel for Hybrid Rocket Engines

    R. V. Singh


    Full Text Available A solid fuel was synthesised by condensing aniline with furfuraldehyde. The product was directly cast in the rocket motor casing. After curing a hard solid mass was obtained. This was found to have good hypergolicity with RFNA (Red Fuming Nitric Acid, good storability at room temperature and the mechanical properties. The paper presented the techniques of casting, ignition delay measurements and indicates the future programme for this study.

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

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


    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.

  4. Introduction of laser initiation for the 48-inch Advanced Solid Rocket Motor (ASRM) test motors at Marshall Space Flight Center (MSFC)

    Zimmerman, Chris J.; Litzinger, Gerald E.


    The Advanced Solid Rocket Motor is a new design for the Space Shuttle Solid Rocket Booster. The new design will provide more thrust and more payload capability, as well as incorporating many design improvements in all facets of the design and manufacturing process. A 48-inch (diameter) test motor program is part of the ASRM development program. This program has multiple purposes for testing of propellent, insulation, nozzle characteristics, etc. An overview of the evolution of the 48-inch ASRM test motor ignition system which culminated with the implementation of a laser ignition system is presented. The laser system requirements, development, and operation configuration are reviewed in detail.

  5. Particle behavior in solid propellant rockets

    Netzer, D. W.; Diloreto, V. D.; Dubrov, E.


    The use of holography, high speed motion pictures, light scattering measurements, and post-fire particle collection/scanning electron microscopic examination to study the combustion of composite solid propellants is discussed. The relative advantages and disadvantages of the different experimental techniques for obtaining two-phase flow characteristics within the combustion environment of a solid propellant grain are evaluated. Combustion bomb studies using high speed motion pictures and post-fire residue analysis were completed for six low metal content propellants. Resolution capabilities and the relationships between post-fire residue and motion picture data are determined. Initial testing using a holocamera together with a 2D windowed motor is also described.

  6. Performance of high area ratio nozzles for a small rocket thruster

    Kushida, R. O.; Hermel, J.; Apfel, S.; Zydowicz, M.


    Theoretical estimates of supersonic nozzle performance have been compared to experimental test data for nozzles with an area ratio of 100:1 conical and 300:1 optimum contour, and 300:1 nozzles cut off at 200:1 and 100:1. These tests were done on a Hughes Aircraft Company 5 lbf monopropellant hydrazine thruster with chamber pressures ranging from 25 to 135 psia. The analytic method used is the conventional inviscid method of characteristic with correction for laminar boundary layer displacement and drag. Replacing the 100:1 conical nozzle with the 300:1 contoured nozzle resulted in an improvement in thrust performance of 0.74 percent at chamber pressure of 25 psia to 2.14 percent at chamber pressure of 135 psia. The data is significant because it is experimental verification that conventional nozzle design techniques are applicable even where the boundary layer is laminar and displaces as much as 35 percent of the flow at the nozzle exit plane.

  7. Time-Frequency Analysis of Rocket Nozzle Wall Pressures During Start-up Transients

    Baars, Woutijn J.; Tinney, Charles E.; Ruf, Joseph H.


    Surveys of the fluctuating wall pressure were conducted on a sub-scale, thrust- optimized parabolic nozzle in order to develop a physical intuition for its Fourier-azimuthal mode behavior during fixed and transient start-up conditions. These unsteady signatures are driven by shock wave turbulent boundary layer interactions which depend on the nozzle pressure ratio and nozzle geometry. The focus however, is on the degree of similarity between the spectral footprints of these modes obtained from transient start-ups as opposed to a sequence of fixed nozzle pressure ratio conditions. For the latter, statistically converged spectra are computed using conventional Fourier analyses techniques, whereas the former are investigated by way of time-frequency analysis. The findings suggest that at low nozzle pressure ratios -- where the flow resides in a Free Shock Separation state -- strong spectral similarities occur between fixed and transient conditions. Conversely, at higher nozzle pressure ratios -- where the flow resides in Restricted Shock Separation -- stark differences are observed between the fixed and transient conditions and depends greatly on the ramping rate of the transient period. And so, it appears that an understanding of the dynamics during transient start-up conditions cannot be furnished by a way of fixed flow analysis.

  8. Application of Optical Measurement Techniques During Fabrication and Testing of Liquid Rocket Nozzles

    Gradl, Paul R.


    This paper presents a series of optical measurement techniques that were developed for use during large-scale fabrication and testing of nozzle components. A thorough understanding of hardware throughout the fabrication cycle and hotfire testing is critical to meet component design intent. Regeneratively cooled nozzles and associated tooling require tight control of tolerances during the fabrication process to ensure optimal performance. Additionally, changes in geometry during testing can affect performance of the nozzle and mating components. Structured light scanning and digital image correlation techniques were used to collect data during the fabrication and test of nozzles, in addition to other engine components. This data was used to analyze deformations data during machining, heat treatment, assembly and testing operations. A series of feasibility experiments were conducted for these techniques that led to use on full scale nozzles during the J-2X upper stage engine program in addition to other engine development programs. This paper discusses the methods and results of these measurement techniques throughout the nozzle life cycle and application to other components.

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

    Laubacher, Brian A.


    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

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

    Moore, Dennis R.; Phelps, Willie J.


    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

  11. Advanced Computer Science on Internal Ballistics of Solid Rocket Motors

    Shimada, Toru; Kato, Kazushige; Sekino, Nobuhiro; Tsuboi, Nobuyuki; Seike, Yoshio; Fukunaga, Mihoko; Daimon, Yu; Hasegawa, Hiroshi; Asakawa, Hiroya

    In this paper, described is the development of a numerical simulation system, what we call “Advanced Computer Science on SRM Internal Ballistics (ACSSIB)”, for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solid propellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.

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

    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

  13. Advanced Multi-Phase Flow CFD Model Development for Solid Rocket Motor Flowfield Analysis

    Liaw, Paul; Chen, Y. S.; Shang, H. M.; Doran, Denise


    It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better

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

    Martin, James A.


    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.

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

    Martin, James A.


    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.

  16. Performance of a UTC FW-4S solid propellant rocket motor under the command effects of simulated altitude and rotational spin

    Merryman, H. L.; Smith, L. R.


    One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.

  17. Effect of Temperature on Mechanical Properties of Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical properties of solid rocket propellants are dependent on temperature. Any change in temperature brings significant change in the tensile strength, percentage elongation, and elastic modulus of the propellant. Different classes of operational solid rocket propellants namely extruded double-base propellants, composite, extruded composite and nitrarte ester polyester propellants were evaluated at different temperatures in the operating range of the rockets and missiles preferably in the range of –50 oC to +55 oC. It was observed that for each class of propellant, as temperature reduces, propellant becomes hard. This is depicted by increase in elastic modulus and tensile strength of the material. However, trend of percentage elongation is not very uniform. Extruded double-base propellants show less percentage elongation (around 1 per cent at reduced temperature (–50 oC probably due to brittleness. So is the trend with case-bonded composite propellants. However, reverse trend is exhibited by cartridge-loaded composite propellants and nitrate ester polyester propellants. Such propellants show higher percentage elongation (6 per cent for CLCP and 35 per cent for NEPE at reduced temperature (–50 oC. This makes such propellants tough and more area under stress-strain curve at reduced temperature is observed.Defence Science Journal, 2011, 61(6, pp.529-533, DOI:

  18. MEMS-Based Solid Propellant Rocket Array Thruster

    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.

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

    Fischbach, Sean


    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.

  20. Rocket Solid Propellant Alternative Based on Ammonium Dinitramide

    Grigore CICAN


    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.

  1. Propellant development for the Advanced Solid Rocket Motor

    Landers, L. C.; Stanley, C. B.; Ricks, D. W.


    The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

  2. Coupled Solid Rocket Motor Ballistics and Trajectory Modeling for Higher Fidelity Launch Vehicle Design

    Ables, Brett


    Multi-stage launch vehicles with solid rocket motors (SRMs) face design optimization challenges, especially when the mission scope changes frequently. Significant performance benefits can be realized if the solid rocket motors are optimized to the changing requirements. While SRMs represent a fixed performance at launch, rapid design iterations enable flexibility at design time, yielding significant performance gains. The streamlining and integration of SRM design and analysis can be achieved with improved analysis tools. While powerful and versatile, the Solid Performance Program (SPP) is not conducive to rapid design iteration. Performing a design iteration with SPP and a trajectory solver is a labor intensive process. To enable a better workflow, SPP, the Program to Optimize Simulated Trajectories (POST), and the interfaces between them have been improved and automated, and a graphical user interface (GUI) has been developed. The GUI enables real-time visual feedback of grain and nozzle design inputs, enforces parameter dependencies, removes redundancies, and simplifies manipulation of SPP and POST's numerous options. Automating the analysis also simplifies batch analyses and trade studies. Finally, the GUI provides post-processing, visualization, and comparison of results. Wrapping legacy high-fidelity analysis codes with modern software provides the improved interface necessary to enable rapid coupled SRM ballistics and vehicle trajectory analysis. Low cost trade studies demonstrate the sensitivities of flight performance metrics to propulsion characteristics. Incorporating high fidelity analysis from SPP into vehicle design reduces performance margins and improves reliability. By flying an SRM designed with the same assumptions as the rest of the vehicle, accurate comparisons can be made between competing architectures. In summary, this flexible workflow is a critical component to designing a versatile launch vehicle model that can accommodate a volatile

  3. Solid Rocket Fuel Constitutive Theory and Polymer Cure

    Ream, Robert


    Solid Rocket Fuel is a complex composite material for which no general constitutive theory, based on first principles, has been developed. One of the principles such a relation would depend on is the morphology of the binder. A theory of polymer curing is required to determine this morphology. During work on such a theory an algorithm was developed for counting the number of ways a polymer chain could assemble. The methods used to develop and check this algorithm led to an analytic solution to the problem. This solution is used in a probability distribution function which characterizes the morphology of the polymer.

  4. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    Moore, Carleton J.


    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

  5. Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant%Nmerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

    SHAFQAT Wahab; XIE Kan; LIU Yu


    Multi-phase flow field simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM)is in- vestigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and form liquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flow simulations have been performed on a large scale motor,to observe the effect of the flowfield on the chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an expo- nential relationship exists between the droplet diameter,d and mass fraction of droplets with diameter greater than d have been used for the simulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-1 00μm are used,as being the most common droplets.In this approach the complete range of particle sizes is divided into a set of discrete size ranges,each to be defined by single stream that is part of the group.Roe scheme-flux differencing splitting based on approximate Riemann problem has been used to simulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect, Spalart-Allmaras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRM flow dynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the parti- cles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

  6. Sensitivity of solid rocket propellants for card gap test

    Kimura, Eishu; Oyumi, Yoshio (Japan Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solid propellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP- and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here. (orig.)

  7. High-speed schlieren imaging of rocket exhaust plumes

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


    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.

  8. Solid rocket booster internal flow analysis by highly accurate adaptive computational methods

    Huang, C. Y.; Tworzydlo, W.; Oden, J. T.; Bass, J. M.; Cullen, C.; Vadaketh, S.


    The primary objective of this project was to develop an adaptive finite element flow solver for simulating internal flows in the solid rocket booster. Described here is a unique flow simulator code for analyzing highly complex flow phenomena in the solid rocket booster. New methodologies and features incorporated into this analysis tool are described.

  9. Environmental impact statement Space Shuttle advanced solid rocket motor program


    The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site. Sites being considered for the new facilities include John C. Stennis Space Center, Hancock County, Mississippi; the Yellow Creek site in Tishomingo County, Mississippi, which is currently in the custody and control of the Tennessee Valley Authority; and John F. Kennedy Space Center, Brevard County, Florida. TVA proposes to transfer its site to the custody and control of NASA if it is the selected site. All facilities need not be located at the same site. Existing facilities which may provide support for the program include Michoud Assembly Facility, New Orleans Parish, Louisiana; and Slidell Computer Center, St. Tammany Parish, Louisiana. NASA's preferred production location is the Yellow Creek site, and the preferred test location is the Stennis Space Center.

  10. Exploring the Solid Rocket Boosters and Properties of Matter

    Moffett, Amy


    I worked for the United Space Alliance, LLC (USA) with the Solid Rocket Booster (SRB) Materials and Process engineers (M&P). I was assigned a project in which I needed to research and collect chemical and physical properties information, material safety data sheets (MSDS), and other product information from the vendor's websites and existing "inhouse" files for a select group of materials used in building and refurbishing the SRBs. This information was then compiled in a report that summarized the information collected. My work site was at the Kennedy Space Center (KSC). This allowed for many opportunities to visit and tour sites operated by NASA, by USA, and by the Air Force. This included the vehicle assembly building (VAB), orbital processing facilities (OPF), the crawler with the mobile launch pad (MLP), and the SRB assembly and refurbishment facility (ARF), to name a few. In addition, the launch, of STS- 117 took place within the first week of employment allowing a day by day following of that mission including post flight operations for the SRBs. Two Delta II rockets were also launched during these 7 weeks. The sights were incredible and the operations witnessed were amazing. I learned so many things I never knew about the entire program and the shuttle itself. The entire experience, especially my work with the SRB materials, inspired my plan for implementation into the classroom.

  11. Real-time radiography of Titan IV Solid Rocket Motor Upgrade (SRMU) static firing test QM-2

    Dolan, K.W.; Curnow, G.M.; Perkins, D.E.; Schneberk, D.J.; Costerus, B.W.; La Chapell, M.J.; Turner, D.E.; Wallace, P.W.


    Real-time radiography was successfully applied to the Titan-IV Solid Rocket Motor Upgrade (SRMU) static firing test QM-2 conducted February 22, 1993 at Phillips Laboratory, Edwards AFB, CA. The real-time video data obtained in this test gave the first incontrovertible evidence that the molten slag pool is low (less than 5 to 6 inches in depth referenced to the bottom of the aft dome cavity) before T + 55 seconds, builds fairly linearly from this point in time reaching a quasi-equilibrium depth of 16 to 17 inches at about T + 97 seconds, which is well below the top of the vectored nozzle, and maintains that level until T + 125 near the end motor burn. From T + 125 seconds to motor burn-out at T + 140 seconds the slag pool builds to a maximum depth of about 20 to 21 inches, still well below the top of the nozzle. The molten slag pool was observed to interact with motions of the vectored nozzle, and exhibit slosh and wave mode oscillations. A few slag ejection events were also observed.

  12. Scale Effects on Solid Rocket Combustion Instability Behaviour

    David R. Greatrix


    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.

  13. Ecological effects and environmental fate of solid rocket exhaust

    Nimmo, B.; Stout, I. J.; Mickus, J.; Vickers, D.; Madsen, B.


    Specific target processes were classified as to the chemical, chemical-physical, and biological reactions and toxic effects of solid rocket emissions within selected ecosystems at Kennedy Space Center. Exposure of Citris seedlings, English peas, and bush beans to SRM exhaust under laboratory conditions demonstrated reduced growth rates, but at very high concentrations. Field studies of natural plant populations in three diverse ecosystems failed to reveal any structural damage at the concentration levels tested. Background information on elemental composition of selected woody plants from two terrestrial ecosystems is reported. LD sub 50 for a native mouse (peromysous gossypinus) exposed to SRM exhaust was determined to be 50 ppm/g body weight. Results strongly indicate that other components of the SRM exhaust act synergically to enhance the toxic effects of HCl gas when inhaled. A brief summary is given regarding the work on SRM exhaust and its possible impact on hatchability of incubating bird eggs.

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

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


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

  15. Internal Flow Simulation of High-Performance Solid Rockets using a k-ωTurbulence Model



    @@ For technological reasons many high-performance solid rocket motors are made from segmented propellant grains with non-uniform port geometry. In this paper parametric studies have been carried out to examine the geometric dependence of transient flow features in solid rockets with non-uniform ports. Numerical computations have been carried out in an inert simulator of solid propellant rocket motor with the aid of a standard k-ω turbulence model. It was seen that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that the damping of the flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the starting transient of solid rockets.

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

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


    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.

  17. Solid propellant rocket motor internal ballistics performance variation analysis, phase 3

    Sforzini, R. H.; Foster, W. A., Jr.; Murph, J. E.; Adams, G. W., Jr.


    Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

  18. Rocket

    K. Karmarkar


    Full Text Available The rockets of World War II represented, not the invention of a new weapon, but the modernization of a very old one. As early as 1232 A.D, the Chinese launched rockets against the Mongols. About a hundred years later the knowledge of ledge of rockets was quite widespread and they were used to set fire to buildings and to terrorize the enemy. But as cannon developed, rockets declined in warfare. However rockets were used occasionally as weapons till about 1530 A.D. About this time improvements in artillery-rifled gun barrel and mechanism to absorb recoil-established a standard of efficiency with which rockets could not compare until World War II brought pew conditions

  19. Shock wave fabricated ceramic-metal nozzles

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


    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. Shock wave fabricated ceramic-metal nozzles

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


    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

  1. Thermodynamic cycle analysis of solid propellant air-turbo-rocket

    CHEN Xiang; CHEN Yu-chun; TU Qiu-ye; ZHANG Hong; CAI Yuan-hu


    Solid propellant air-turbo-rocket (SPATR) is an air-breathing propulsion system. A numerical model of performance and characteristics analysis for SPATR was presented and the corresponding computer program was written according to the operation characteristics of SPATR. The influence on the SPATR performance at design point caused by the gas generator exit parameters and compressor pressure ratio had been computed and analyzed in detail. The off-design perform-ance of SPATR at sea level and high altitude had also been computed. The performance of thrust and specific impulse for SPATR with different solid propellant had been compared at off-design points, and the off-design performance comparison had been made between fuel-rich and oxygen-rich. The computation results indicated that SPATR operates within wide range of Maeh number (0 ~3) and altitude (0~12 km), and SPATR possesses high specific thrust (1 200 N/(kg/s)) and high specific impulse (7000 N/ (kg/s)) when fuel-air ratio of combustor equals fuel-air ratio.

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

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


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

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

    Si-Yuan Chen


    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.

  4. Solid Rocket Booster Hydraulic Pump Port Cap Joint Load Testing

    Gamwell, W. R.; Murphy, N. C.


    The solid rocket booster uses hydraulic pumps fabricated from cast C355 aluminum alloy, with 17-4 PH stainless steel pump port caps. Corrosion-resistant steel, MS51830 CA204L self-locking screw thread inserts are installed into C355 pump housings, with A286 stainless steel fasteners installed into the insert to secure the pump port cap to the housing. In the past, pump port cap fasteners were installed to a torque of 33 Nm (300 in-lb). However, the structural analyses used a significantly higher nut factor than indicated during tests conducted by Boeing Space Systems. When the torque values were reassessed using Boeing's nut factor, the fastener preload had a factor of safety of less than 1, with potential for overloading the joint. This paper describes how behavior was determined for a preloaded joint with a steel bolt threaded into steel inserts in aluminum parts. Finite element models were compared with test results. For all initial bolt preloads, bolt loads increased as external applied loads increased. For higher initial bolt preloads, less load was transferred into the bolt, due to external applied loading. Lower torque limits were established for pump port cap fasteners and additional limits were placed on insert axial deformation under operating conditions after seating the insert with an initial preload.

  5. NDE of Space Shuttle Solid Rocket Motor field joint

    Johnston, Patrick H.

    One of the most critical areas for inspection in the Space Shuttle Solid Rocket Motors is the bond between the steel case and rubber insulation in the region of the field joints. The tang-and-clevis geometry of the field joints is sufficiently complex to prohibit the use of resonance-based techniques. One approach we are investigating is to interrogate the steel-insulation bondline in the tang and clevis regions using surface-travelling waves. A low-frequency contact surface wave transmitting array transducer is under development at our laboratory for this purpose. The array is placed in acoustic contact with the steel and surface waves are launched on the inside surface or the clevis leg which propagate along the steel-insulation interface. As these surface waves propagate along the bonded surface, the magnitude of the ultrasonic energy leaking into the steel is monitored on the outer surface of the case. Our working hypothesis is that the magnitude of energy received at the outer surface of the case is dependent upon the integrity of the case-insulation bond, with less attenuation for propagation along a disbond due to imperfect acoustic coupling between the steel and rubber. Measurements on test specimens indicate a linear relationship between received signal amplitude and the length of good bend between the transmitter and receiver, suggesting the validity of this working hypothesis.

  6. Lidar measurements of solid rocket propellant fire particle plumes.

    Brown, David M; Brown, Andrea M; Willitsford, Adam H; Dinello-Fass, Ryan; Airola, Marc B; Siegrist, Karen M; Thomas, Michael E; Chang, Yale


    This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air.

  7. Viscoelastic Modelling of Solid Rocket Propellants using Maxwell Fluid Model

    Himanshu Shekhar


    Full Text Available Maxwell fluid model consisting of a spring and a dashpot in series is applied for viscoelastic characterisation of solid rocket propellants. Suitable values of spring constant and damping coefficient wereemployed by least square variation of errors for generation of complete stress-strain curve in uniaxial tensile mode for case-bonded solid propellant formulations. Propellants from the same lot were tested at different strain rates. It was observed that change in spring constant, representing elastic part was very small with strain rate but damping constant varies significantly with variation in strain rate. For a typical propellant formulation, when strain rate was raised from 0.00037/s to 0.185/s, spring constant K changed from 5.5 MPato 7.9 MPa, but damping coefficient D was reduced from 1400 MPa-s to 4 MPa-s. For all strain rates, stress-strain curve was generated using Maxwell model and close matching with actual test curve was observed.This indicates validity of Maxwell fluid model for uniaxial tensile testing curves of case-bonded solid propellant formulations. It was established that at higher strain rate, damping coefficient becomes negligible as compared to spring constant. It was also observed that variation of spring constant is logarithmic with strain rate and that of damping coefficient follows power law. The correlation coefficients were introduced to ascertain spring constants and damping coefficients at any strain rate from that at a reference strain rate. Correlationfor spring constant needs a coefficient H, which is function of propellant formulation alone and not of test conditions and the equation developeds K2 = K1 + H ´ ln{(de2/dt/(de1/dt}. Similarly for damping coefficient D also another constant S is introduced and prediction formula is given by D2 = D1 ´ {(de2/dt/(de1/dt}S.Evaluating constants H and S at different strain rates validate this mathematical formulation for differentpropellant formulations

  8. Shuttle Redesigned Solid Rocket Motor aluminum oxide investigations

    Blomshield, Fred S.; Kraeutle, Karl J.; Stalnaker, Richard A.


    During the launch of STS-54, a 15 psi pressure blip was observed in the ballistic pressure trace of one of the two Space Shuttle Redesigned Solid Rocket Motors (RSRM). One possible scenario for the observed pressure increase deals with aluminum oxide slag formation in the RSRM. The purpose of this investigation was to examine changes which may have occurred in the aluminum oxide formation in shuttle solid propellant due to changes in the ammonium perchlorate. Aluminum oxide formation from three propellants, all having the same formulation, but containing ammonium perchlorate from different manufacturers, will be compared. Three methods have been used to look for possible differences among the propellants. The first method was to examine window bomb movies of the propellants burning at 100, 300 and 600 psia. The motor operating pressure during the pressure blip was around 600 psia. The second method used small samples of propellant which were fired in a combustion bomb which quenched the burning aluminum particles soon after they left the propellant surface. The bomb was fired in both argon and Nitrogen atmospheres at various pressures. Products from this device were examined by optical microscopy. The third method used larger propellant samples fired into a particle collection device which allowed the aluminum to react and combust more completely. This device was pressurized with Nitrogen to motor operating pressures. The collected products were subdivided into size fractions by screening and sedimentation and analyzed optically with an optical microscope. the results from all three methods indicate very small changes in the size distribution of combustion products.

  9. Maturation of Structural Health Management Systems for Solid Rocket Motors Project

    National Aeronautics and Space Administration — Solid rocket motor cases are subject to a variety of external environmental and loading conditions from cradle-to-grave. These conditions can significantly impact...

  10. Electrets used in measuring rocket exhaust effluents from the space shuttle's solid rocket booster during static test firing, DM-3

    Susko, M.


    The purpose of this experimental research was to compare Marshall Space Flight Center's electrets with Thiokol's fixed flow air samplers during the Space Shuttle Solid Rocket Booster Demonstration Model-3 static test firing on October 19, 1978. The measurement of rocket exhaust effluents by Thiokol's samplers and MSFC's electrets indicated that the firing of the Solid Rocket Booster had no significant effect on the quality of the air sampled. The highest measurement by Thiokol's samplers was obtained at Plant 3 (site 11) approximately 8 km at a 113 degree heading from the static test stand. At sites 11, 12, and 5, Thiokol's fixed flow air samplers measured 0.0048, 0.00016, and 0.00012 mg/m3 of CI. Alongside the fixed flow measurements, the electret counts from X-ray spectroscopy were 685, 894, and 719 counts. After background corrections, the counts were 334, 543, and 368, or an average of 415 counts. An additional electred, E20, which was the only measurement device at a site approximately 20 km northeast from the test site where no power was available, obtained 901 counts. After background correction, the count was 550. Again this data indicate there was no measurement of significant rocket exhaust effluents at the test site.

  11. Ballistic anomalies in solid rocket motors due to migration effects

    Pröbster, M.; Schmucker, R. H.

    Double base and composite propellants are generally used for rocket motors, whereby double base propellants basically consist of nitrocellulose plasticized with an explosive plasticizer, mostly nitroglycerine, and in some cases with an additional inert plasticizer and ballistic additives. Composite propellants consist of an oxidizer like ammonium perchlorate and of aluminum, binder and plasticizer and often contain liquid or solid burning rate catalysts. A common feature of both propellants is that they contain smaller or larger amounts of chemically unbonded liquid species which tend to migrate. If these propellants loose part of the plasticizer by migration into the insulation layer, not only will there be a change in mechanical propellant properties but also the bond between propellant and insulation may degrade. However, depending on the severity of these effects, the change in the ballistic properties of the propellant grain caused by plasticizer migration may be of even more importance. In the past, most emphasis was placed on the behaviour of end-burning configurations. However, more recent theoretical and experimental studies revealed that not only for end-burning grain configurations but also for internal burning configurations there is a common effect which is responsible for ballistic anomalies: migration of liquid species from the propellant into the insulation. By using a plasticizer equilibrated insulation in an internal burning configuration the liquid species migration and thus the previously observed ballistic anomalies are avoided. Using this approach for end-burning configurations provides similar positive results. The various factors affecting plasticizer migration are studied and discussed, and several methods to prevent liquid species migration are described as well as methods to obtain plasticizer resistant insulations.

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

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


    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


    Muhammad; M.; R.; Qureshi; Chao; Zhu; Chao-Hsin; Lin; Liang-Shih; Fan


    A three-dimensional simulation study is performed for investigating the hydrodynamic behaviors of a cross-flow liquid nitrogen spray injected into an air-fluidized catalytic cracking (FCC) riser of rectangular cross-section. Rectangular nozzles with a fixed aspect ratio but different fan angles are used for the spray feeding. While our numerical simulation reveals a generic three-phase flow structure with strong three-phase interactions under rapid vaporization of sprays, this paper tends to focus on the study of the effect of nozzle fan angle on the spray coverage as well as vapor flux distribution by spray vaporization inside the riser flow. The gas-solid (air-FCC) flow is simulated using the multi-fluid method while the evaporating sprays (liquid nitrogen) are calculated using the Lagrangian trajectory method, with a strong two-way coupling between the Eulerian gas-solid flow and the Lagrangian trajectories of spray. Our simulation shows that the spray coverage is basically dominated by the spray fan angle. The spray fan angle has a very minor effect on spray penetration. The spray vaporization flux per unit area of spray coverage is highly non-linearly distributed along the spray penetration. The convection of gas-solid flow in a riser leads to a significant downward deviation of vapor generated by droplet vaporization, causing a strong recirculating wake region in the immediate downstream area of the spray.

  14. Solid rocket motor fire tests: Phases 1 and 2

    Chang, Yale; Hunter, Lawrence W.; Han, David K.; Thomas, Michael E.; Cain, Russell P.; Lennon, Andrew M.


    JHU/APL conducted a series of open-air burns of small blocks (3 to 10 kg) of solid rocket motor (SRM) propellant at the Thiokol Elkton MD facility to elucidate the thermal environment under burning propellant. The propellant was TP-H-3340A for the STAR 48 motor, with a weight ratio of 71/18/11 for the ammonium perchlorate, aluminum, and HTPB binder. Combustion inhibitor applied on the blocks allowed burning on the bottom and/or sides only. Burns were conducted on sand and concrete to simulate near-launch pad surfaces, and on graphite to simulate a low-recession surface. Unique test fixturing allowed propellant self-levitation while constraining lateral motion. Optics instrumentation consisted of a longwave infrared imaging pyrometer, a midwave spectroradiometer, and a UV/visible spectroradiometer. In-situ instrumentation consisted of rod calorimeters, Gardon gauges, elevated thermocouples, flush thermocouples, a two-color pyrometer, and Knudsen cells. Witness materials consisted of yttria, ceria, alumina, tungsten, iridium, and platinum/rhodium. Objectives of the tests were to determine propellant burn characteristics such as burn rate and self-levitation, to determine heat fluxes and temperatures, and to carry out materials analyses. A summary of qualitative results: alumina coated almost all surfaces, the concrete spalled, sand moisture content matters, the propellant self-levitated, the test fixtures worked as designed, and bottom-burning propellant does not self-extinguish. A summary of quantitative results: burn rate averaged 1.15 mm/s, thermocouples peaked at 2070 C, pyrometer readings matched MWIR data at about 2400 C, the volume-averaged plume temperatures were 2300-2400 C with peaks of 2400-2600 C, and the heat fluxes peaked at 125 W/cm2. These results are higher than other researchers' measurements of top-burning propellant in chimneys, and will be used, along with Phase 3 test results, to analyze hardware response to these environments, including General




    Full Text Available In this study, three dimensional modelling of extrusion forming of a double base solid rocket propellant is performed on Ansys® finite element simulation package. For the purpose of initial model construction and later comparisons with elastoviscoplastik model, the solid propellant is assumed to obey the elastic-plastic material response during the direct extrusion process. Taking into account the contact surface behavior with Coulomb friction and geometric and material nonlinearities, an incremental large large strain solution methodology has been adapted in the simulation. The hydrostatic pressure, stress, strain, and displacement values during extrusion of the solid rocket propellant are obtained from the simulation.

  16. Multi-Dimensional Combustion Instability Analysis of Solid Propellant Rocket Motors.


    RI D-R159 314 MULTI-DIMENSIONAL COMBUSTION INSTABLITY ANALYSIS OF 1/1 I SOLID PROPELLANT ROCK.. (U) ALABAMA UNIY IN HUNTSVILLE I DEPT OF MECHANICAL...STANDARDS MlICROCOPY RESOLUTION TEST CHART 0 0 0 03 V.%% f iSR.TR. 85-0567 NULTI-DIMNSIONAL COMBUSTION INSTABILITY ANALYSIS OF SOLID PROPELLANT ROCKET...analysis of solid propellant rocket motors. This research was motivated by the need for im- provement of the current practice in combustion instability

  17. Mathematical Modelling of In-Chamber Processes in Hydrocombined Propellant Solid Rocket Motors

    Nikolai A. Obukhov


    Full Text Available The special conditions of employment of commercial rockets in the sea environment has opened up new possibilities of improving motor performance. The interesting method suggests supplying water into the running motor. This paper reports the calculations and experiments carried out with solid propellant model setups. The results prove the validity of the proposed method and allow the refinement of calculation techniques for the prediction of solid rocket motor performance characteristics. The serviceability of the solid propellant charges working in combination with water is demonstrated. A mathematical model is proposed for the operation of a hydrocombined propellant motor with water and powdered additives applied to the combustion chamber."

  18. Rocket engine high-enthalpy flow simulation using heated CO2 gas to verify the development of a rocket nozzle and combustion tests

    Takeishi, K.; Ishizaka, K.; Okamoto, J.; Watanabe, Y.


    The LE-7A engine is the first-stage engine of the Japanese-made H-IIA launch vehicle. This engine has been developed by improving and reducing the price of the LE-7 engine used in the H-II launch vehicle. In the qualification combustion tests, the original designed LE-7A (LE-7A-OR) engine experienced two major problems, a large side load in the transient state of engine start and stop and melt on nozzle generative cooling tubes. The reason for the troubles of the LE-7A-OR engine was investigated by conducting experimental and numerical studies. In actual engine conditions, the main hot gas stream is a heated steam. Furthermore, the main stream temperature in the nozzle changes from approximately 3500 K at the throat to 500 K at the exit. In such a case, the specific heat ratio changes depending on the temperature. A similarity of the Mach number should be considered when conducting a model flow test with a similar flow condition of the Mach number between an actual engine combustion test and a model flow test. High-speed flow tests were conducted using CO2 gas heated up to 673 K as a working fluid and a 1:12 sub-scaled model nozzle of the LE-7A-OR engine configuration. The problems of the side force and the conducted form of the shock waves generated in the nozzle of the LE-7A-OR engine during engine start and stop were reproduced by the model tests of experimental and numerical investigations. This study presented that the model flow test using heated CO2 gas is useful and effective in verifying the numerical analysis and the design verification before actual engine combustion tests.

  19. Solar Thermal Rocket Propulsion

    Sercel, J. C.


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

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

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


    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.

  1. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.


    Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  2. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.; Borisov Boris V.; Bondarchuk Sergey S.; Zhukov Alexander S.


    On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  3. Solid propellant processing factor in rocket motor design


    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.

  4. Formation of Vortex Structures in the Prenozzle Space of an Engine with a Vectorable Thrust Nozzle

    Volkov, K. N.; Emel'yanov, V. N.; Denisikhin, S. V.


    A numerical simulation of the hydrodynamic effects arising in the process of work of the vectorable thrust nozzle of a solid-propellant rocket engine has been performed. The fields of the flows of combustion products in the channel of a charge, the prenozzle space, and the nozzle unit were calculated for different angles of vectoring of the nozzle. The distributions of the gasdynamic parameters of the flow of combustion products in the prenozzle space, corresponding to their efflux from the cylindrical and star-shaped channels of charges, were compared. The formation of a vortex flow in the neighborhood of the back cover of the nozzle was considered.

  5. Thermographic inspection of solid-fuel rocket booster field joint components

    Thompson, Karen G.; Crisman, Elton M.


    Thermographic nondestructive evaluation techniques were investigated for possible application on Space Shuttle solid rocket booster field joint hardware. This investigation included evaluation of the clevis and tang mating surfaces for scratches and measurement of grease film thickness. The field joint insulation system was inspected for voids and disbonds.

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


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

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


    3950°C and 3928°C, respectively. Compared to refractory metal candidates such as Rhenium (W) and Tungsten (W), these carbides offer significant...45HfC-10VC at 1538 W/cm² in the LHMEL Facility laser Sample gone within ~3 seconds, ablated and blew away; carbon foam backing remained intact for...Response of 45TaC-45HfC-10VC at 1529 W/cm² in the LHMEL Facility laser Sample gone within 6 seconds, ablated and blew away; carbon foam backing

  8. Research on Instantaneous Thrust Measurement for Attitude-control Solid Rocket Motor

    OUYANG Hua-bing; WANG Jian-ping; LIN Feng; XU Wen-gan


    In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.

  9. Draft environmental impact statement: Space Shuttle Advanced Solid Rocket Motor Program


    The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site.

  10. FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor%固体火箭发动机柔性喷管有限元建模及摆动分析

    王成林; 刘勇; 文立华


    The methods of FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor were explores. The FEM model of flexible nozzle using the equivalent model of flexible joint based on the three-direction custom spring elements is build; modify the FEM model of the flexible joint according to the experiment data; and carry out the oscillation analysis of solid propellant rocket motor.%研究了固态火箭发动机柔性喷管有限元建模及摆动分析方法.利用基于自定义三向弹簧单元的柔性接头线性等效模型,建立了发动机柔性喷管有限元模型.根据试验数据对柔性接头模型进行修正,并对发动机柔性喷管进行了摆动分析.

  11. Chemically Collapsible Mandrel for Solid Rocket Motor Processing

    Dey,Abhijit; Kumar, Arvind; Sikder,Arun K; Gupta, Manoj


    ABSTRACT: Composite propellant mainly consists of two parts, binder matrix (prepolymer, plasticizer, cross linker, antioxidant and curative etc.) and solid ingredients (oxidizer, metal fuel, burn rate modifier, combustion stabilizer etc.). Its processing involves several stages like ingredient preparation (grinding, 1.1 Hazard Division - 1.1 HD), mixing (1.1 HD), casting (1.1 HD), curing (1.3 HD) and extraction (1.3 HD). Each and every process is very hazardous. Removal of any of the mentione...

  12. Impact of the Surface Morphology on the Combustion of Simulated Solid Rocket Motor

    Abdelkarim M. Hegab


    Full Text Available An advanced and intensive computational solution development is integrated with an asymptotic technique, to examine the impact of the combustion surface morphology on the generated rotational flow field in a solid rocket chamber with wide ranges of forcing frequencies. The simulated rectangular chamber is closed at one end and is open at the aft end. The upper and lower walls are permeable to allow steady and unsteady injected air to generate internal flow mimicking the flow field of the combustion gases in real rocket chamber. The frequencies of the unsteady injected flow are chosen to be very close or away from the resonance frequencies of the adapted chamber. The current study accounts for a wide range of wave numbers that reflect the complexity of real burning processes. Detailed derivation for Navier-Stokes equations at the four boundaries of the chamber is introduced in the current study. Qualitative comparison is performed with recent experimental work carried out on a two-inch hybrid rocket motor using a mixture of polyethylene and aluminum powder. The higher the percentage of aluminum powder in the mixture, the more the corrugations of the combustion surface. This trend is almost similar to the computational and analytical results of a simulated solid rocket chamber.

  13. Estimation of Pressure Index and Temperature Sensitivity Coefficient of Solid Rocket Propellants by Static Evaluation

    Himanshu Shekhar


    Full Text Available Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T-To}]Pn was considered and a method was evolved for determination of pressure index (n and temperature sensitivity coefficient (p of burning rate for solid rocket propellants from static evaluation data. Effect of pressure index and temperature sensitivity coefficient on firing curve is also depicted. Propellant grain was fired in progressive mode to cover a very wide pressure range of 50 kg/cm2 to 250 kg/cm2 and propellant burning rate index was calculated to be 0.32 in the given pressure range. Propellant grain was fired at +35 °C and –20 °C temperatures and temperature sensitivity coefficient of burning rate was calculated to be 0.27 % per °C. Since both the values were evaluated from realised static evaluation curves, these are more realistic and accurate compared to data generated by conventional methods.Defence Science Journal, 2009, 59(6, pp.666-669, DOI:

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

    Clark, John S.; Walton, James T.; Mcguire, Melissa L.


    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.

  15. ASRM radiation and flowfield prediction status. [Advanced Solid Rocket Motor plume radiation prediction

    Reardon, J. E.; Everson, J.; Smith, S. D.; Sulyma, P. R.


    Existing and proposed methods for the prediction of plume radiation are discussed in terms of their application to the NASA Advanced Solid Rocket Motor (ASRM) and Space Shuttle Main Engine (SSME) projects. Extrapolations of the Solid Rocket Motor (SRM) are discussed with respect to preliminary predictions of the primary and secondary radiation environments. The methodology for radiation and initial plume property predictions are set forth, including a new code for scattering media and independent secondary source models based on flight data. The Monte Carlo code employs a reverse-evaluation approach which traces rays back to their point of absorption in the plume. The SRM sea-level plume model is modified to account for the increased radiation in the ASRM plume due to the ASRM's propellant chemistry. The ASRM cycle-1 environment predictions are shown to identify a potential reason for the shutdown spike identified with pre-SRM staging.

  16. Probabilistic Fracture Mechanics and Optimum Fracture Control Analytical Procedures for a Reusable Solid Rocket Motor Case

    Hanagud, S.; Uppaluri, B.


    A methodology for the reliability analysis of a reusable solid rocket motor case is discussed. The analysis is based on probabilistic fracture mechanics and probability distribution for initial flaw sizes. The developed reliability analysis is used to select the structural design variables of the solid rocket motor case on the basis of minimum expected cost and specified reliability bounds during the projected design life of the case. Effects of failure prevention plans such as nondestructive inspection and the material erosion between missions are also considered in the developed procedure for selection of design variables. The reliability-based procedure can be modified to consider other similar structures of reusable space vehicle systems with different failure prevention plans.

  17. Palynological Investigation of Post-Flight Solid Rocket Booster Foreign Material

    Nelson, Linda; Jarzen, David


    Investigations of foreign material in a drain tube, from the Solid Rocket Booster (SRB) of a recent Space Shuttle mission, was identified as pollen. The source of the pollen is from deposits made by bees, collecting pollen from plants found at the Kennedy Space Center, Cape Canaveral, Florida. The pollen is determined to have been present in the frustum drain tubes before the shuttle flight. During the flight the pollen did not undergo thermal maturation.

  18. Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

    Martin, Heath Thomas


    Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

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

    Zhang Xiaoying; Chen Huandong


    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 2–6μm. Conditions at wavelengths 2.7μm and 4.3μm 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 sig-nificantly in peak radiation spectra of plume gases CO and CO2. Al2O3 particles are the major radi-ation 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 radi-ation intensity at wavebands 2.7–3.0μm and 4.2–4.6μm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle.

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

    Harris, Joshua A.; Patterson-Hine, Ann


    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.

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

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


    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.

  2. Response of selected plant and insect species to simulated solid rocket exhaust mixtures and to exhaust components from solid rocket fuels

    Heck, W. W.; Knott, W. M.; Stahel, E. P.; Ambrose, J. T.; Mccrimmon, J. N.; Engle, M.; Romanow, L. A.; Sawyer, A. G.; Tyson, J. D.


    The effects of solid rocket fuel (SRF) exhaust on selected plant and and insect species in the Merritt Island, Florida area was investigated in order to determine if the exhaust clouds generated by shuttle launches would adversely affect the native, plants of the Merritt Island Wildlife Refuge, the citrus production, or the beekeeping industry of the island. Conditions were simulated in greenhouse exposure chambers and field chambers constructed to model the ideal continuous stirred tank reactor. A plant exposure system was developed for dispensing and monitoring the two major chemicals in SRF exhaust, HCl and Al203, and for dispensing and monitoring SRF exhaust (controlled fuel burns). Plants native to Merritt Island, Florida were grown and used as test species. Dose-response relationships were determined for short term exposure of selected plant species to HCl, Al203, and mixtures of the two to SRF exhaust.

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

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


    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.

  4. Mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene

    Gligorijević Nikola I.


    Full Text Available This paper presents the procedure of uniaxial mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene (HTPB, whose mechanical properties strongly depend on temperature, strain rate, natural aging and accumulated damage. A method of processing data is presented in order to determine time-temperature shift factor and master curves for tensile strength, ultimate strain and relaxation modulus, depending on reduced time. Functional dependences of these features represent an input for structural analysis of a rocket motor propellant grain. The effects of natural aging on the mechanical properties are also considered. [Projekat Ministarstva nauke Republike Srbije, br. TR 36050: Research and development of unmanned aircraft in support of traffic infrastructure monitoring

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

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


    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.

  6. Digital Machining System for Nozzle Cooling Channel of Large Liquid Rocket Engine%大型液体火箭发动机喷管数字化铣槽加工系统

    王永青; 刘海波; 李护林; 贾振元


    Rocket nozzle is a key structural part of high-thrust liquid rocket engine. There are a hundreds of cooling channels around the nozzle, to ensure the reliable cooling and preheat the fuel. However, it is very difficult to machine the cooling channel due to large size, complex profile, low rigidity, etc. In this article, an integrated digital method for cooling channel machining composed of profile measuring, data processing and channel milling is proposed. Because of large difference between the actual contour and the design model, the channel bottom should be redesigned by using the measured geometric information. Therefore, the varying-thickness and varying-depth cooling channel of nozzle with high order contour or parametric shapes can be machined. Further, a special digital machining system is developed based on an open numerical control platform for the dual-channel vertical milling machine. Finally, an experiment utilizing a typical rocket nozzle is implemented to verify the feasibility of the system. It has been proved that the digital machining system can meet the machining requirements for liquid rocket engine nozzle.%针对大型液体火箭发动机喷管几何尺寸大、廓形复杂、结构刚度低致使其冷却通道加工质量难以保证的难题,提出一种集“测量-数据处理-铣槽”于一体的喷管冷却通道数字化加工新方法,并在开放式数控平台上开发出喷管专用数字化铣槽加工系统.该方法利用喷管几何外廓的实际测量信息再设计出槽底曲面,进而实现高次曲线或参数曲线廓形、变壁厚变槽深喷管冷却通道的数字化加工.通过某型号火箭发动喷管的实际加工,表明所研制的双通道立式铣槽加工专用装备与系统可满足我国新一代大推力液体火箭发动机喷管冷却通道高质量、高效、高可靠的制造要求.

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

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


    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

  8. Ignition of Liquid Fuel Spray and Simulated Solid Rocket Fuel by Photoignition of Carbon Nanotube Utilizing a Camera Flash


    Badakhshan A1 , Danczyk S. A.2, Wirth D.3 and Pilon L. 3 Abstract We have studied the ignition of fuel sprays and simulated solid rocket fuels (SRF...photoignition of solid oxidizer/CNT mixtures exposed to a flash of light. The flash source was a commercial studio flash lamp with a rated maximum

  9. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

  10. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

  11. Internal Ballistic Code for Solid Rocket Motors using Minimum Distance Function for Grain Burnback

    Afroz Javed


    Full Text Available A computer code has been developed for internal ballistic performance evaluation of solid rocket motors, using minimum distance function (MDF approach for prediction of geometry evolution. This method can handle any complex geometry without the need to define different geometrical shapes and their evolution as used in several existing analytical geometry evolution-based methodologies. The code is validated with both experimental results published in literature, as well as for solid rocket motors of tactical and strategic missiles and a very good match is obtained with static test results. The output of the code gives p-t (pressure-time curve as well as the detailed parameters of the flow along the axial direction, and geometries in the form of mesh file, which can be further used as input to codes for CFD analysis.Defence Science Journal, Vol. 65, No. 3, May 2015, pp.181-188, DOI:

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

    Yang, H. Q.; West, Jeff


    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.

  13. Nozzle fabrication technique

    Wells, Dennis L. (Inventor)


    This invention relates to techniques for fabricating hour glass throat or convergent divergent nozzle shapes, and more particularly to new and improved techniques for forming rocket nozzles from electrically conductive material and forming cooling channels in the wall thereof. The concept of positioning a block of electrically conductive material so that its axis is set at a predetermined skew angle with relation to a travelling electron discharge machine electrode and thereafter revolving the body about its own axis to generate a hyperbolic surface of revolution, either internal or external is novel. The method will generate a rocket nozzle which may be provided with cooling channels using the same control and positioning system. The configuration of the cooling channels so produced are unique and novel. Also the method is adaptable to nonmetallic material using analogous cutting tools, such as, water jet, laser, abrasive wire and hot wire.

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

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


    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

  15. Regarding the evaluation of the solid rocket propellant response function to pressure coupling

    Ioan ION


    Full Text Available High frequency combustion instabilities imply a major risk for the solid rocket motor stableworking and they are directly linked to the propellant response to chamber pressure coupling. Thisarticle discusses a laboratory testing method for the measurement and evaluation of the pressurecoupled response for non-metalized propellants in a first stage. Experimental researches were donewith an adequate setup, built and improved in our lab, able to evaluate the propellant response byinterpreting the pressure oscillations damping in terms of propellant response. Our paper aims atdefining a linearized one-dimensional flow study model to analyze the disturbed operation of the solidpropellant rocket motors. Based on the applied model we can assert that the real part of propellantresponse is a function of the oscillations damping, acoustic energy in the motor chamber and variouslosses in the burning chamber. The imaginary part of propellant response mainly depends on thenormalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency.Our research purpose was obviously to minimize the risk of the combustion instabilities effects on therocket motors working, by experimental investigations using jet modulating techniques and sustainedby an interesting study model based on the perturbation method.

  16. Ozone depletion in the plume of a solid-fuelled rocket

    B. C. Krüger

    Full Text Available The local effects of the emission of a solid-fuelled rocket on the stratospheric ozone concentration have been investigated by photochemical model calculations. A one-dimensional horizontal model has been applied which calculates the trace gas composition at a single atmospheric altitude spatially resolved around the exhaust plume. Different cases were tested for the emissions of the Space Shuttle concerning the composition of the exhaust and the effects of heterogeneous reactions on atmospheric background aerosol.

    The strongest depletion of ozone is achieved when a high amount of the emitted chlorine is Cl2. If it is purely HCl, the effect is smallest, though in this case the heterogeneous reactions show their largest influence. From the results it may be estimated whether ozone depletion caused by rocket launches can be detected by satellite instruments. It appears that the chance of coincidental detection of such an event is rather small.

  17. Pressure oscillations and instability of working processes in the combustion chambers of solid rocket motors

    Emelyanov, V. N.; Teterina, I. V.; Volkov, K. N.; Garkushev, A. U.


    Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

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

    Zolotorev Nikolay


    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.

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

    David R. Greatrix


    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.

  20. Numerical techniques for solving nonlinear instability problems in smokeless tactical solid rocket motors. [finite difference technique

    Baum, J. D.; Levine, J. N.


    The selection of a satisfactory numerical method for calculating the propagation of steep fronted shock life waveforms in a solid rocket motor combustion chamber is discussed. A number of different numerical schemes were evaluated by comparing the results obtained for three problems: the shock tube problems; the linear wave equation, and nonlinear wave propagation in a closed tube. The most promising method--a combination of the Lax-Wendroff, Hybrid and Artificial Compression techniques, was incorporated into an existing nonlinear instability program. The capability of the modified program to treat steep fronted wave instabilities in low smoke tactical motors was verified by solving a number of motor test cases with disturbance amplitudes as high as 80% of the mean pressure.

  1. Scale Effects on Quasi-Steady Solid Rocket Internal Ballistic Behaviour

    David R. Greatrix


    Full Text Available 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.

  2. Alternate propellants for the space shuttle solid rocket booster motors. [for reducing environmental impact of launches


    As part of the Shuttle Exhaust Effects Panel (SEEP) program for fiscal year 1973, a limited study was performed to determine the feasibility of minimizing the environmental impact associated with the operation of the solid rocket booster motors (SRBMs) in projected space shuttle launches. Eleven hypothetical and two existing limited-experience propellants were evaluated as possible alternates to a well-proven state-of-the-art reference propellant with respect to reducing emissions of primary concern: namely, hydrogen chloride (HCl) and aluminum oxide (Al2O3). The study showed that it would be possible to develop a new propellant to effect a considerable reduction of HCl or Al2O3 emissions. At the one extreme, a 23% reduction of HCl is possible along with a ll% reduction in Al2O3, whereas, at the other extreme, a 75% reduction of Al2O3 is possible, but with a resultant 5% increase in HCl.

  3. Numerical Simulation of a Dual Pulse Solid Rocket Motor Flow Field

    Afroz Javed


    Full Text Available Numerical simulations are carried out for the internal flow field of a dual pulse solid rocket motor port to understand the flow behaviour. Three dimensional Reynolds Averaged Navier Stokes equations are solved alongwith shear stress transport turbulence model using commercial code. The combustion gas is assumed as a mixture of alumina and gases and single phase flow calculations are done with the thermo chemical properties provided for the mixture. The simulation captures all the essential features of the flow field. The flow accelerates through the pulse separation device (PSD port and high temperature and high velocity gas is seen to impinge the motor wall near the PSD port. The overall total pressure drop through motor port and through PSD is found to be moderate.Defence Science Journal, 2012, 62(6, pp.369-374, DOI:

  4. Preliminary design and optimization of slotted tube grain for solid rocket motor


    In this paper,design and optimization technique of slotted tube grain for solid rocket motors has been discussed.In doing so,the design objectives and constraints have been set,geometric parameters identified,performance prediction parameters calculated,thereafter preliminary designs completed and finally optimal design reached.Geometric model for slotted tube grain configuration has been developed.Average thrust has been taken as the objective function with constraints of burning time,mass of propellant,fixed length and diameter of chamber case.Lumped parameter method has been used for calculating the performance prediction parameters.A set of preliminary designs has been completed and an analysis of these results conducted.Although all the preliminary results fulfill the design requirements in terms of objective function and constraints,however in order tO attain the optimal design,Sequen-tial quadratic programming optimization technique has been adopted.As the slotted tube grain ge-ometry is totally dependent upon various independent variables and each of these variables has a bearing on explicit characteristic of grain designing,hence affects of the independent variables on performance parameters have been examined,thus variation laws have been developed.Basing on the variation laws and the analysis of preliminary design results,upper and lower limits have been defined for the independent geometric variables and an initial guess provided for conducting optimi-zation.Resuhs attained exhibits that an optimal result has been attained and the value of objective function has been maximized.All the design constraint limits have also been met while ensuring sound values of volumetric loading fraction,web fraction and neutrality.This methodology of design and optimization of slotted tube grain for solid rocket motors can be used by engineers as a reference guide for actual design and engineering purposes.

  5. The materials and elements production practice of counter-erosional and thermal protection system of the SPR-solid-propellant sustainer nozzle

    Shkurenko, V. M.


    This paper presents the production scheme for heat- and erosion-protective carbon plastic materials for heat shield elements of solid-propellant nozzles. Attention is also given the method of manufacturing adhesive joint assemblies, and the production scheme is included.

  6. Multi-phase flow effect on SRM nozzle flow field and thermal protection materials

    SHAFQAT Wahab; XIE Kan; LIU Yu


    Multi-phase flow effect generated from the combustion of aluminum based com-posite propellant was performed on the thermal protection material of solid rocket motor (SRM) nozzle. Injection of alumina (Al2O3) particles from 5% to 10% was tried on SRM nozzle flow field to see the influence of multiphase flow on heat transfer computations. A coupled, time resolved CFD (computational fluid dynamics) approach was adopted to solve the conjugate problem of multi-phase fluid flow and heat transfer in the solid rocket motor nozzle. The governing equations are discretized by using the finite volume method. Spalart-Allmaras (S-A) turbulence model was employed. The computation was executed on the dif-ferent models selected for the analysis to validate the temperature variation in the throat in-serts and baking material of SRM nozzle. Comparison for temperatures variations were also carried out at different expansion ratios of nozzle. This paper also characterized the advanced SRM nozzle composites material for their high thermo stability and their high thermo me-chanical capabilities to make it more reliable simpler and lighter.

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

    Heitkotter, Robert H


    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.

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

    Stefanski, Philip L.


    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.

  9. Solid-propellant rocket motor internal ballistic performance variation analysis, phase 2

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


    The Monte Carlo method was used to investigate thrust imbalance and its first time derivative throughtout the burning time of pairs of solid rocket motors firing in parallel. Results obtained compare favorably with Titan 3 C flight performance data. Statistical correlations of the thrust imbalance at various times with corresponding nominal trace slopes suggest several alternative methods of predicting thrust imbalance. The effect of circular-perforated grain deformation on internal ballistics is discussed, and a modified design analysis computer program which permits such an evaluation is presented. Comparisons with SRM firings indicate that grain deformation may account for a portion of the so-called scale factor on burning rate between large motors and strand burners or small ballistic test motors. Thermoelastic effects on burning rate are also investigated. Burning surface temperature is calculated by coupling the solid phase energy equation containing a strain rate term with a model of gas phase combustion zone using the Zeldovich-Novozhilov technique. Comparisons of solutions with and without the strain rate term indicate a small but possibly significant effect of the thermoelastic coupling.

  10. Studies on Stress-Strain Curves of Aged Composite Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical property evaluation of composite solid rocket propellants is used as a quick quality control tool for propellant development and production. However, stress-strain curves from uni-axial tensile testing can be utilised to assess the shelf-life of propellants also. Composite propellants (CP of two varieties cartridge-loaded (CLCP and case-bonded (CBCP are utilized in rocket and missile applications. Both classes of propellants were evaluated for mechanical properties namely tensile strength, modulus and percentage elongation using specimens conforming to ASTM D638 type IV at different ageing time. Both classes of propellants show almost identical variation in various mechanical properties with time. Tensile strength increases with time for both classes of propellants and percentage elongation reduces. Initial modulus is also found to decrease with time. Tensile strength is taken as degradation criteria and it is observed that CLCP has slower degradation rate than CBCP. This is because of two facts–(i higher initial tensile strength of CLCP (1.39 MPa compared to CBCP (0.665 MPa and (ii lower degradation rate of CLCP (0.0014 MPa/day with respect to CBCP (0.0025 MPa/day. For the studied composite propellants, a degradation criterion in the form of percentage change in tensile strength is evaluated and shelf life for different degradation criteria is tabulated for quick reference.Defence Science Journal, 2012, 62(2, pp.90-94, DOI:

  11. High temperature reformation of aluminum and chlorine compounds behind the Mach disk of a solid-fuel rocket exhaust

    Park, C.


    Chemical reactions expected to occur among the constituents of solid-fuel rocket engine effluents in the hot region behind a Mach disk are analyzed theoretically. With the use of a rocket plume model that assumes the flow to be separated in the base region, and a chemical reaction scheme that includes evaporation of alumina and the associated reactions of 17 gas species, the reformation of the effluent is calculated. It is shown that AlClO and AlOH are produced in exchange for a corresponding reduction in the amounts of HCl and Al2O3. For the case of the space shuttle booster engines, up to 2% of the original mass of the rocket fuel can possibly be converted to these two new species and deposited in the atmosphere between the altitudes of 10 and 40 km. No adverse effects on the atmospheric environment are anticipated with the addition of these two new species.

  12. Integral throat entrance development, qualification and production for the Antares 3 nozzle

    Clayton, F. I.; Dirling, R. B.; Eitman, D. A.; Loomis, W. C.


    Although design analyses of a G-90 graphite integral throat entrance for the Antares 3 solid rocket motor nozzle indicated acceptable margins of safety, the nozzle throat insert suffered a thermostructural failure during the first development firing. Subsequent re-analysis using properties measured on material from the same billet as the nozzle throat insert showed negative margins. Carbon-carbon was investigated and found to result in large positive margins of safety. The G-90 graphite was replaced by SAI fast processed 4-D material which uses Hercules HM 10000 fiber as the reinforcement. Its construction allows powder filling of the interstices after preform fabrication which accelerates the densification process. Allied 15V coal tar pitch is then used to complete densification. The properties were extensively characterized on this material and six nozzles were subjected to demonstration, development and qualification firings.

  13. Feasibility Assessment of Thermal Barriers for RSRM Nozzle Joint Locations

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.


    Solid rockets, including the Space Shuttle solid rocket motor, are generally manufactured in large segments which are then shipped to their final destination where they are assembled. These large segments are sealed with a system of primary and secondary 0-rings to contain combustion gases inside the rocket which are at pressures of up to 900 psi and temperatures of up to 5500 F. The seals are protected from hot combustion gases by thick layers of phenolic insulation and by joint-filling compounds between these layers. Recently, though, routine inspections of nozzle-to-case joints in the Shuttle solid rocket motors during disassembly revealed erosion of the primary O-rings. Jets of hot gas leaked through gaps in the joint-filling compound between the layers of insulation and impinged on the O-rings. This is not supposed to take place, so NASA and Thiokol, the manufacturer of the rockets, initiated an investigation and found that design improvements could be made in this joint. One such improvement would involve using NASA Lewis braided thermal barriers as another level of protection for the O-ring seals against the hot combustion gases.

  14. Flow Simulation of Solid Rocket Motors. 2; Sub-Scale Air Flow Simulation of Port Flows

    Yeh, Y. P.; Ramandran, N.; Smith, A. W.; Heaman, J. P.


    The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

  15. Design and optimization of solid rocket motor Finocyl grain using simulated annealing

    Ali Kamran; LIANG Guo-zhu


    The research effort outlined the application of a computer aided design (CAD)-centric technique to the design and optimization of solid rocket motor Finocyl (fin in cylinder) grain using simulated annealing.The proper method for constructing the grain configuration model, ballistic performance and optimizer integration for analysis was presented. Finoeyl is a complex grain configuration, requiring thirteen variables to define the geometry. The large number of variables not only complicates the geometrical construction but also optimization process. CAD representation encapsulates all of the geometric entities pertinent to the grain design in a parametric way, allowing manipulation of grain entity (web), performing regression and automating geometrical data calculations. Robustness to avoid local minima and efficient capacity to explore design space makes simulated annealing an attractive choice as optimizer. It is demonstrated with a constrained optimization of Finocyl grain geometry for homogeneous, isotropic propellant, uniform regression, and a quasi-steady, bulk mode internal ballistics model that maximizes average thrust for required deviations from neutrality.

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

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


    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.

  17. Study of plasticizer diffusion in a solid rocket motor´s bondline

    Juliano Libardi


    Full Text Available This work aims to determine the diffusion coefficient of the plasticizers dibutyl phthalate (DBP, dioctyl phthalate (DOP and dioctyl azelate (DOZ on the internal insulating layer of solid rocket motors. These plasticizers are originally present in the layers of rubber, liner and propellant, respectively. This species are not chemically bonded and tend to diffuse from propellant to insulating and vice versa. A computer program based on the mathematical model of Fick’s second Law of diffusion was developed to perform the calculus from the concentration data obtained by gas chromatographic (GC analyses. The samples were prepared with two different adhesive liners; one conventional (LHNA and the other with barrier properties (LHNT. A common feature of both liners was that they were synthesized by the reaction of hydroxyl-terminated polybutadiene (HTPB and diisocyanates. However, a bond promoter was used to increase the crosslink density of the LHNT liner and to improve its performance as barrier against the diffusion. The effects of the diffusion of the plasticizers were also investigated by hardness analyses, which were executed on samples aged at room temperature and at 80ºC. The results showed an increase trend for the samples aged at room temperature and an opposite behavior for the tests carried out at 80ºC.

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

    David Greatrix


    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.

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

    S. Saha


    Full Text Available 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 compare closely with the literarture data. The growth rate of pressure oscillations of a cylindrical grain solid rocket motor is determined for different response functions at the fundamental longitudinal frequency. It is observed that for response function more than a critical value, the motor exhibits exponential growth rate of pressure oscillations.

  20. Experimental determination of the particle sizes in a subscale motor for application to the Ariane 5 solid rocket booster

    Traineau, J. C.; Kuentzmann, P.; Prevost, M.; Tarrin, P.; Delfour, A.

    The knowledge of the aluminum oxide particle size distribution inside the combustion chamber of a solid propellant rocket motor is an important factor for assessing the combustion stability or the slag mass accumulation in the motor. A representative subscale motor for the Ariane 5 P230 Solid Rocket Booster (SRB), in which helium is injected to quench the condensed phase reactions, has been designed and manufactured. Its use for combustion stability purpose has given the aluminum oxide particle size distribution in conditions representative of the actual Ariane 5 SRB. The experimental techniques, optical and particle capturing, have been found to give results in good agreement. A stretched distribution, with particles ranging from 1 micron to 120 microns and a maximum around 45 microns, has been demonstrated.

  1. Numerical Simulations of Canted Nozzle and Scarfed Nozzle Flow Fields

    Javed, Afroz; Chakraborty, Debasis


    Computational fluid dynamics (CFD) techniques are used for the analysis of issues concerning non-conventional (canted and scarfed) nozzle flow fields. Numerical simulations are carried out for the quality of flow in terms of axisymmetric nature at the inlet of canted nozzles of a rocket motor. Two different nozzle geometries are examined. The analysis of these simulation results shows that the flow field at the entry of the nozzles is non axisymmetric at the start of the motor. With time this asymmetry diminishes, also the flow becomes symmetric before the nozzle throat, indicating no misalignment of thrust vector with the nozzle axis. The qualitative flow fields at the inlet of the nozzles are used in selecting the geometry with lesser flow asymmetry. Further CFD methodology is used to analyse flow field of a scarfed nozzle for the evaluation of thrust developed and its direction. This work demonstrates the capability of the CFD based methods for the nozzle analysis problems which were earlier solved only approximately by making simplifying assumptions and semi empirical methods.

  2. Vibration, acoustic, and shock design and test criteria for components on the Solid Rocket Boosters (SRB), Lightweight External Tank (LWT), and Space Shuttle Main Engines (SSME)


    The vibration, acoustics, and shock design and test criteria for components and subassemblies on the space shuttle solid rocket booster (SRB), lightweight tank (LWT), and main engines (SSME) are presented. Specifications for transportation, handling, and acceptance testing are also provided.

  3. Solid Propellant Test Article (SPTA) Test Firing


    The Marshall Space Flight Center (MSFC) engineers test fired a 26-foot long, 100,000-pound-thrust solid rocket motor for 30 seconds at the MSFC east test area, the first test firing of the Modified NASA Motor (M-NASA Motor). The M-NASA Motor was fired in a newly constructed stand. The motor is 48-inches in diameter and was loaded with two propellant cartridges weighing a total of approximately 12,000 pounds. The purpose of the test was to learn more about solid rocket motor insulation and nozzle materials and to provide young engineers additional hands-on expertise in solid rocket motor technology. The test is a part of NASA's Solid Propulsion Integrity Program, that is to provide NASA engineers with the techniques, engineering tools, and computer programs to be able to better design, build, and verify solid rocket motors.

  4. Development of high temperature materials for solid propellant rocket nozzle applications. [tantalum carbides-tungsten fiber composites

    Manning, C. R., Jr.; Honeycutt, L., III


    Evaluation of tantalum carbide-tungsten fiber composites has been completed as far as weight percent carbon additions and weight percent additions of tungsten fiber. Extensive studies were undertaken concerning Young's Modulus and fracture strength of this material. Also, in-depth analysis of the embrittling effects of the extra carbon additions on the tungsten fibers has been completed. The complete fabrication procedure for the tantalum carbide-tungsten fiber composites with extra carbon additions is given. Microprobe and metallographic studies showed the effect of extra carbon on the tungsten fibers, and evaluation of the thermal shock parameter fracture strength/Young's Modulus is included.

  5. Asbestos Free Insulation Development for the Space Shuttle Solid Propellant Rocket Motor (RSRM)

    Allred, Larry D.; Eddy, Norman F.; McCool, A. A. (Technical Monitor)


    Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

  6. Time-Accurate Computational Fluid Dynamics Simulation of a Pair of Moving Solid Rocket Boosters

    Strutzenberg, Louise L.; Williams, Brandon R.


    Since the Columbia accident, the threat to the Shuttle launch vehicle from debris during the liftoff timeframe has been assessed by the Liftoff Debris Team at NASA/MSFC. In addition to engineering methods of analysis, CFD-generated flow fields during the liftoff timeframe have been used in conjunction with 3-DOF debris transport methods to predict the motion of liftoff debris. Early models made use of a quasi-steady flow field approximation with the vehicle positioned at a fixed location relative to the ground; however, a moving overset mesh capability has recently been developed for the Loci/CHEM CFD software which enables higher-fidelity simulation of the Shuttle transient plume startup and liftoff environment. The present work details the simulation of the launch pad and mobile launch platform (MLP) with truncated solid rocket boosters (SRBs) moving in a prescribed liftoff trajectory derived from Shuttle flight measurements. Using Loci/CHEM, time-accurate RANS and hybrid RANS/LES simulations were performed for the timeframe T0+0 to T0+3.5 seconds, which consists of SRB startup to a vehicle altitude of approximately 90 feet above the MLP. Analysis of the transient flowfield focuses on the evolution of the SRB plumes in the MLP plume holes and the flame trench, impingement on the flame deflector, and especially impingment on the MLP deck resulting in upward flow which is a transport mechanism for debris. The results show excellent qualitative agreement with the visual record from past Shuttle flights, and comparisons to pressure measurements in the flame trench and on the MLP provide confidence in these simulation capabilities.

  7. Feasibility evaluation of the monolithic braided ablative nozzle

    Director, Mark N.; McPherson, Douglass J., Sr.


    The feasibility of the monolithic braided ablative nozzle was evaluated as part of an independent research and development (IR&D) program complementary to the National Aeronautics and Space Administration/Marshall Space Flight Center (NASA/MSFC) Low-Cost, High-Reliability Case, Insulation and Nozzle for Large Solid Rocket Motors (LOCCIN) Program. The monolithic braided ablative nozzle is a new concept that utilizes a continuous, ablative, monolithic flame surface that extends from the nozzle entrance, through the throat, to the exit plane. The flame surface is fabricated using a Through-the-Thickness braided carbon-fiber preform, which is impregnated with a phenolic or phenolic-like resin. During operation, the braided-carbon fiber/resin material ablates, leaving the structural backside at temperatures which are sufficiently low to preclude the need for any additional insulative materials. The monolithic braided nozzle derives its potential for low life cycle cost through the use of automated processing, one-component fabrication, low material scrap, low process scrap, inexpensive raw materials, and simplified case attachment. It also has the potential for high reliability because its construction prevents delamination, has no nozzle bondlines or leak paths along the flame surface, is amenable to simplified analysis, and is readily inspectable. In addition, the braided construction has inherent toughness and is damage-tolerant. Two static-firing tests were conducted using subscale, 1.8 - 2.0-inch throat diameter, hardware. Tests were approximately 15 seconds in duration, using a conventional 18 percent aluminum/ammonium perchlorate propellant. The first of these tests evaluated the braided ablative as an integral backside insulator and exit cone; the second test evaluated the monolithic braided ablative as an integral entrance/throat/exit cone nozzle. Both tests met their objectives. Radial ablation rates at the throat were as predicted, approximately 0.017 in

  8. Hybrid Rocket Technology

    Sankaran Venugopal; K K Rajesh; V Ramanujachari


    With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems...

  9. Enhanced Large Solid Rocket Motor Understanding Through Performance Margin Testing: RSRM Five-Segment Engineering Test Motor (ETM-3)

    Huppi, Hal; Tobias, Mark; Seiler, James


    The Five-Segment Engineering Test Motor (ETM-3) is an extended length reusable solid rocket motor (RSRM) intended to increase motor performance and internal environments above the current four-segment RSRM flight motor. The principal purpose of ETM-3 is to provide a test article for RSRM component margin testing. As the RSRM and Space Shuttle in general continue to age, replacing obsolete materials becomes an ever-increasing issue. Having a five-segment motor that provides environments in excess of normal opera- tion allows a mechanism to subject replacement materials to a more severe environment than experienced in flight. Additionally, ETM-3 offers a second design data point from which to develop and/or validate analytical models that currently have some level of empiricism associated with them. These enhanced models have the potential to further the understanding of RSRM motor performance and solid rocket motor (SRM) propulsion in general. Furthermore, these data could be leveraged to support a five-segment booster (FSB) development program should the Space Shuttle program choose to pursue this option for abort mode enhancements during the ascent phase. A tertiary goal of ETM-3 is to challenge both the ATK Thiokol Propulsion and NASA MSFC technical personnel through the design and analysis of a large solid rocket motor without the benefit of a well-established performance database such as the RSRM. The end result of this undertaking will be a more competent and experienced workforce for both organizations. Of particular interest are the motor design characteristics and the systems engineering approach used to conduct a complex yet successful large motor static test. These aspects of ETM-3 and more will be summarized.

  10. Quality assurance and control in the production and static tests of the solid rocket boosters for the Space Shuttle

    Cerny, O. F.


    The paper surveys the various aspects of design and overhaul of the solid rocket boosters. It is noted that quality control is an integral part of the design specifications. Attention is given to the production process which is optimized towards highest quality. Also discussed is the role of the DCA (Defense Contract Administration) in inspecting the products of subcontractors, noting that the USAF performs this role for prime contractors. Fabrication and construction of the booster is detailed with attention given to the lining of the booster cylinder and the mixing of the propellant and the subsequent X-ray inspection.

  11. Burn-back Equations for High Volumetric Loading Single-grain Dual-thrust Rocket Propellant Configuration (Review Paper

    Himanshu Shekhar


    Full Text Available Dual-thrust mode is adopted in solid propellant rocket propulsion through tailoring of burning area, nozzle, rocket motor chamber, propellant type, multiple propellant blocks. In the present study, mathematical formulation has been evolved for generation of burning surface area with web burnt for a simple central blind hole in a solid cylindrical propellant geometry with proper partial inhibition on external and lateral surfaces. The burn-back equation has been validated by static firing and parametric study was conducted to understand effect of various control geometrical parameters. The system is utilised for high volumetric loading, single propellant, single composition, single-chamber, single nozzle dual-thrust mode of burning profiles in rocket application.Defence Science Journal, 2011, 61(2, pp.165-170, DOI:

  12. Filled Ethylene-propylene Diene Terpolymer Elastomer as ThermalInsulator for Case-bonded Solid Rocket Motors

    C. M. Bhuvaneswari


    Full Text Available Ethylene-propylene diene terpolymer (EPDM-based insulation system is being globallyused for case-bonded solid rocket motors. A study was undertaken using EPDM as base polymer,blended with hypalon and liquid EPDM and filled with fibrous and non-fibrous fillers. Theseformulations were evaluated as rocket motor insulation system. The basic objective of the studywas to develop an insulation system based on EPDM for case-bonded applications. A series ofrocket motor insulator compositions based on EPDM, filled with particulate and fibrous fillerslike precipitated silica, fumed silica, aramid, and carbon fibres have been studied for mechanical,rheological, thermal, and interface properties. Compositions based on particulate fillers wereoptimised for the filler content. Comparatively, fumed silica was found to be superior as fillerin terms of mechanical and interface properties. Addition of fibrous filler (5 parts improved thepeel strength, and reduced the thermal conductivity and erosion rate. All the compositions wereevaluated for sulphur and peroxide curing. Superior mechanical properties were achieved forsulphur-cured products, whereas peroxide-cured products exhibited an excellent ageing resistance.Rocket motors were insulated with optimised composition and propellant cast, and the motorswere evaluated by conducting static test in end-burning mode.Defence Science Journal, 2008, 58(1, pp.94-102, DOI :

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

    Goldberg, Ben E.; Wiley, Dan R.


    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.

  14. Assessment of Various Flow Solvers Used to Predict the Thermal Environment inside Space Shuttle Solid Rocket Motor Joints

    Wang, Qun-Zhen; Cash, Steve (Technical Monitor)


    It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.

  15. Solid amine-boranes as high performance hypergolic hybrid rocket fuels

    Pfeil, Mark A.

    Hypergolic hybrid rockets have the potential of providing systems that are simple, reliable, have high performance, and allow for energy management. Such a propulsion system can be applied to fields that need a single tactical motor with flexible mission requirements of either high speed to target or extended loitering. They also provide the possibility for alternative fast response dynamic altitude control systems if ignition delays are sufficiently short. Amines are the traditional fuel of choice when selecting a hypergolic combination as these tend to react readily with both nitric acid and dinitrogen tertroxide based oxidizers. It has been found that the addition of a borane adduct to an amine fuel tends to reduce the ignition delay by up to an order of magnitude with white fuming nitric acid (WFNA). The borane addition has resulted in fuels with very short ignition delays between 2-10 ms - the fastest times for an amine based fuel reacting with nitric acid based oxidizers. The incorporation of these amine-boranes, specifically ethylenediamine bisborane (EDBB), into various fuel binders has also been found to result in ignition delays between 3-10 ms - the fastest times again for amine based fuels. It was found that the addition of a borane to an amine increased theoretical performance of the amine resulting in high performance fuels. The amine-borane/fuel binder combinations also produced higher theoretical performance values than previously used hypergolic hybrid rockets. Some of the theoretical values are on par or higher than the current toxic liquid hypergolic fuels, making amine boranes an attractive replacement. The higher performing amine-borane/fuel binder combinations also have higher performance values than the traditional rocket fuels, excluding liquid hydrogen. Thus, amine-borane based fuels have the potential to influence various area in the rocket field. An EDBB/ferrocene/epoxy fuel was tested in a hypergolic hybrid with pure nitric acid as the

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

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


    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.

  17. Design, manufacture and test of the composite case for ERINT-1 solid rocket motor

    Mard, Francis


    SEP is in charge since 1989 of the ERINT-1 motor case and nozzle. The stringent missile weight and volume requirements coupled with the specification to provide an aerodynamically stable configuration over a very large Mach number range led to the need to develop a high-performance composite motor case. Development of this SRM case presented a variety of technical challenges that were solved by an original design: (1) integral skirts, high bending stiffness, and bending loads are required; (2) high temperature composite stiffness and loads are required up to 160 C; (3) integral fin lugs attachments high aerodynamic loading is required on fin lugs; (4) enclosed fore dome; and (5) aft-pinned joint: a large rear opening is required to cast the propellant. Structural testing in ultimate conditions confirmed the soundness of the design. Positive safety margins were demonstrated on both internal pressure and mechanical loads requirements.

  18. Coupled CFD-Thermal Analysis of Erosion Patterns Resulting from Nozzle Wedgeouts on the SRTMV-N2

    Ables, Catherine; Davis, Philip


    The objective of this analysis was to study the effects of the erosion patterns from the introduction of nozzle flaws machined into the nozzle of the SRTMV-N2 (Solid Rocket Test Motor V Nozzle 2). The SRTMV-N2 motor was a single segment static subscale solid rocket motor used to further develop the RSRMV (Redesigned Solid Rocket Motor V Segment). Two flaws or "wedgeouts" were placed in the nozzle inlet parallel to the ply angles of that section to study erosion effects. One wedgeout was placed in the nose cap region and the other placed in the inlet ring on the opposite side of the bondline, separated 180 degrees circumferentially. A coupled CFD (Computational Fluid Analysis)-thermal iterative analytical approach was utilized at the wedgeouts to analyze the erosion profile during the burn time. The iterative CFD thermal approach was applied at five second intervals throughout the motor burn. The coupled fluid thermal boundary conditions were derived from a steady state CFD solution at the beginning of the interval. The derived heat fluxes were then applied along the surface and a transient thermal solution was developed to characterize the material response over the specified interval. Eroded profiles of each of the nozzle's wedgeouts and the original contour were created at each of the specified intervals. The final iteration of the erosion profile showed that both wedgeouts were "washedout," indicating that the erosion profile of the wedgeout had rejoined the original eroded contour, leaving no trace of the wedgeouts post fire. This analytical assessment agreed with post-fire observations made of the SRTMV-N2 wedgeouts, which noted a smooth eroded contour.

  19. Manufacturing Advanced Channel Wall Rocket Liners Project

    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...

  20. Influence of Structural Parameters on the Performance of Vortex Valve Variable-Thrust Solid Rocket Motor

    Wei, Xianggeng; Li, Jiang; He, Guoqiang


    The vortex valve solid variable thrust motor is a new solid motor which can achieve Vehicle system trajectory optimization and motor energy management. Numerical calculation was performed to investigate the influence of vortex chamber diameter, vortex chamber shape, and vortex chamber height of the vortex valve solid variable thrust motor on modulation performance. The test results verified that the calculation results are consistent with laboratory results with a maximum error of 9.5%. The research drew the following major conclusions: the optimal modulation performance was achieved in a cylindrical vortex chamber, increasing the vortex chamber diameter improved the modulation performance of the vortex valve solid variable thrust motor, optimal modulation performance could be achieved when the height of the vortex chamber is half of the vortex chamber outlet diameter, and the hot gas control flow could result in an enhancement of modulation performance. The results can provide the basis for establishing the design method of the vortex valve solid variable thrust motor.

  1. An Evaluation Of Rocket Parameters

    J. N. Beri


    Full Text Available The dependence of conventional parameters of internal ballistics of Solid Propellant Rockets using external burning cruciform charge, on the geometry of charge aad rocket motor is discussed and results applied in a special case.

  2. Navier-Stokes calculation of solid-propellant rocket motor internal flowfields

    Hsieh, Kwang-Chung; Yang, Vigor; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the detailed physical and chemical processes involved in the combustion of homogeneous propellant in a rocket motor. The formulation is based on the time-dependent full Navier-Stokes equations, with special attention devoted to the chemical reactions in both gas and condensed phases. The turbulence closure is achieved using both the Baldwin-Lomax algebraic model and a modified k-epsilon two-equation scheme with a low Reynolds number and near-wall treatment. The effects of variable thermodynamic and transport properties are also included. The system of governing equations are solved using a multi-stage Runge-Kutta shceme with the source terms treated implicitly. Preliminary results clearly demonstrate the presence of various combustion regimes in the vicinity of propellant surface. The effects of propellant combustion on the motor internal flowfields are investigated in detail.

  3. Velocity-coupled flow oscillations in a simulated solid-propellant rocket environment

    Yang, Vigor; Hsieh, Kwang-Chung; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the unsteady flowfields in a simulated rocket-motor environment. The model is based on the time-dependent compressible Navier-Stokes equations with a two-equation turbulence closure scheme. Various important aspects of the coupling between acoustic oscillations and mean flowfields, including flow reversal, modification of transport properties, etc., are addressed. Results indicate that multi-dimensional effects play important roles in determining local flow structures and wave characteristics. In much of the domain, acoustic velocity nodal points are observed in the near-wall region. The classical one-dimensional theory fails to describe several important mechanisms associated with velocity-induced flow instabilities.

  4. Advanced Flow Analysis Tools for Transient Solid Rocket Motor Simulations Project

    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...

  5. Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method

    Hueser, J. E.; Brock, F. J.; Melfi, L. T., Jr.; Bird, G. A.


    A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).

  6. Scramjet Nozzles


    integration et gestion thermique ) 14. ABSTRACT The lecture is given in four parts, each being a step in the process of nozzle design, and within each part...project and applied to the conceptual design of a Mach 3.5 transport aircraft. The result is depicted in figure 4. The central feature of the concept is

  7. Transient Burning Rate Model for Solid Rocket Motor Internal Ballistic Simulations

    David R. Greatrix


    Full Text Available A general numerical model based on the Zeldovich-Novozhilov solid-phase energy conservation result for unsteady solid-propellant burning is presented in this paper. Unlike past models, the integrated temperature distribution in the solid phase is utilized directly for estimating instantaneous burning rate (rather than the thermal gradient at the burning surface. The burning model is general in the sense that the model may be incorporated for various propellant burning-rate mechanisms. Given the availability of pressure-related experimental data in the open literature, varying static pressure is the principal mechanism of interest in this study. The example predicted results presented in this paper are to a substantial extent consistent with the corresponding experimental firing response data.

  8. Feasibility of an advanced thrust termination assembly for a solid propellant rocket motor


    A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.

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

    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.

  10. Hot wire anemometer measurements in the unheated air flow tests of the SRB nozzle-to-case joint

    Ramachandran, N.


    Hot-Wire Anemometer measurements made in the Solid Rocket Booster (SRB) nozzle-to-case joint are discussed. The study was undertaken to glean additional information on the circumferential flow induced in the SRB nozzle joint and the effect of this flow on the insulation bonding flaws. The tests were conducted on a full-scale, 2-D representation of a 65-in long segment of the SRB nozzle joint, with unheated air as the working fluid. Both the flight Mach number and Reynolds number were matched simultaneously and different pressure gradients imposed along the joint face were investigated. Hot-wire anemometers were used to obtain velocity data for different joint gaps and debond configurations. The procedure adopted for hot-wire calibration and use is outlined and the results from the tests summarized.

  11. Nuclear magnetic resonance imaging of solid rocket propellants at 14.1 T.

    Maas, W E; Merwin, L H; Cory, D G


    Proton NMR images of solid propellant materials, consisting of a polybutadiene binder material filled with 82% solid particles, have been obtained at a magnetic field strength of 14.1 T and at a resolution of 8.5 x 8.5 micron. The images are the first of elastomeric materials obtained at a proton frequency of 600 MHz and have the highest spatial resolution yet reported. The images display a high contrast and are rich in information content. They reveal the distribution of individual filler particles in the polymer matrix as well as a thin polymer film of about 10-30 micron which is found to surround some of the larger filler particles.

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

    Cross, Robert


    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. Propellant grain dynamics in aft attach ring of shuttle solid rocket booster

    Verderaime, V.


    An analytical technique for implementing simultaneously the temperature, dynamic strain, real modulus, and frequency properties of solid propellant in an unsymmetrical vibrating ring mode is presented. All dynamic parameters and sources are defined for a free vibrating ring-grain structure with initial displacement and related to a forced vibrating system to determine the change in real modulus. Propellant test data application is discussed. The technique was developed to determine the aft attach ring stiffness of the shuttle booster at lift-off.

  14. Test on launch jet noise of liquid rocket with single nozzle%单喷管液体火箭发射喷流噪声模拟试验研究

    陈劲松; 曾玲芳; 胡小伟; 范虹


    Based on the distinguishing launch technology of trapped launch vehicle,a simpli-fied launch jet noise test system is designed and developed.The test system can be used to simu-late jet noise and jet flow field about liquid rocket launching with single nozzle.Then the series of launch jet noise tests are accomplished.The test results show that the height domain SPL curves of launch jet noise are different from that of free jet noise,while changes of SPL along with time among different test points show the similar tendency,which is caused by the launch pad distur-bing.The frequency domain SPL curves indicate that there are wide frequency characteristics and distinct screams about the launch jet noise,and the screams of launch jet noise usually have har-monic multiple frequencies or monophonic frequency.The launch jet flow field tests accomplished with the launch jet noise tests show that the developing tendencies of time domain SPL curves of launch jet noise are also similar with that of the engine working pressure and the launch jet flow pressure.%针对捆绑式运载火箭发射噪声问题,研制了一种相对简化的单喷管液体火箭发射喷流噪声模拟试验系统,开展了发射喷流噪声模拟试验研究。研究表明:受发射平台结构扰动效应影响,空间高度方向发射喷流噪声变化规律不同于自由喷流噪声变化规律,但不同测点之间噪声声压级随时间变化规律存在相似性;发射喷流噪声频谱存在宽频特性,同时还存在突出倍谐频啸叫特征或突出单基频啸叫特征。发射喷流噪声模拟试验过程中综合了喷流流场研究,研究发现:喷流噪声声压时域变化规律与发动机工作压力、喷流流场压力时域变化规律也存在相似性。

  15. Design, analysis, fabrication and test of the Space Shuttle solid rocket booster motor case

    Kapp, J. R.


    The motor case used in the solid propellant booster for the Space Shuttle is unique in many respects, most of which are indigenous to size and special design requirements. The evolution of the case design from initial requirements to finished product is discussed, with increased emphasis of reuse capability, special design features, fracture mechanics and corrosion control. Case fabrication history and the resulting procedure are briefly reviewed with respect to material development, processing techniques and special problem areas. Case assembly, behavior and performance during the DM-1 static firing are reviewed, with appropriate comments and conclusions.

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

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


    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.

  17. Methane Dual Expander Aerospike Nozzle Rocket Engine


    v Acknowledgments I would like to thank my thesis advisor, Lt. Col Hartsfield, who has put great effort in discussing, teaching , and...42 NPSS Thermochemistry ................................................................................................44 Methane FPT Generation...Next is a more focused description of how NPSS uses thermochemistry and an explanation on the development of the different fluid property tables

  18. The starting transient of solid propellant rocket motors with high internal gas velocities. Ph.D. Thesis

    Peretz, A.; Caveny, L. H.; Kuo, K. K.; Summerfield, M.


    A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

  19. Design of Multi-Propellant Star Grains for Solid Propellant Rockets

    S. Krishnan


    Full Text Available A new approach to solve the geometry-problem of solid propellant star is presented. The basis of the approach is to take the web-thickness (a ballistic as well as a geometrical property as the characteristic length. The nondimensional characteristic parameters representing diameter, length, slenderness-ratio, and ignitor accommodation of the grain are all identified. Many particular cases of star configurations (from the configurations of single propellant to those of four different propellants can be analysed through the identified characteristic parameters. A better way of representing the single-propellant-star-performance in a design graph is explained. Two types of dual propellant grains are analysed in detail. The first type is characterised by its two distinct stages of burning (initially by single propellant burning and then by dual propellant burning; the second type has the dual propellant burning throughout. Suitability of the identified characteristic parameters to an optimisation study is demonstrated through examples.

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

    Terry, Brandon C., E-mail: [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)


    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.

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

    Young, Philip R.


    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.

  2. Application Analysis of Multi-functional Structure Design for Solid Rocket%多功能结构在固体火箭上的应用分析

    曹莉; 介党阳; 熊楚杨; 徐嘉


    It was very limited to realize light-weight design for solid rocket adopting traditional mechanical structure optimization method. Due to the design philosophy of mechanical, electrical, thermal integration, multi-functional structure design broke a new path for solid rocket performance promotion. The method and thought of multi-functional structure design were introduced systematically. Combining with the development of a new onboard recording equipment prototype, the application prospect of multi-functional structure was discussed deeply on solid rocket area.%为实现固体火箭轻质化设计,采用传统机械结构优化的方法能力十分有限。多功能结构优化设计方法,由于采用机、电、热一体化的系统工程级设计思路,为提升火箭整体性能开辟了一个全新的方向。介绍多功能结构设计方法及思路,结合一种新型箭载记录设备原型样机的研制,对多功能结构在固体火箭领域应用的前景进行深入探讨。

  3. Gas only nozzle

    Bechtel, William Theodore; Fitts, David Orus; DeLeonardo, Guy Wayne


    A diffusion flame nozzle gas tip is provided to convert a dual fuel nozzle to a gas only nozzle. The nozzle tip diverts compressor discharge air from the passage feeding the diffusion nozzle air swirl vanes to a region vacated by removal of the dual fuel components, so that the diverted compressor discharge air can flow to and through effusion holes in the end cap plate of the nozzle tip. In a preferred embodiment, the nozzle gas tip defines a cavity for receiving the compressor discharge air from a peripheral passage of the nozzle for flow through the effusion openings defined in the end cap plate.

  4. Thrust Control Features for Aerodynamic Throat for Solid Rocket Motor%固体火箭发动机气动喉部的推力调控特性

    谢侃; 李博; 郭常超; 王宁飞


    为了研究固体火箭发动机气动喉部推力调节的一般规律,利用氮气作为介质对气动喉部喷管进行了冷流实验研究。研究了该种喷管的扼流性能,二次流嘴的面积、个数对其扼流性能的影响以及空腔容积与喷管压强调节时间的关系。掌握了气动喉部喷管的有效喉部面积随流量比变化的一般规律。结果表明,二次流与主流流量比越大,气动喉部面积越小。小的面积比具有更高的扼流性能,而当流量比大于0.4时,面积比对扼流性能无明显影响。空腔体积越小压强调节时间越短。%Adjust:In order to study the general rules of the thrust control features for aerodynamic throat for solid rocket motor, a cold-flow test research was performed on the aerodynamic nozzle throat using the nitrogen as working gas. Firstly, the choke performance of the aerodynamic throat was studied. Then, the performance of throat modification was studied considering the influence of area and number of secondary flow injectors. Final⁃ly,the relation about the volume of cavity with adjustment time of nozzle was also studied. The general law of ef⁃fective throat area along with the flow rate change was obtained through study. The results show that the larger the flow ratio of the secondary flow and main flow is, the smaller the pneumatic throat area is. The smaller area ratio has higher choke performance,and the influence of area is not obvious when the flow rate is greater than 0.4. Be⁃sides,the smaller the cavity volume is,the shorter the accommodation time of pressure is.

  5. A numerical simulation on the infrared radiation of hot exhausting nozzles with a coupled flow and heat transfer model


    A coupled model among flow field,solid temperature,species concentration and gas radiation,which was based on statistical narrow-band correlated-k model,was employed to predict the infrared radiations from hot exhausting nozzles. The parameters of narrow-band model were deduced from HITEMP line-by-line database. Several methods to increase computational efficiency and to save computational resources were employed,thus all the complicated computations could be operated on a personal computer. The predictions for three cases have been conducted to validate the accuracy of the methods mentioned above,including the temperature distribution of a water-cooling nozzle in rocket engines,the carbon dioxide absorptivity at the wavelength of 4.3 micron and the infrared radiation of a cylindrical furnace. Finally,the aerothermodynamic and infrared characteristics of two nozzles were predicted. It was shown that the infrared radiation intensity of chevron ejecting nozzle were obviously smaller than that of common axisymmetric convergent-divergent nozzle.

  6. Fully Coupled Aero-Thermochemical-Elastic Simulations of an Eroding Graphite Nozzle

    Blades, E. L.; Reveles, N. D.; Nucci, M.; Maclean, M.


    A multiphysics simulation capability has been developed that incorporates mutual interactions between aerodynamics, structural response from aero/thermal loading, ablation/pyrolysis, heating, and surface-to-surface radiation to perform high-fidelity, fully coupled aerothermoelastic ablation simulations, which to date had been unattainable. The multiphysics framework couples CHAR (a 3-D implicit charring ablator solver), Loci/CHEM (a computational fluid dynamics solver for high-speed chemically reacting flows), and Abaqus (a nonlinear structural dynamics solver) to create a fully coupled aerothermoelastic charring ablative solver. The solvers are tightly coupled in a fully integrated fashion to resolve the effects of the ablation pyrolysis and charring process and chemistry products upon the flow field, the changes in surface geometry due to recession upon the flow field, and thermal-structural analysis of the body from the induced aerodynamic heating from the flow field. The multiphysics framework was successfully demonstrated on a solid rocket motor graphite nozzle erosion application. Comparisons were made with available experimental data that measured the throat erosion during the motor firing. The erosion data is well characterized, as the test rig was 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 initially undergoes a nozzle contraction due to thermal expansion before ablation effects are able to widen the throat. A series of parameters studies were conducted using the coupled simulation capability to determine the sensitivity of the nozzle erosion to different parameters. The parameter studies included the shape of the nozzle throat (flat versus rounded), the material properties, the effect of the choice of turbulence model, and the inclusion or exclusion of the mechanical thermal expansion. Overall, the predicted results match

  7. Implementation of microwave transmissions for rocket exhaust plume diagnostics

    Coutu, Nicholas George

    Rocket-launched vehicles produce a trail of exhaust that contains ions, free electrons, and soot. The exhaust plume increases the effective conductor length of the rocket. A conductor in the presence of an electric field (e.g. near the electric charge stored within a cloud) can channel an electric discharge. The electrical conductivity of the exhaust plume is related to its concentration of free electrons. The risk of a lightning strike in-flight is a function of both the conductivity of the body and its effective length. This paper presents an approach that relates the electron number density of the exhaust plume to its propagation constant. Estimated values of the collision frequency and electron number density generated from a numerical simulation of a rocket plume are used to guide the design of the experimental apparatus. Test par meters are identified for the apparatus designed to transmit a signal sweep form 4 GHz to 7 GHz through the exhaust plume of a J-class solid rocket motor. Measurements of the scattering parameters imply that the transmission does not penetrate the plume, but instead diffracts around it. The electron density 20 cm downstream from the nozzle exit is estimated to be between 2.7x1014 m--3 and 5.6x10 15 m--3.

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

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


    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

  9. Solid Propellant Test Article (SPTA) Test Stand


    This photograph shows the Solid Propellant Test Article (SPTA) test stand with the Modified Nasa Motor (M-NASA) test article at the Marshall Space Flight Center (MSFC). The SPTA test stand, 12-feet wide by 12-feet long by 24-feet high, was built in 1989 to provide comparative performance data on nozzle and case insulation material and to verify thermostructural analysis models. A modified NASA 48-inch solid motor (M-NASA motor) with a 12-foot blast tube and 10-inch throat makes up the SPTA. The M-NASA motor is being used to evaluate solid rocket motor internal non-asbestos insulation materials, nozzle designs, materials, and new inspection techniques. New internal motor case instrumentation techniques are also being evaluated.

  10. 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.

    Schramm, Harry F.; Sullivan, Kenneth W.


    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.

  11. Enantioselective synthesis of aziridines using asymmetric transfer hydrogenation as a precursor for chiral derivatives used as bonding agent for rocket solid propellants

    Aparecida M. Kawamoto


    Full Text Available A rapid, expedient and enantioselective method for the synthesis of beta-hydroxy amines and monosubstituted aziridines in up to 99% e.e., via asymmetric transfer hydrogenation of a-amino ketones and cyclisation through treatment with tosyl chloride and base, is described. (1R,2R-N-(para-toluenesulfonyl-1,2-ethylenediamine with formic acid has been utilised as a ligand for the Ruthenium (II catalysed enantioselective transfer hydrogenation of the ketones.The chiral 2-methyl aziridine, which is a potentially more efficient bonding agent for Rocket Solid Propellant has been successfully achieved.

  12. Rocket Flight.

    Van Evera, Bill; Sterling, Donna R.


    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)

  13. Unique nuclear thermal rocket engine

    Culver, D.W. (Aerojet Propulsion Division, P.O. Box 13222, Sacramento, California 95813-6000 (United States)); Rochow, R. (Babcock Wilcox Space Nuclear Systems, P.O. Box 11165, Lynchburg, Virginia 24506-1165 (United States))


    Earlier this year Aerojet Propulsion Division (APD) introduced a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars. 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 (E-D) 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)Reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2)Eliminate need for a new, uncooled nozzle throat material suitable for long life application; (3)Practical provision for reactor power control; and (4)Use near term, long life turbopumps.

  14. Unique nuclear thermal rocket engine

    Culver, Donald W.; Rochow, Richard


    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.

  15. Computational Fluid Dynamics Simulation of Dual Bell Nozzle Film Cooling

    Braman, Kalen; Garcia, Christian; Ruf, Joseph; Bui, Trong


    Marshall Space Flight Center (MSFC) and Armstrong Flight Research Center (AFRC) are working together to advance the technology readiness level (TRL) of the dual bell nozzle concept. Dual bell nozzles are a form of altitude compensating nozzle that consists of two connecting bell contours. At low altitude the nozzle flows fully in the first, relatively lower area ratio, nozzle. The nozzle flow separates from the wall at the inflection point which joins the two bell contours. This relatively low expansion results in higher nozzle efficiency during the low altitude portion of the launch. As ambient pressure decreases with increasing altitude, the nozzle flow will expand to fill the relatively large area ratio second nozzle. The larger area ratio of the second bell enables higher Isp during the high altitude and vacuum portions of the launch. Despite a long history of theoretical consideration and promise towards improving rocket performance, dual bell nozzles have yet to be developed for practical use and have seen only limited testing. One barrier to use of dual bell nozzles is the lack of control over the nozzle flow transition from the first bell to the second bell during operation. A method that this team is pursuing to enhance the controllability of the nozzle flow transition is manipulation of the film coolant that is injected near the inflection between the two bell contours. Computational fluid dynamics (CFD) analysis is being run to assess the degree of control over nozzle flow transition generated via manipulation of the film injection. A cold flow dual bell nozzle, without film coolant, was tested over a range of simulated altitudes in 2004 in MSFC's nozzle test facility. Both NASA centers have performed a series of simulations of that dual bell to validate their computational models. Those CFD results are compared to the experimental results within this paper. MSFC then proceeded to add film injection to the CFD grid of the dual bell nozzle. A series of

  16. Technology Method Design of Assembly and Testing for Solid Propellant Rocket Engine of Aviation Seat%航空座椅固体火箭发动机装配及检测工艺技术设计



    本文对航空座椅某型固体火箭发动机部装、总装及检测、试验、包装技术难点等进行了工艺分析;介绍了固体火箭发动机装配全过程工艺流程、检测、试验方法及注意事项等,对于同类及新型火箭发动机的装配制造过程具有良好的借鉴、推广应用意义。%Aiming at the difficulty of solid propellant rocket engine of aviation seat to assembly, testing and packaging technology, the assembly, testing process and method for solid propellant rocket engine were introduced. It can be regarded as reference with application for solid propellant rocket engine assembly process.

  17. Cold spray nozzle design

    Haynes, Jeffrey D.; Sanders, Stuart A.


    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.

  18. Design and Analysis of Metal-to-Composite Nozzle Extension Joints Project

    National Aeronautics and Space Administration — As the operational demands of liquid rocket engines increases, so too does the need for improved design and manufacturing methods for metal-to-composite nozzle...

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

    Walton, James T.


    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.

  20. Application of Dry Coupling Ultrasonic Detecting Technology for the Rocket Engine Nozzle%干耦合超声检测技术在某火箭发动机喷管在役检测中的应用

    穆洪彬; 吴朝军; 吴晨; 李剑


    应用干耦合超声检测方法对某火箭发动机喷管进行了在役检测试验,验证了干耦合超声检测技术对该火箭发动机喷管的在役检测能力.结果表明,干耦合超声检测方法能够有效地检测出该型号发动机喷管中φ 20 mm以上的脱粘缺陷,满足在役检测要求.自动与手动,两种检测方法的检测能力相当,但在操作方式和工作效率等方面各有优劣.%In order to study the feasibility of the in-service dry coupling ultrasonic test technology for rocket jet,the principle of the dry coupling ultrasonic testing was analyzed,and both the manual and automatic testing methods were applied to test the rocket jet specimen and the rocket jet parts in service.The results show that the debond defect,whose normal size is larger than 20mm,can be found using the two kinds of the dry coupling ultrasonic testing methods.It satisfies the quality request of the rocket jet being tested in service.It is the same to the manual testing and automatic testing in the ability of finding defect.And there is difference between the two testing methods in the operation and efficiency.The study is useful for testing rocket jet in service.

  1. Simulation of Cold Flow in a Truncated Ideal Nozzle with Film Cooling

    Braman, K. E.; Ruf, J. H.


    Flow transients during rocket start-up and shut-down can lead to significant side loads on rocket nozzles. The capability to estimate these side loads computationally can streamline the nozzle design process. Towards this goal, the flow in a truncated ideal contour (TIC) nozzle has been simulated using RANS and URANS for a range of nozzle pressure ratios (NPRs) aimed to match a series of cold flow experiments performed at the NASA MSFC Nozzle Test Facility. These simulations were performed with varying turbulence model choices and for four approximations of the supersonic film injection geometry, each of which was created with a different simplification of the test article geometry. The results show that although a reasonable match to experiment can be obtained with varying levels of geometric fidelity, the modeling choices made do not fully represent the physics of flow separation in a TIC nozzle with film cooling.

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

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


    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

  3. Spiral cooled fuel nozzle

    Fox, Timothy; Schilp, Reinhard


    A fuel nozzle for delivery of fuel to a gas turbine engine. The fuel nozzle includes an outer nozzle wall and a center body located centrally within the nozzle wall. A gap is defined between an inner wall surface of the nozzle wall and an outer body surface of the center body for providing fuel flow in a longitudinal direction from an inlet end to an outlet end of the fuel nozzle. A turbulating feature is defined on at least one of the central body and the inner wall for causing at least a portion of the fuel flow in the gap to flow transverse to the longitudinal direction. The gap is effective to provide a substantially uniform temperature distribution along the nozzle wall in the circumferential direction.

  4. 动能拦截器姿控固体小火箭点火算法设计%Ignition algorithm for attitude control solid-propellant nozzles in kinetic interceptor

    李广华; 张洪波; 汤国建


    使用固体姿控小火箭是实现动能拦截器快响应和高精度姿态控制的最佳方案之一。针对一种新型动能拦截器姿控小火箭布局,提出了点火组合混合搜索算法。描述了动能拦截器姿控小火箭的配置方案,分析了弹体自旋需求。设计了一种结合目标排序法和区间搜索法的点火组合混合搜索算法:当可用小火箭个数较少时,采用目标排序法;当可用小火箭个数较多时,采用区间搜索法。指令力矩近似仿真结果及姿态控制数值仿真结果表明:该算法能够有效地近似指令力矩,实现快速高精度的姿态跟踪。%Using solid-propellant nozzles is one of the best schemes for kinetic interceptor to realize the fast response and high precision of attitude control.A mixed searching algorithm for ignition combination was presented for a novel attitude control solid-propellant nozzle in kinetic interceptor.Firstly,the configuration of solid-propellant nozzles was described and spin requirements of the kinetic interceptor were analyzed.Then the mixed searching algorithm was designed by a combination of sorting method and interval searching method.Sorting method is selected when the number of available nozzles is small and interval searching method is chosen on the contrary.Results of instruction torque approximation simulation and attitude control numerical simulation suggest that the algorithm can track the instruction torque effectively and achieve attitude tracking quickly and with a high precision.

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

    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


    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.

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

    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

  7. The sky is falling: chemical characterization and corrosion evaluation of deposition produced during the static testing of solid rocket motors.

    Doucette, William J; McNeill, Laurie S; Mendenhall, Scout; Hancock, Paul V; Wells, Jason E; Thackeray, Kevin J; Gosen, David P


    Static tests of horizontally restrained rocket motors at the ATK facility in Promontory UT, USA result in the deposition of entrained soil and fuel combustion products, referred to as Test Fire Soil (TFS), over areas as large as 30-50 mile (80-130 km) and at distances up to 10-12 miles (16-20 km) from the test site. Chloride is the main combustion product generated from the ammonium perchlorate-aluminum based composite propellant. Deposition sampling/characterization and a 6-month field corrosivity study using mild steel coupons were conducted in conjunction with the February 25th 2010 FSM-17 static test. The TFS deposition rates at the three study sites ranged from 1 to 5 g/min/m. TFS contained significantly more chloride than the surface soil collected from the test site. The TFS collected during two subsequent tests had similarly elevated chloride, suggesting that the results obtained in this study are applicable to other tests assuming that the rocket fuel composition remains similar. The field-deployed coupons exposed to the TFS had higher corrosion rates (3.6-5.0 mpy) than paired non-exposed coupons (1.6-1.8 mpy). Corrosion rates for all coupons decreased over time, but coupons exposed to the TFS always had a higher rate than the non-exposed. Differences in corrosion rates between the three study sites were also observed, with sites receiving more TFS deposition having higher corrosion rates.

  8. 固体火箭发动机预固化技术及其应用%Pre-cure Technique and Its Applications for Solid Rocket Motors

    苏昌银; 张爱科


    Based on interface properties of HTPB propellant, the cross linking level is regulated through the temperature and time of cure reaction. The chemical reaction takes place gradually in remainder functional groups of the system to form chemical bonds and hydrogen bond, so as to improve the mechanical properties of the resultant. Pre-cure technique and bonding model are described in the paper. They can be used in propellant-liner bonding, propellant loading and integral repairing of the propellant grain of the solid rocket motor (SRM). These test results have been qualified by the successful static firing tests, flight tests of motors and storage tests of specimens for ten-years. The performances of the motor meet the design requirements with good reproducibilities.

  9. Solid propellants for rockets. Rocket suishin yaku

    Kubota, N. (Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Physical and chemical ProPerties and combustion characteristics of propellants differ according to the combination of oxidizers and fuel components. Composite smoke propellant, having crystalline ammonium perchlorate as an oxidizer and hydrocarbon Polymer as a fuel, has higher specific impulse and improved mechanical properties compared to smokeless double base propellant consisting of nitroglycerin and nirocellulose. Double base propellants with low specific impulse are combined with nitramines( RDX or HMX ) to make composite modified double based( CMDB ) propellants, as a result the smokeless property of double base propellant is preserved and the combustion efficiency is increased. With the combination of oxidizing agents and fuels, formation of various high functional propellants has been possible and energetic azide polymers have provided possibilities for fuels of propellants. 3 refs., 6 figs., 3 tabs.

  10. Hybrid Rocket Technology

    Sankaran Venugopal


    Full Text Available With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nontoxic, and not hazardous to store and transport. Solid fuel grains are not highly susceptible to cracks, imperfections, and environmental temperature and are therefore safer to manufacture, store, transport, and use for launch. The status of development based on the experience of the last few decades indicating the maturity of the hybrid rocket technology is given in brief.Defence Science Journal, 2011, 61(3, pp.193-200, DOI:

  11. Analysis of Viscous Heating in a Micro-Rocket Flow and Performance

    José A. Morí(n)igo; José Hermida Quesada


    Micro-rockets for propulsion of small spacecrafts exhibit significant differences with regard to their macroscale counterparts, mainly caused by the role of the viscous dissipation and heat transfer processes in the micron-sized scale. The goal of this work is to simulate the transient operation of a micro-rocket to investigate the effects of viscous heating on the flow and performance for four configurations of the expanding gas and wafer material. The modelling follows a multiphysics approach that solves the fluid and solid regions fully coupled. A continuum-based description that incorporates the effects of gas rarefaction through the micro-nozzle, viscous dissipation and heat transfer at the solid-gas interface is presented. Non-equilibrium is addressed with the implementation of a 2nd-order slip-model for the velocity and temperature at the walls. The results stress that solid-fluid coupling exerts a strong influence on the flowfield and performance as well as the effect of the wafer during the first instants of the transient in micro-rockets made of low and high thermal conductivity materials.

  12. Rocket noise - A review

    McInerny, S. A.


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

  13. Gas only nozzle fuel tip

    Bechtel, William Theodore (Scotia, NY); Fitts, David Orus (Ballston Spa, NY); DeLeonardo, Guy Wayne (Glenville, NY)


    A diffusion flame nozzle gas tip is provided to convert a dual fuel nozzle to a gas only nozzle. The nozzle tip diverts compressor discharge air from the passage feeding the diffusion nozzle air swirl vanes to a region vacated by removal of the dual fuel components, so that the diverted compressor discharge air can flow to and through effusion holes in the end cap plate of the nozzle tip. In a preferred embodiment, the nozzle gas tip defines a cavity for receiving the compressor discharge air from a peripheral passage of the nozzle for flow through the effusion openings defined in the end cap plate.

  14. Firefighter Nozzle Reaction

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


    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...

  15. Thermal Hydraulics Design and Analysis Methodology for a Solid-Core Nuclear Thermal Rocket Engine Thrust Chamber

    Wang, Ten-See; Canabal, Francisco; Chen, Yen-Sen; Cheng, Gary; Ito, Yasushi


    Nuclear thermal propulsion is a leading candidate for in-space propulsion for human Mars missions. This chapter describes a thermal hydraulics design and analysis methodology developed at the NASA Marshall Space Flight Center, in support of the nuclear thermal propulsion development effort. The objective of this campaign is to bridge the design methods in the Rover/NERVA era, with a modern computational fluid dynamics and heat transfer methodology, to predict thermal, fluid, and hydrogen environments of a hypothetical solid-core, nuclear thermal engine the Small Engine, designed in the 1960s. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics and heat transfer platform, while formulations of flow and heat transfer through porous and solid media were implemented to describe those of hydrogen flow channels inside the solid24 core. Design analyses of a single flow element and the entire solid-core thrust chamber of the Small Engine were performed and the results are presented herein


    Walton, J. T.


    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.

  17. Analysis of velocity-coupled response function data from the dual rotating valve. [combustion stability of solid rocket propellants

    Brown, R. S.; Waugh, R. C.


    The results of a re-evaluation of the propellant combustion data obtained using the dual valve approach for measuring velocity-coupling characteristics of solid propellants are presented. Data analysis and testing procedures are described. The velocity response is compared to pressure-coupled response data within the context of thermal wave response theory. This comparison shows important inconsistencies which cast doubt on inferring the velocity response from pressure-coupled response functions.

  18. Transition nozzle combustion system

    Kim, Won-Wook; McMahan, Kevin Weston; Maldonado, Jaime Javier


    The present application provides a combustion system for use with a cooling flow. The combustion system may include a head end, an aft end, a transition nozzle extending from the head end to the aft end, and an impingement sleeve surrounding the transition nozzle. The impingement sleeve may define a first cavity in communication with the head end for a first portion of the cooling flow and a second cavity in communication with the aft end for a second portion of the cooling flow. The transition nozzle may include a number of cooling holes thereon in communication with the second portion of the cooling flow.

  19. Wear characterization of abrasive waterjet nozzles and nozzle materials

    Nanduri, Madhusarathi

    Parameters that influence nozzle wear in the abrasive water jet (AWJ) environment were identified and classified into nozzle geometric, AWJ system, and nozzle material categories. Regular and accelerated wear test procedures were developed to study nozzle wear under actual and simulated conditions, respectively. Long term tests, using garnet abrasive, were conducted to validate the accelerated test procedure. In addition to exit diameter growth, two new measures of wear, nozzle weight loss and nozzle bore profiles were shown to be invaluable in characterizing and explaining the phenomena of nozzle wear. By conducting nozzle wear tests, the effects of nozzle geometric, and AWJ system parameters on nozzle wear were systematically investigated. An empirical model was developed for nozzle weight loss rate. To understand the response of nozzle materials under varying AWJ system conditions, erosion tests were conducted on samples of typical nozzle materials. The effect of factors such as jet impingement angle, abrasive type, abrasive size, abrasive flow rate, water pressure, traverse speed, and target material was evaluated. Scanning electron microscopy was performed on eroded samples as well as worn nozzles to understand the wear mechanisms. The dominant wear mechanism observed was grain pullout. Erosion models were reviewed and along the lines of classical erosion theories a semi-empirical model, suitable for erosion of nozzle materials under AWJ impact, was developed. The erosion data correlated very well with the developed model. Finally, the cutting efficiency of AWJ nozzles was investigated in conjunction with nozzle wear. The cutting efficiency of a nozzle deteriorates as it wears. There is a direct correlation between nozzle wear and cutting efficiency. The operating conditions that produce the most efficient jets also cause the most wear in the nozzle.

  20. MHD thrust vectoring of a rocket engine

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


    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.

  1. Fundamentals of aircraft and rocket propulsion

    El-Sayed, Ahmed F


    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...

  2. The Analysis for Grain Structural Integrity of a Certain Solid Rocket Motor%某发动机装药结构完整性分析

    张亮; 邢国强


    Based on three-dimension viscoelastic finite element method, by MSC/NASTRAN software system, the grain structural integrity of the motor under internal pressure load, under thermal load and under the combined action of the two load is analyzed and evaluatedrespectively. The results show that the grain structural integrity of the solid rocket motor meets the design requirement satisfaction.%基于三维粘弹性有限元模型,应用MSC/NASTRAN软件对某发动机分别在固化降温、燃气内压载荷条件下的装药结构完整性进行分析,并对该发动机在固化降温、燃气内压两种载荷联合作用下的装药结构完整性进行评估。结果表明,该发动机的装药结构完整性满足要求。

  3. Adsorption and chemical reaction of gaseous mixtures of hydrogen chloride and water on aluminum oxide and application to solid-propellant rocket exhaust clouds

    Cofer, W. R., III; Pellett, G. L.


    Hydrogen chloride (HCl) and aluminum oxide (Al2O3) are major exhaust products of solid rocket motors (SRM). Samples of calcination-produced alumina were exposed to continuously flowing mixtures of gaseous HCl/H2O in nitrogen. Transient sorption rates, as well as maximum sorptive capacities, were found to be largely controlled by specific surface area for samples of alpha, theta, and gamma alumina. Sorption rates for small samples were characterized linearly with an empirical relationship that accounted for specific area and logarithmic time. Chemisorption occurred on all aluminas studied and appeared to form from the sorption of about a 2/5 HCl-to-H2O mole ratio. The chemisorbed phase was predominantly water soluble, yielding chloride/aluminum III ion mole ratios of about 3.3/1 suggestive of dissolved surface chlorides and/or oxychlorides. Isopiestic experiments in hydrochloric acid indicated that dissolution of alumina led to an increase in water-vapor pressure. Dissolution in aqueous SRM acid aerosol droplets, therefore, might be expected to promote evaporation.

  4. Vacuum plasma spray applications on liquid fuel rocket engines

    Mckechnie, T. N.; Zimmerman, F. R.; Bryant, M. A.


    The vacuum plasma spray process (VPS) has been developed by NASA and Rocketdyne for a variety of applications on liquid fuel rocket engines, including the Space Shuttle Main Engine. These applications encompass thermal barrier coatings which are thermal shock resistant for turbopump blades and nozzles; bond coatings for cryogenic titanium components; wear resistant coatings and materials; high conductivity copper, NaRloy-Z, combustion chamber liners, and structural nickel base material, Inconel 718, for nozzle and combustion chamber support jackets.

  5. Metal atomization spray nozzle

    Huxford, Theodore J.


    A spray nozzle for a magnetohydrodynamic atomization apparatus has a feed passage for molten metal and a pair of spray electrodes mounted in the feed passage. The electrodes, diverging surfaces which define a nozzle throat and diverge at an acute angle from the throat. Current passes through molten metal when fed through the throat which creates the Lorentz force necessary to provide atomization of the molten metal.

  6. Transonic swirling nozzle flow

    Keith, Theo G., Jr.; Pawlas, Gary E.


    A numerical model of viscous transonic swirling flow in axisymmetric nozzles is developed. MacCormack's implicit Gauss-Seidel method is applied to the thin-layer Navier-Stokes equations in transformed coordinates. Numerical results are compared with experimental data to validate the method. The effect of swirl and viscosity on nozzle performance are demonstrated by examining wall pressures, Mach contours, and integral parameters.

  7. Hybrid rocket propulsion systems for outer planet exploration missions

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


    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.

  8. 气-固-液混凝土喷射机喷头结构的力学性能分析%Analysis of the Mechanics Characteristic in Gas-Solid-Liquid Concrete Jet Nozzle

    王金明; 谢旭时; 陈海; 叶谋平; 郭金基


    阐述气-固-液混凝土喷射机的结构,计算气-固-液流体的密度和质量流率,研究混合流体的动力学方程,讨论喷头结构对喷射性能的影响,最后介绍混凝土喷射机在建筑基坑边坡支护工程及喷锚网等的应用。%The structures of gas-solid-liquid concrete jet are described in this paper. The density and mass velocity of gas-solid-liquid fluid are calculated here. The equations of the mixed fluid dynamics are studied. The influences of the structures of nozzle on the jet characteristic are discussed. At last , the applications of the concrete jet in the brace engineering of constructional trench slope and jet anchored mesh etc. are introduced.

  9. 高精度小型固体火箭发动机184性能检测系统研究%Research on the Performance Detecting System of High-precision Small-scale Solid Rocket Motor

    冯喜平; 董韬; 李进贤; 曹琪


    针对固体火箭发动机研制和生产中的性能检测需求,基于柔性试验架建立试验平台,采用虚拟仪器技术搭建测控平台,使用LabvieW7.1开发一套包含参数标定、数据测量、数据处理等模块的固体火箭发动机性能检测试验测控软件,构建了固体火箭发动机性能检测系统。通过对标准发动机进行测试,结果表明:该系统实现了发动机参数的现场方便标定、发动机数据的高速采集和实验数据的快速处理,测量精度达到0.3%的工程要求,并同步监测了整个发动机的工作过程,满足了发动机性能检测的高精度要求。%According to the performance detecting requirements in development and manufacture of solid rocket motor, a test platform has been set up based on flexible test stand, in which the measurement and control platform has been built using virtual instrument technology ; a set of solid rocket motor ground test performance detecting software including parameter calibration model, data measure model and data processing model have been developed based on the LabViewT. 1 ; and the solid rocket motor performance detecting system was built. The application results through testing normative solid rocket motor show that: this system could actualize parameter calibration expediently, experimental data collection and processing rapidly, even reach the measurement precision by 0.3% required in projects and could monitor the testing process of solid rocket motor simultaneously, which satisfied high-precision requirements of motor performance detection.

  10. 固体火箭发动机环缝式气动喉部研究%Investigation of ring aerodynamic throat for solid rocket motor

    谢侃; 刘宇; 王一白


    对固体火箭发动机气体二次流控制的环缝式气动喉部方案进行了数值模拟.研究了二次流不同喷射位置、角度、流率及喷嘴几何参数对气动喉部调节性能的影响规律.计算得到了气动喉部的流场特征,即气动喉部的声速线起点在二次流喷口的下游,并得到了气动喉部特征存在的喷注范围.结果还表明使二次流的喷入位置越靠近喉部、增大二次流流量或减小喷射角度都能明显增加气动喉部调节性能.%A ring aerodynamic throat concept of solid rocket motor controlled by secondary injection was simulated numerically. The influence and rules of secondary flow injection positions, injection angle, secondary flow rate and injector geometry parameters on the performance of aerodynamic throat were studied. The flow field character of aerodynamic throat was attained by calculation, namely: the starting point of sonic line from the downstream of secondary injection outlet. And the scope of injection positions for the existence of aerodynamic throat character was also attained. The results show that, the performance of aerodynamic throat can be obviously increased by making secondary flow injector close to the throat, increasing secondary flow rate or decreasing secondary flow injection angle.

  11. 某发动机石墨喉衬的裂纹成因分析%Graphite throat insert cracks analysis of a solid rocket motor

    熊波; 白彦军; 唐敏


    确保石墨喉衬在工作期间的结构完整性是某级间分离固体发动机设计的重要任务,初步设计方案经试车考核后,部分石墨喉衬出现了裂纹。针对该发动机喷管进行了喉衬热结构仿真分析,获得了不同危险点的应力大小和状态,并探讨了该喉衬裂纹的成因。结果表明,喷管仿真分析结果与实际试车结果吻合较好,点火初期石墨喉衬在大梯度的温度和压力冲击下,内部易产生较大的热应力,若石墨材料轴向拉伸强度不足,将可能导致裂纹出现。%A solid rocket motor is used for the separation of new launch vehicle stage. One of the important tasks is to ensure the integrality of graphite throat insert in the firing time. After experiment of the initial design, crack was observed in some of the throat insert was proposed. A simulation analysis method for throat insert thermal structure was proposed. The reason for the crack was explored. The results show that, in the initial time of firing, interior of graphite throat may produce great thermal stress under large gradient temperature and pressure, and the insufficiency of tensile strength in axial direction gives birth to the crack.

  12. 固体火箭发动机药柱加压固化仿真%Simulation on pressure cure of solid rocket motor grain

    宗路航; 杜聪; 卢山; 姚东; 郜婕; 沙宝林


    For those case-bonded casting solid rocket motors ( SRMs) with large outside/inside radius ratio of the grain, high thermal strain will be generated in the propellant grain subjected to thermal loading, which significantly limit the performance of SRM. Pressure cure is an effective method to reduce the thermal strain in the grain. In this paper, the theory of pressure cure was analyzed and the relationship between the desired pressure and the parameters of the SRM was deduced. Then, a finite element method ( FEM) of pressure cure named two step method was proposed. Theoretical calculation and FEM simulation were carried out on a tube motor with four different material cases. The recommended pressure of pressure cure of four cases were given out.%对于药柱外/内径比(m数)很大的贴壁浇注式固体火箭发动机,在固化降温后,推进剂药柱内会产生显著的热应变,这严重限制了发动机的进一步高性能化. 加压固化是一种降低推进剂药柱内热应变的有效方法. 文中分析了加压固化的原理,推导出了加压固化所需压强与发动机参数之间的关系式,提出了一种两步分析法的加压固化有限元分析方法. 针对4种不同壳体材料的圆管发动机,进行了加压固化理论计算与有限元仿真分析,给出了4种壳体加压固化时的推荐压强.

  13. Rocket Tablet,


    is a vast and desolate world, this is a strip of mir- aculous land! How many struggling dramas full of power and * grandeur were cheered, resisted and...rocket officers and men, a group enormous and powerful , marched into this land soaked with the fresh blood of our ancestors. This place is about to...and tough pestering said he wanted an American aircraft ob- tained on the battlefield to transport goods from Lanzhou, Xian, Beijing, Guangzhou and

  14. 3D Reacting Flow Analysis of LANTR Nozzles

    Stewart, Mark E. M.; Krivanek, Thomas M.; Hemminger, Joseph A.; Bulman, M. J.


    This paper presents performance predictions for LANTR nozzles and the system implications for their use in a manned Mars mission. The LANTR concept is rocket thrust augmentation by injecting Oxygen into the nozzle to combust the Hydrogen exhaust of a Nuclear Thermal Rocket. The performance predictions are based on three-dimensional reacting flow simulations using VULCAN. These simulations explore a range of O2/H2 mixture ratios, injector configurations, and concepts. These performance predictions are used for a trade analysis within a system study for a manned Mars mission. Results indicate that the greatest benefit of LANTR will occur with In-Situ Resource Utilization (ISRU). However, Hydrogen propellant volume reductions may allow greater margins for fitting tanks within the launch vehicle where packaging issues occur.

  15. Numerical Analysis on the Thermal Safety of Solid Rocket Motor Propellant%固体发动机装药热安全性数值分析

    刘文一; 焦冀光


    Objective To investigate the safety of solid rocket motor(SRM)when it was cook-off. Methods Finite element model of solid SRM was established, and the temperature distribution and the explosion delay time of propellant in fast cook-off mode and slow cook- off mode were computed. Results Propellant reached its critical temperature (352 ℃) after 47 h slow cook-off, while it reached its critical temperature (355 ℃) after 697 s fast cook-off. Conclusion It was proven that the thermal diffusivity in fast cook-off mode was greater than that in slow cook-off mode, while the temperature gradient had an opposite trend. The reaction position of propellant was different in the two different working modes when it reached critical temperature, and the thermal storage capacity of propellant was dependent on its thickness.%目的:研究固体火箭发动机遭受火烤时的安全性。方法建立发动机有限元模型,计算推进剂在慢速烤燃和快速烤燃工况下的温度分布和爆炸延迟时间。结果推进剂慢烤47 h后达到临界温度,其值为352℃;快烤推进剂加热697 s后达到临界温度,临界温度为355℃。结论推进剂在快速烤燃模式下的热扩散速率大于慢速烤燃工况下,但是温度梯度则相反。两种工况下推进剂达到临界温度后开始反应的位置不同,推进剂厚度决定了其储热能力。

  16. The four INTA-300 rocket prototypes

    Calero, J. S.


    A development history and performance capability assessment is presented for the INTA-300 'Flamenco' sounding rocket prototype specimens. The Flamenco is a two-stage solid fuel rocket, based on British sounding rocket technology, that can lift 50 km payloads to altitudes of about 300 km. The flight of the first two prototypes, in 1974 and 1975, pointed to vibration problems which reduced the achievable apogee, and the third prototype's flight was marred by a premature detonation that destroyed the rocket. The fourth Flamenco flight, however, yielded much reliable data.

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

    Mohamed Sellam


    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.

  18. The TICTOP nozzle: a new nozzle contouring concept

    Frey, Manuel; Makowka, Konrad; Aichner, Thomas


    Currently, mainly two types of nozzle contouring methods are applied in space propulsion: the truncated ideal contour (TIC) and the thrust-optimized parabola (TOP). This article presents a new nozzle contouring method called TICTOP, combining elements of TIC and TOP design. The resulting nozzle is shock-free as the TIC and therefore does not induce restricted shock separation leading to excessive side-loads. Simultaneously, the TICTOP nozzle will allow higher nozzle wall exit pressures and hence give a better separation margin than is the case for a TIC. Hence, this new nozzle type combines the good properties of TIC and TOP nozzles and eliminates their drawbacks. It is especially suited for first stage application in launchers where flow separation and side-loads are design drivers.

  19. High regression rate hybrid rocket fuel grains with helical port structures

    Walker, Sean D.

    Hybrid rockets are popular in the aerospace industry due to their storage safety, simplicity, and controllability during rocket motor burn. However, they produce fuel regression rates typically 25% lower than solid fuel motors of the same thrust level. These lowered regression rates produce unacceptably high oxidizer-to-fuel (O/F) ratios that produce a potential for motor instability, nozzle erosion, and reduced motor duty cycles. To achieve O/F ratios that produce acceptable combustion characteristics, traditional cylindrical fuel ports are fabricated with very long length-to-diameter ratios to increase the total burning area. These high aspect ratios produce further reduced fuel regression rate and thrust levels, poor volumetric efficiency, and a potential for lateral structural loading issues during high thrust burns. In place of traditional cylindrical fuel ports, it is proposed that by researching the effects of centrifugal flow patterns introduced by embedded helical fuel port structures, a significant increase in fuel regression rates can be observed. The benefits of increasing volumetric efficiencies by lengthening the internal flow path will also be observed. The mechanisms of this increased fuel regression rate are driven by enhancing surface skin friction and reducing the effect of boundary layer "blowing" to enhance convective heat transfer to the fuel surface. Preliminary results using additive manufacturing to fabricate hybrid rocket fuel grains from acrylonitrile-butadiene-styrene (ABS) with embedded helical fuel port structures have been obtained, with burn-rate amplifications up to 3.0x than that of cylindrical fuel ports.

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


    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 e...

  1. Serrating Nozzle Surfaces for Complete Transfer of Droplets

    Kim, Chang-Jin " CJ" Yi, Uichong


    print surface. The basic principle of the present method is to reduce the liquid-solid surface energy of the nozzle to a level sufficiently below the intrinsic solid-liquid surface energy of the nozzle material so that the droplet is not pulled apart and, instead, the entire droplet volume becomes transferred to the print surface. In this method, the liquid-solid surface energy is reduced by introducing artificial surface roughness in the form of micromachined serrations on the inner nozzle surface (see figure). The method was tested in experiments on soft printing of DNA solutions and of deionized water through 0.5-mm-diameter nozzles, of which some were not serrated, some were partially serrated, and some were fully serrated. In the nozzles without serrations, transfer was incomplete; that is, residual liquids remained in the nozzles after printing. However, in every nozzle in which at least half the inner surface was serrated, complete transfer of droplets to the print surface was achieved.

  2. Design of Solid-fuel Rocket Attitude Control System Based on Monte Carlo Method%基于蒙特卡罗方法的固体火箭姿态控制系统设计

    王辰琳; 赵长见; 宋志国


    在固体火箭姿态控制系统设计过程中,为保证设计结果的可靠性,需要针对发动机性能、全箭质量及气动参数等进行拉偏仿真分析,各项偏差的大小及使用方法直接影响对固体火箭控制能力的需求。传统固体火箭姿态控制系统设计时,一般针对各项偏差进行极限拉偏组合仿真,导致设计结果较为保守。针对总体各项偏差量,建立概率模型,采用蒙特卡罗方法进行控制力分析。数学仿真结果表明,相比传统设计方法,在保证系统具有一定的可靠度情况下,大幅降低了对姿态控制系统的需求,优化了系统方案。%In the design process of solid-fuel rocket attitude control system, it is necessary to simulate based on population deviations of engine performance, whole solid-fuel rocket mass and aerodynamic parameter in order to assure the reliability of design results, because the using method of deviation factors are accounted for the demand of solid-fuel rocket control. The extreme value of population deviations are taken in the traditional design method, but it leads to more conservative design results. The probability models of population deviations are established, and then Monte Carlo methods are introduced to analysis the controlling force. The simulated results show that, compared to the traditional design method, the probability design method reduces the demand of solid-fuel rocket attitude control system and optimizes the system design scheme obviously.

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

    Davidian, Kenneth O.


    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.

  4. Carbon Nano-Composite Ablative Rocket Nozzles Project

    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...

  5. An Experimental Investigation of Rocket Ramjet Nozzle Assembly Base Pressures.


    psia as measured by the mercury manometer connected to the vacuum tank. The transducer was connected to a 10 volt D.C. power supply to provide the...pressure was allowed to rise in small steps to atmospheric pressure. The mercury manometer reading was subtracted from the barometric pressure...excitation voltage and to a digital voltmeter to record the voltage across the transducer. The vacuum tank was also connected to a 100 inch mercury

  6. An Experimental Study of a Rocket-Ramjet Nozzle Cluster


    approximately 0.5 psia which was easily within the transducer’s range. The vacuum chamber was also connected to a 100 in mercury manometer to provide a pressure...had dissipated and atmospheric conditions were reached. The mercury manometer reading was subtracted from the barometric pressure to yield the


    class investigated consisted of an AISI Type 316 stainless steel matrix incorporating a hard phase of titanium carbide ranging in content from 20% to...55% by volume. The results of the study indicated that under the test conditions, increases in the titanium carbide constituents did increase the

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

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


    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.

  9. Numerical study on drop formation through a micro nozzle

    Kim, Sung Il; Son, Gi Hun [Sogang Univ., Seoul (Korea, Republic of)


    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.

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

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


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

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

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


    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.

  12. Thermal analysis of the MC-1 chamber/nozzle

    Davis, Darrell


    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 .

  13. Fuel nozzle tube retention

    Cihlar, David William; Melton, Patrick Benedict


    A system for retaining a fuel nozzle premix tube includes a retention plate and a premix tube which extends downstream from an outlet of a premix passage defined along an aft side of a fuel plenum body. The premix tube includes an inlet end and a spring support feature which is disposed proximate to the inlet end. The premix tube extends through the retention plate. The spring retention feature is disposed between an aft side of the fuel plenum and the retention plate. The system further includes a spring which extends between the spring retention feature and the retention plate.

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

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


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

  15. Improved hybrid rocket fuel

    Dean, David L.


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

  16. Rocket propulsion elements

    Sutton, George P


    The definitive text on rocket propulsion-now revised to reflect advancements in the field For sixty years, Sutton's Rocket Propulsion Elements has been regarded as the single most authoritative sourcebook on rocket propulsion technology. As with the previous edition, coauthored with Oscar Biblarz, the Eighth Edition of Rocket Propulsion Elements offers a thorough introduction to basic principles of rocket propulsion for guided missiles, space flight, or satellite flight. It describes the physical mechanisms and designs for various types of rockets' and provides an unders

  17. Study on Horizontal Docking Assembly Method for Segment Thin-walled Solid Rocket Motor with Large Opening%薄壁大开口分段固体发动机卧式对接装配研究

    彭莎莎; 刘永盛; 宗路航; 罗玲莉; 苏昌银


    对薄壁大开口分段固体发动机卧式对接装配进行研究,应用ANSYS软件对燃烧室在重力作用下的应力、应变云图进行分析,提出变形识别与安全校正是解决结构件形变的一种方法。采用径向圆周10点均分校正法,采集燃烧室对接径向U型件边沿的变形量,制定了燃烧室变形安全校正值,经校正后中段与前段燃烧室对接径向边沿的10个点之间距离最大值为0.02mm。实现了Φ2m/分段式大型发动机对接装配。经Φ2m/分段式发动机地面试车获得成功证实校正值合理。%In this paper, the horizontal docking assembly of segment thin-walled large opening solid rocket motor is analyzed. The stress and strain of combustion chamber under the gravity is analyzed by ANSYS software and then deformation recognition and safe correction are proposed to prevent deformation of structural parts. The radial circular 10 average points correction method is used to collect the deformation amount of docking radial direction of edge of the U-shaped part, and to devise the safety correction value of deformation of single and double combustion chamber. Finally it is proved that the biggest gap of docking radial edge of 10 points between front and middle chamber is 0.02mm using different angle correction. Docking assembly of 2m/ segment solid rocket motor is realized and the correction value is proved reasonable after the success of ground fire test of 2m/segment solid rocket motor.

  18. Computation of ablation of thermal-protection layer in long-time working solid rocket motors%长时间工作固体火箭发动机燃烧室热防护层烧蚀计算

    张斌; 刘宇; 王长辉; 任军学


    为了研究长时间工作固体火箭发动机燃烧室的热防护性能,运用三方程烧蚀模型和运动边界显示差分格式,对长时间固体火箭发动机内绝热层烧蚀及温度场进行了耦合计算.计算得到了化学烧蚀率、扩散烧蚀率、燃烧室内壁温度等参数.计算结果表明,所研究的长时间工作发动机燃烧室烧蚀由扩散过程控制.此外,在求解烧蚀子程序时,提出了一种简便有效的赋初值方法.采用文中方法在得到合理计算结果的同时,使得烧蚀计算时间大大缩短.该项研究为长时间工作固体火箭发动机燃烧室热防护层设计提供了有效的分析手段.%In order to study thermal-protection function of long-time working solid rocket motors, insulation ablation and temperature field of long-time working solid rocket motors were coupling calculated by moving boundary explicit-difference method and three equations ablation model. Ablation velocity caused by chemical reaction or diffnsion process and temperature of inner wall were calculated. Numerical results show that the ablation is restricted by the diffusion process. In addition, when solving the ablation subroutine,a convenient and effective method for setting initial value was proposed. Using this new method, the reasonable results were obtained and the calculation time was greatly decreased. This research can offer an effective analytical method for long-time working solid rocket motors.

  19. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Zubrin, Robert; Snyder, Gary


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

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

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


    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.

  1. Factors of Influencing Bond Characteristics at II Interface of a Single Chamber Dual Thrust Solid Rocket Motor Grains%一种单室双推力发动机装药Ⅱ界面粘接性能研究

    何德伟; 刘戎; 侯少锋


    Based on the properties of single chamber dual thrust solid rocket motor grain,the effects of the thickness of thermal insulation,liner pre-curing and the vertical storage of pre-cured liner in vacuum on the bond characteristics at II interface were discussed in this paper.Some technical methods were proposed for improving the bond properties at interface.%根据单室双推力发动机装药的特点,对厚度绝热层、衬层的预反应及预固化衬层在真空状态下垂直存放等绝热衬层加工工艺条件对装药Ⅱ界面粘接性能的影响进行了研究,并提出了改善界面性能的技术途径。

  2. Effects of Ignition Process on the Internal Ballistics of Small-size Solid Rocket Motor%点火过程对小型固体火箭发动机内弹道影响

    刘赟; 王浩; 陶如意; 朱德龙


    为了研究某小型固体火箭发动机点火过程对内弹道性能的影响,建立包含点火过程的小型固体火箭发动机的内弹道数值研究模型和试验验证方案,对点火药量为1.0g、0.8g、0.6g和0.4g的发动机进行了内弹道数值研究,试验研究了点火药量为1.0g和0.8g两种情况,数值计算结果与试验结果基本一致.研究结果表明:小型固体火箭发动机由于燃烧室体积小,点火过程对内弹道影响明显;点火药量越大,点火药装填密度越大,引起压力峰值越大,稳定工作时间越短;经验估算得到的1.0g点火药量产生了过高的压力,是稳定压力的三倍,0.8g的点火药量能够满足点火可靠性和总体设计要求,产生最大压力为27.08 MPa,稳定工作时长159 ms,建议该小型火箭发动机的点火药量为0.8g.%To study the effects of ignition process on interior ballistic performance of a small-size solid rocket motor,the model for interior ballistic calculation including ignition process of small-size solid rocket motor and verification plan were set up. The numerical calculations interior ballistic performance with 1.0 g,0.8 g,0,6 g and 0.4 g igniter masses were carried out. The tests of 1.0 g and 0.8 g igniter mass were done. The compute result and test data were basically consistent. The results indicated that; the effect of ignition process on interior ballistic about small-size solid rocket motor is obvious for the small combustion chamber volume. The igniting charge density and pressure peak increase .stable operating time of motor decreases as the igniter mass increases. The 1.0 g igniter mass is estimated by empirical formula,1.0 g igniter mass brings too high pressure to motor,the value of pressure reaches three times the stable pressure. 0.8 g igniter mass meets ignition reliability and general design requirements,and the maximum pressure is 27.08 MPa,and stable work time is 159 ms,and 0.8 g igniter mass is suggested for the

  3. 固体火箭发动机自动回转系统的设计与实现%Design and Implementation of Auto-rotation System for Solid Rocket Engine

    赵锴; 何敏; 于殿泓; 郑毅


    In the process of high energy X-ray radiography detection, the disadvantages of low control accuracy and difficult radiation-proof for operators exist in manually controlling the rotation of solid rocket engine. Thus, the auto-rotation system based on OMRON CQM1H PLC has been designed. Two operating modes: auto and manual are equipped in this system to implement remote and high accurate automatic rotating function for solid rocket engine. The practice shows that the system features high stability and reliability, ease maintenance, and satisfies the requirements of explosion-proof, safety and high reliability.%在对固体火箭发动机进行高能X射线照相检测的过程中,针对采用人工方式存在回转固体火箭发动机存在控制精度不高、人员辐射防护困难等问题,设计了一种基于OMRON CQM1H PLC的自动回转系统.系统具备手动和自动两种运行模式,实现了固体火箭发动机的远程、高精度和自动化回转控制功能.实际应用表明,系统稳定性好、可靠性高且易于维护,符合检测现场防爆安全和高可靠性的要求.

  4. Mechanics Numerical Simulation of Filament ̄wound Solid Rocket Motor Shell%纤维缠绕固体火箭发动机壳体的力学数值仿真

    王凯; 鞠玉涛


    针对某种纤维缠绕复合材料固体火箭发动机壳体,依据原壳体在ABAQUS中建立了有限元模型,建模过程中将纤维缠绕层视为层合板来处理,简化了模型,并对发射瞬态进行有限元数值计算,分别求出了环向层6层的应力和缠绕层6层的应力,以便于强度分析,并对于铺层方案进行求解,得出数据结论,数据结果与实际情况符合,研究结果为纤维缠绕固体火箭发动机优化分析提供了理论依据。%Aiming at some certain kind of filament ̄wound composite solid rocket motor shel,on the basis of the original shel,the fi ̄nite element model is established. During the process of modeling,the filament winding layers are regarded as the laminated plate,and the finite element numerical calculation of the instantaneous state of the firing is done. Throug the stress calculation,its strength is analyzed to find the way of laying down that it needs to meet the strength requirement. The results provide a theoretical basis for better analysis of filament ̄wound solid rocket motor.

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

    Wang, K. C.


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

  6. Rockets two classic papers

    Goddard, Robert


    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

  7. PAR Analysis of HSR Nozzles

    Georgiadis, Nicholas J.


    Only recently has computational fluid dynamics (CFD) been relied upon to predict the flow details of advanced nozzle concepts. Computer hardware technology and flow solving techniques are advancing rapidly and CFD is now being used to analyze such complex flows. Validation studies are needed to assess the accuracy, reliability, and cost of such CFD analyses. At NASA Lewis, the PARC2D/3D full Navier-Stokes (FNS) codes are being applied to HSR-type nozzles. This report presents the results of two such PARC FNS analyses. The first is an analysis of the Pratt and Whitney 2D mixer-ejector nozzle, conducted by Dr. Yunho Choi (formerly of Sverdrup Technology-NASA Lewis Group). The second is an analysis of NASA-Langley's axisymmetric single flow plug nozzle, conducted by the author.

  8. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    Morgan, Morris H.; Gilinsky, Mikhail M.


    Three connected sub-projects were conducted under reported project. Partially, these sub-projects are directed to solving the problems conducted by the HU/FM&AL under two other NASA grants. The fundamental idea uniting these projects is to use untraditional 3D corrugated nozzle designs and additional methods for exhaust jet noise reduction without essential thrust lost and even with thrust augmentation. Such additional approaches are: (1) to add some solid, fluid, or gas mass at discrete locations to the main supersonic gas stream to minimize the negative influence of strong shock waves forming in propulsion systems; this mass addition may be accompanied by heat addition to the main stream as a result of the fuel combustion or by cooling of this stream as a result of the liquid mass evaporation and boiling; (2) to use porous or permeable nozzles and additional shells at the nozzle exit for preliminary cooling of exhaust hot jet and pressure compensation for non-design conditions (so-called continuous ejector with small mass flow rate; and (3) to propose and analyze new effective methods fuel injection into flow stream in air-breathing engines. Note that all these problems were formulated based on detailed descriptions of the main experimental facts observed at NASA Glenn Research Center. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of finding theoretical explanations for experimental facts and the creation of the accurate numerical simulation technique and prediction theory for solutions for current problems in propulsion systems solved by NASA and Navy agencies. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analysis for advanced aircraft and rocket engines. The F&AL Team uses analytical methods, numerical simulations, and possible experimental tests at the Hampton University campus. We will present some management activity

  9. Heterogeneous fuel for hybrid rocket

    Stickler, David B. (Inventor)


    Heterogeneous fuel compositions suitable for use in hybrid rocket engines and solid-fuel ramjet engines, The compositions include mixtures of a continuous phase, which forms a solid matrix, and a dispersed phase permanently distributed therein. The dispersed phase or the matrix vaporizes (or melts) and disperses into the gas flow much more rapidly than the other, creating depressions, voids and bumps within and on the surface of the remaining bulk material that continuously roughen its surface, This effect substantially enhances heat transfer from the combusting gas flow to the fuel surface, producing a correspondingly high burning rate, The dispersed phase may include solid particles, entrained liquid droplets, or gas-phase voids having dimensions roughly similar to the displacement scale height of the gas-flow boundary layer generated during combustion.

  10. Determination of 1-methyl-1H-1,2,4-triazole in soils contaminated by rocket fuel using solid-phase microextraction, isotope dilution and gas chromatography-mass spectrometry.

    Yegemova, Saltanat; Bakaikina, Nadezhda V; Kenessov, Bulat; Koziel, Jacek A; Nauryzbayev, Mikhail


    Environmental monitoring of Central Kazakhstan territories where heavy space booster rockets land requires fast, efficient, and inexpensive analytical methods. The goal of this study was to develop a method for quantitation of the most stable transformation product of rocket fuel, i.e., highly toxic unsymmetrical dimethylhydrazine - 1-methyl-1H-1,2,4-triazole (MTA) in soils using solid-phase microextraction (SPME) in combination with gas chromatography-mass spectrometry. Quantitation of organic compounds in soil samples by SPME is complicated by a matrix effect. Thus, an isotope dilution method was chosen using deuterated analyte (1-(trideuteromethyl)-1H-1,2,4-triazole; MTA-d3) for matrix effect control. The work included study of the matrix effect, optimization of a sample equilibration stage (time and temperature) after spiking MTA-d3 and validation of the developed method. Soils of different type and water content showed an order of magnitude difference in SPME effectiveness of the analyte. Isotope dilution minimized matrix effects. However, proper equilibration of MTA-d3 in soil was required. Complete MTA-d3 equilibration at temperatures below 40°C was not observed. Increase of temperature to 60°C and 80°C enhanced equilibration reaching theoretical MTA/MTA-d3 response ratios after 13 and 3h, respectively. Recoveries of MTA depended on concentrations of spiked MTA-d3 during method validation. Lowest spiked MTA-d3 concentration (0.24 mg kg(-1)) provided best MTA recoveries (91-121%). Addition of excess water to soil sample prior to SPME increased equilibration rate, but it also decreased method sensitivity. Method detection limit depended on soil type, water content, and was always below 1 mg kg(-1). The newly developed method is fully automated, and requires much lower time, labor and financial resources compared to known methods.

  11. Mixing and reaction processes in rocket based combined cycle and conventional rocket engines

    Lehman, Matthew Kurt

    Raman spectroscopy was used to make species measurements in two rocket engines. An airbreathing rocket, the rocket based combined cycle (RBCC) engine, and a conventional rocket were investigated. A supersonic rocket plume mixing with subsonic coflowing air characterizes the ejector mode of the RBCC engine. The mixing length required for the air and plume to become homogenous is a critical dimension. For the conventional rocket experiments, a gaseous oxygen/gaseous hydrogen single-element shear coaxial injector was used. Three chamber Mach number conditions, 0.1, 0.2 and 0.3, were chosen to assess the effect of Mach number on mixing. The flow within the chamber was entirely subsonic. For the RBCC experiments, vertical Raman line measurements were made at multiple axial locations downstream from the rocket nozzle plane. Species profiles assessed the mixing progress between the supersonic plume and subsonic air. For the conventional rocket, Raman line measurements were made downstream from the injector face. The goal was to evaluate the effect of increased chamber Mach number on injector mixing/reaction. For both engines, quantitative and qualitative information was collected for computational fluid dynamics (CFD development. The RBCC experiments were conducted for three distinct geometries. The primary flow path was a diffuse and afterburner design with a direct-connect air supply. A sea-level static (SLS) version and a thermally choked variant were also tested. The experimental results show that mixing length increases with additional coflow air in the DAB geometry. Operation of variable rocket mixture ratios at identical air flow rates did not significantly affect the mixing length. The thermally choked variant had a longer mixing length compared to the DAB geometry, and the SLS modification had a shorter mixing length due to a reduced air flow. The conventional rocket studies focused on the effect of chamber Mach number on primary injector mixing. Chamber Mach

  12. Experimental Study on Shock Wave Structures in Constant-area Passage of Cold Spray Nozzle

    Hiroshi KATANODA; Takeshi MATSUOKA; Kazuyasu MATSUO


    Cold spray is a technique to make a coating on a wide variety of mechanical or electric parts by spraying solid particles accelerated through a high-speed gas flow in a converging-diverging nozzle. In this study, pseudo-shock waves in a modeled cold spray nozzle as well as high-speed gas jets are visualized by schlieren technique. The schlieren photographs reveals the supersonic flow with shock train in the nozzle. Static pressure along the barrel wall is also measured. The location of the head of pseudo-shock wave and its pressure distribution along the nozzle wall are analytically explained by using a formula of pseudo-shock wave. The analytical results show that the supersonic flow accompanying shock wave in the nozzle should be treated as pseudo-shock wave instead of normal shock wave.

  13. The flight of uncontrolled rockets

    Gantmakher, F R; Dryden, H L


    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

  14. Nozzle geometry for organic vapor jet printing

    Forrest, Stephen R; McGraw, Gregory


    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.

  15. Particle Streak Velocimetry of Supersonic Nozzle Flows

    Willits, J. D.; Pourpoint, T. L.


    A novel velocimetry technique to probe the exhaust flow of a laboratory scale combustor is being developed. The technique combines the advantages of standard particle velocimetry techniques and the ultra-fast imaging capabilities of a streak camera to probe high speed flows near continuously with improved spatial and velocity resolution. This "Particle Streak Velocimetry" technique tracks laser illuminated seed particles at up to 236 picosecond temporal resolution allowing time-resolved measurement of one-dimensional flows exceeding 2000 m/s as are found in rocket nozzles and many other applications. Developmental tests with cold nitrogen have been performed to validate and troubleshoot the technique with supersonic flows of much lower velocity and without background noise due to combusting flow. Flow velocities on the order of 500 m/s have been probed with titanium dioxide particles and a continuous-wave laser diode. Single frame images containing multiple streaks are analyzed to find the average slope of all incident particles corresponding to the centerline axial flow velocity. Long term objectives for these tests are correlation of specific impulse to theoretical combustion predictions and direct comparisons between candidate green fuels and the industry standard, monomethylhydrazine, each tested under identical conditions.

  16. Potential Climate and Ozone Impacts From Hybrid Rocket Engine Emissions

    Ross, M.


    Hybrid rocket engines that use N2O as an oxidizer and a solid hydrocarbon (such as rubber) as a fuel are relatively new. Little is known about the composition of such hybrid engine emissions. General principles and visual inspection of hybrid plumes suggest significant soot and possibly NO emissions. Understanding hybrid rocket emissions is important because of the possibility that a fleet of hybrid powered suborbital rockets will be flying on the order of 1000 flights per year by 2020. The annual stratospheric emission for these rockets would be about 10 kilotons, equal to present day solid rocket motor (SRM) emissions. We present a preliminary analysis of the magnitude of (1) the radiative forcing from soot emissions and (2) the ozone depletion from soot and NO emissions associated with such a fleet of suborbital hybrid rockets. Because the details of the composition of hybrid emissions are unknown, it is not clear if the ozone depletion caused by these hybrid rockets would be more or less than the ozone depletion from SRMs. We also consider the climate implications associated with the N2O production and use requirements for hybrid rockets. Finally, we identify the most important data collection and modeling needs that are required to reliably assess the complete range of environmental impacts of a fleet of hybrid rockets.

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

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


    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.

  18. Numerical investigation on operation process of solid ducted rocket with postpositional gas generator%后置燃气发生器的新型固冲发动机工作过程数值模拟

    王云霞; 陈林泉; 杨向明; 张胜勇


    进行了后置燃气发生器的新型固体火箭冲压发动机直连式试验,并对实验演示用发动机补燃室三维内流场进行了数值模拟,将试验结果与数值模拟结果进行对比,验证了数值模拟的准确性.采用单因素比较分析的方法,研究了一次燃气喷射方式与补燃室长度对固冲发动机性能的影响.结果表明,一次燃气喷射角度为150°时的燃烧效率比60°时高14%,补燃室燃烧效率在一次燃气喷射角度为180°时达到最大值;8喷口的燃烧效率高于4喷口;补燃室长度增加,燃烧效率增大,补燃室长度为149 mm时的燃烧效率比99 mm仅高5%.%Direct-connected test of solid ducted rocket with postpositional gas generator was completed. Numerical investigation on the field of solid ducted rocket secondary chamber was presented. The simulation model was proved by comparing the test and CFD results. Furthermore, the effect of fuel injection style and length of secondary chamber on combustion efficiency was carried out by analysis method of single factor. The numerical results show that the combustion efficiency increases with either the number of fuel-inlet or the length of secondary chamber. In addition, the combustion efficiency was increased by 14 % when the fuel injection angle was equal to 150°, then increased to the maximum while the fuel injection angle was 180° , and the combustion efficiency was increased by only 5% when the length of secondary chamber increased to 149 mm.

  19. Fluidized-Solid-Fuel Injection Process

    Taylor, William


    Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

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

    Bulat Kenessov


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

    Jichao Hu; Juntao Chang; Wen Bao


    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-je...

  2. Porous solid backbone impregnation for electrochemical energy conversion systems

    Boulfrad, Samir


    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.

  3. Solid propellants.

    Marsh, H. E., Jr.; Hutchison, J. J.


    The basic principles underlying propulsion by rocket motor are examined together with the configuration of a solid propellant motor. Solid propellants and their preparation are discussed, giving attention to homogeneous propellants, composite propellants, energetic considerations in choosing a solid propellant, the processing of composite propellants, and some examples of new developments. The performance of solid propellants is investigated, taking into account characteristics velocity, the specific impulse, and performance calculations. Aspects of propellant development considered include nonperformance requirements for solid propellants, the approach to development, propellant mechanical properties, and future trends.

  4. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    Morris, Christopher I.


    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

  5. Ambipolar ion acceleration in an expanding magnetic nozzle

    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)


    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.

  6. The Ion Rocket


    discharge velocity w and the speci- fic impulse lap respectively cannot be increased. At this limit condition the thermal rocket oecouos "choked up...structural quality is 900 t, 3) In the case of an atomic-driven thermal rocket ’,;lth specific Ipipulse ISjy«8C0 sec and thrust to weight ratio « 1, the

  7. Model Rockets and Microchips.

    Fitzsimmons, Charles P.


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

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

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


    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.

  9. Thermodynamic cycle analysis on solid propellant air-turbo-rocket%固体推进剂吸气式涡轮火箭发动机的气动热力循环分析

    屠秋野; 丁朝霞; 陈玉春; 蔡元虎


    A numerical model for calculating performance of solid propellant air-turbo-rocket at design point was set up,and a relationship expression of fuel-air ratio of combustion based on compressor pressure ratio,tubine intet total temperature and turbine expansion ratio was put forward.And a relationship between the turbine expansion ratio and the bypass ratio was given.The effects of compressor pressure ratio, turbine inlet temperature,bypass ratio/turbine expansion ratio and flight Mach number on the specific thrust and specific impulse were analyzed quantitatively.%建立了固体推进剂吸气式涡轮火箭发动机的设计状态数值模型,提出了基于压气机增压比、涡轮前温度和涡轮落压比关系的燃烧室燃气与空气配比表达式,以及涡轮落压比和发动机涵道比的匹配关系.定量分析了压气机增压比、涡轮进口燃气总温、涵道比/涡轮落压比和飞行马赫数对固体推进剂吸气式涡轮火箭发动机的单位推力和比冲的影响.

  10. Kinetic energy of rainfall simulation nozzles

    Different spray nozzles are used frequently to simulate natural rain for soil erosion and chemical transport, particularly phosphorous (P), studies. Oscillating VeeJet nozzles are used mostly in soil erosion research while constant spray FullJet nozzles are commonly used for P transport. Several ch...

  11. Nozzle Bricks and Well Bricks

    Zhang Xiaohui; Peng Xigao


    1 Scope This standard specifies the classification,brand,technical requirements,test methods,inspection rules,marking,packing,transportation,storage,and quality certificate of nozzle bricks and well bricks.This standard is applicable to unfired and fired products.

  12. Nozzle for electric dispersion reactor

    Sisson, Warren G.; Basaran, Osman A.; Harris, Michael T.


    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.

  13. Liquid Atomization out of a Full Cone Pressure Swirl Nozzle

    Rimbert, Nicolas


    A thorough numerical, theoretical and experimental investigation of the liquid atomization in a full cone pressure swirl nozzle is presented. The first part is devoted to the study of the inner flow. CAD and CFD software are used in order to determine the most important parameters of the flow at the exit of nozzle. An important conclusion is the existence of two flow regions: one in relatively slow motion (the boundary layer) and a second nearly in solid rotation at a very high angular rate (about 100 000 rad/s) with a thickness of about 4/5th of the nozzle section. Then, a theoretical and experimental analysis of the flow outside the nozzle is carried out. In the theoretical section, the size of the biggest drops is successfully compared to results stemming from linear instability theory. However, it is also shown that this theory cannot explain the occurrence of small drops observed in the stability domain whose size are close to the Kolmogorov and Taylor turbulent length scale. A Phase Doppler Particle Ana...

  14. Use of Several Thermal Analysis Techniques on a Hypalon Paint Coating for the Solid Rocket Booster (SRB) of the Space Shuttle

    Wingard, Charles D.; Whitaker, Ann F. (Technical Monitor)


    White Hypalon paint is brush-applied as a moisture barrier coating over cork surfaces on each of the two Space Shuttle SRBs. Fine cracks have been observed in the Hypalon coating three times historically on laboratory witness panels, but never on flight hardware. Samples of the cracked and standard ("good") Hypalon were removed from witness panel cork surfaces, and were tested in 1998 by Thermogravimetric Analysis (TGA), TMA and Differential Scanning Calorimetry (DSC) thermal analysis techniques. The TGA data showed that at 700C, where only paint pigment solids remain, the cracked material had about 9 weight percent more material remaining than the standard material, probably indicating incomplete mixing of the paint before it was brush-applied to produce the cracked material. Use of the TMA film/fiber technique showed that the average modulus (stiffness) vs. temperature was about 3 to 6 times higher for the cracked material than for the standard material. The TMA data also showed that an increase in coating thickness for the cracked Hypalon was not a factor in the anomaly.

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

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


    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.

  16. Another Look at Rocket Thrust

    Hester, Brooke; Burris, Jennifer


    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…

  17. Another Look at Rocket Thrust

    Hester, Brooke; Burris, Jennifer


    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…

  18. Rocket University at KSC

    Sullivan, Steven J.


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

  19. Production, properties, and probing of Laval nozzles for cluster-jet targets

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


    A cluster-jet target achieves high and constant beam densities, which can be adjusted during operation. Therefore, it is highly eligible for storage ring experiments. By the expansion of pre-cooled gases within fine Laval nozzles a cluster source produces a continuous flow of cryogenic solid clusters. Essential for the production of clusters are the properties of the Laval nozzle. The production of such a nozzle with its complex inner geometry represents a major technical challenge. To ensure the production of these fine Laval nozzles for future internal targets, an improved production process based on the initial CERN production was recently developed at the University of Muenster. Systematic investigations on Laval nozzles with modified geometries will clarify the outstanding questions of the cluster production process. Moreover, this is very important for the deeper understanding of the cluster beam characteristics, in particular: the density, velocity, and mass, affected by the geometry of the nozzle. The production process and initial measurements with these new nozzles at the anti PANDA cluster-jet target prototype is presented and discussed.

  20. The effect of nozzle layout on droplet ejection of a piezo-electrically actuated micro-atomizer

    Yanying Feng; Zhaoying Zhou; Junhua Zhu; Guibin Du


    We study here effects of nozzle layout on the droplet ejection of a micro atomizer, which was fabricated with the arrayed nozzles by the MEMS technology and actuated by a piezoelectric disc. A theoretical model was first built for this piezoelectric-liquid-structure coupling system to characterize the acoustic wave propagation in the liquid chamber, which determined the droplet formation out of nozzles. The modal analysis was carried out numerically to predict resonant frequencies and simulate the corresponding pressure wave field. By comparing the amplitude contours of pressure wave on the liquid-solid interface at nozzle inlets with the designed nozzle layout, behaviors of the device under different vibration modes can be predicted. Experimentally, an impedance analyzer was used to measure the resonant frequencies of the system. Three types of atomizers with different nozzle layouts were fabricated for measuring the effect of nozzle distribution on the ejection performance. The visualization experiment of droplet generation was carried out and volume flow rates of these devices were measured. The good agreement between the experiment and the prediction proved that only the increase of nozzles may not enhance the droplet generation and a design of nozzle distribution from a viewpoint of frequency is necessary for a resonant related atomizer.

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

    Prisniakov, V.; Sitnikova, N.


    achievements V. Utkin and his pupils are crea- tion unique "mortar" launching of a heavy liquid rocket from shaft, the decision of a complex of prob- lems on maintenance ready for military action (continuous attendance) of liquid rockets in the filled condi-tion for many years, maintenance of stability of rockets at action on them of striking factors of nuclear explosion. With personal participation of academician IAA V. Utkin the following large scien- tific and technical results were received: (a) a military railway rocket complex with intercontinental solid-propellant rocket with starting weight of 105 tons and with 10 warheads; (b) a method of war manage-ment with the help of command rockets; (c) a method of definition of characteristics of means of overcoming of antimissile defense; (d) war intercontinental rockets with the increased accuracy, with the survivability, with the availability for action; (e) a commanding rocket. Design' decisions not ha- ving the analogues in world: (a) managements of flight solid-propellant an intercontinental ballistic missiles by means of a deviating head part; (b) managements solid-propellant rocket by method of inje- ction of gas in supercritical part of nozzle; (c) industrial introduction of the newest materials etc.V. Ut- kin is the active participant of works in the field of the international cooperation in research and deve- lopment of a space. In 1990 V. Utkin hold a high post of the director of ZSNIIMACH which is leading organization of a space-rocket industry of Russia. Under manual V. Utkin the Federal space program of Russia was developed. V. Utkin had huge authority as the chairman of Advice of the Main designers of the USSR. He was the co-chairman combined commission of experts V. Utkin - T. Stafford" on problems of maintenance joint manned flights. He was the chairman of Coordination advice under the program of researches on manned space complexes. V. Utkin dreamed to be the active participant of a new stage of the outer

  2. Dual-fuel, dual-mode rocket engine

    Martin, James A. (Inventor)


    The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

  3. 空空导弹固体发动机内弹道对导弹后体流场非定常影响的数值模拟%Numerical simulation on the unsteady effect of solid rocket motor internal ballistic on aft part flow field of air-to-air missile

    陈伟; 梁国柱


    空空导弹高空工作过程中,外部的超声速来流与其固体火箭发动机的尾部喷流相互作用,形成复杂的非定常尾部干扰流场,影响导弹后体的工作环境。为了探寻发动机内弹道对导弹后体结构的非定常影响,采用双组分气体的非定常CFD仿真模型对某空空导弹发动机工作期间的喷管内流场和导弹外流场进行一体化数值模拟,研究了由多个自由剪切层、激波、膨胀波等组成复杂干扰流场的结构,以及在发动机内弹道和外流速度的非定场效应影响下其变化过程,在此基础上定量分析了由此引起的尾流的温度和燃气的扩散,以及在不同内弹道阶段发动机对导弹后体结构产生的影响。计算结果表明,非定常干扰流场在导弹后体附近产生不断变化的低速涡流区域,加速了温度和燃气的扩散,致使导弹尾端面区域受到高温气体冲刷,进而降低导弹后体结构的安全性。因此,空空导弹的后体设计有必要考虑并减少发动机内弹道与导弹外流的非定场影响对导弹后体安全性所造成的潜在威胁。%In the high⁃altitude working period of missile,the interaction between the supersonic external flow and the tail jet flow of solid rocket motor derives complex unsteady interactive flow at the tail of missile,which affects the working environment of aft missile.In order to explore the unsteady influence of the internal ballistics on the structure of aft missile,an unsteady CFD simulation model of double component gas was adopted to simulate numerically the integration flow field including the inner flow of nozzle and the outer flow of missile during the working period of motor. The structure of interactive flow field composed of multiple free shear layers,shock waves and expansion waves,as well as its changing process under the unsteady effect of internal ballistic and outer flow velocity,were studied.Based on it

  4. Scaled Rocket Testing in Hypersonic Flow

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


    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.

  5. Spray nozzle for fire control

    Papavergos, Panayiotis G.


    The design of a spray nozzle for fire control is described. It produces a spray of gas and liquid having an oval transverse cross section and it comprises a mixing chamber with an oval transverse cross section adapted to induce a toroidal mixing pattern in pressurized gas and liquid introduced to the mixing chamber through a plurality of inlets. In a preferred embodiment the mixing chamber is toroidal. The spray nozzle produces an oval spray pattern for more efficient wetting of narrow passages and is suitable for fire control systems in vehicles or other confined spaces. Vehicles to which this invention may be applied include trains, armoured vehicles, ships, hovercraft, submarines, oil rigs, and most preferably, aircraft.

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

    Kalomba Mboyi


    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.

  7. 二次进气固冲发动机补燃室粒子沉积数值模拟%Numerical simulation of particle deposition in solid rocket ramjet chamber with secondary air inlets

    严聪; 张志峰; 马岑睿; 张成涛


    基于随机颗粒轨道模型和颗粒部分沉积模型,对二次进气固冲发动机补燃室壁面的粒子沉积进行了探索研究,讨论了空燃比、进气间距及流量分配比等参数对沉积的影响。结果表明,随空燃比增大,沉积区域有一个明显的前移,空燃比对沉积的影响分为两个阶段;进气间距对沉积的影响较复杂;流量分配比增大,沉积区域变化不大,但沉积总量增加。%Particle deposition in solid rocket ramjet chamber with secondary air inlets was numerically simulated based on ran-dom particle stochastic trajectory and particle partial deposition model. The effects of several parameters on deposition were dis-cussed,such as air-to-fuel ratio,inlet interval,flow distribution ratio etc.The results show that there is an obvious forward move with accretion of air-to-fuel rate,leading to two sections concerning the effects of air-to-fuel ratio.The effects of inlet interval on deposition are complex.The deposition regions change little but the deposition gross increases with the accretion of flow distribution ratio.

  8. Thermochemical Erosion of Hafnium Carbide Modified Carbon/Carbon Composite Throat in a Small Solid Rocket Motor%碳化铪改性炭/炭复合材料喉衬的热化学烧蚀

    沈学涛; 李克智; 李贺军; 冯涛; 张磊磊; 王斌


    Thermochemical erosion of hafnium carbide (HfC) modified carbon/carbon (C/C) composite throat was investigated using a hot-fire testing system in a small solid rocket motor. Chemical composition of the equilibrium combustion products was calculated by NASA-CEA program based on the principle of free energy minimum. The reaction products between the oxidizing species and HfC in the C/C composites were also calculated by FactSage.The results show that H2O, CO2, and OH are the main oxidizing species to consume carbon and HfC and generate thermochemical erosion to the throat materials. The interface of the fibers and matrix is preferentially ablated, and then erosion extends to the fibers and matrix. The formation of cone-shaped fibers and shell-shaped matrix is attributed to the thermochemical erosion of the flame.%采用小型固体火箭发动机研究了碳化铪(HfC)改性炭/炭复合材料喉衬的热化学烧蚀.借助基于最小自由能原理的NASA-CEA程序计算了燃气组成,借助化学热力学软件FactSage计算了燃气组分与碳、HfC的化学反应.结果表明,燃气中的H2O、CO2和OH是碳和HfC的主要氧化组分,使材料发生热化学烧蚀;纤维-基体界面是烧蚀的薄弱环节,烧蚀沿着界面分别向碳纤维和基体方向推进.热化学作用(氧化)造成纤维变细,顶端呈锥状,基体变薄,呈壳状.

  9. Computation of two-phase reacting flows in solid-liquid rocket ramjets%固液火箭冲压发动机内两相反应流场数值计算

    马智博; 朱建士


    In order to compute three-dimens ional reacting flow fieldsestabl ished in the chambers of solid-liquid rocket ramjets,the block implicit algorit hm was used to solve the Navier-Stokes equations about gas,the Continuum Formul a tion Model and k-ε-Ap model were used to characterize the turbulent fl ow and vaporizatio n of droplets.The modified k-ε-g model was adopted to represent the combu stion of the fuels.Calculations were carried out under different chamber configuration s and initial droplet diameters,from which the effects of these conditions on t he combustion efficiency were analyzed.The numerical results reveal the p r ocesses of droplet vaporization and combustion.%为了计算固液混合式火箭冲压发动机补燃室内的三维反应流场,用块隐式法求解气相Navier-Stokes方程组,用连续介质模型和k-ε-Ap模型计算颗粒相的湍流流动与蒸发过程,用修正的k-ε-g模型描述燃料的燃烧。为了分析发动机设计参数对反应流场的影响,用不同的条件进行计算,并由此分析了补燃室几何结构和液体燃料初始颗粒直径对燃烧效率的影响。算例表明,计算方法有效可行,数值结果能够反映流场结构、液体燃料的蒸发和两种燃料的燃烧过程。

  10. Frozen Chemistry Effects on Nozzle Performance Simulations

    Yoder, Dennis A.; Georgiadis, Nicholas J.; O'Gara, Michael R.


    Simulations of exhaust nozzle flows are typically conducted assuming the gas is calorically perfect, and typically modeled as air. However the gas inside a real nozzle is generally composed of combustion products whose thermodynamic properties may differ. In this study, the effect of gas model assumption on exhaust nozzle simulations is examined. The three methods considered model the nozzle exhaust gas as calorically perfect air, a calorically perfect exhaust gas mixture, and a frozen exhaust gas mixture. In the latter case the individual non-reacting species are tracked and modeled as a gas which is only thermally perfect. Performance parameters such as mass flow rate, gross thrust, and thrust coefficient are compared as are mean flow and turbulence profiles in the jet plume region. Nozzles which operate at low temperatures or have low subsonic exit Mach numbers experience relatively minor temperature variations inside the nozzle, and may be modeled as a calorically perfect gas. In those which operate at the opposite extreme conditions, variations in the thermodynamic properties can lead to different expansion behavior within the nozzle. Modeling these cases as a perfect exhaust gas flow rather than air captures much of the flow features of the frozen chemistry simulations. Use of the exhaust gas reduces the nozzle mass flow rate, but has little effect on the gross thrust. When reporting nozzle thrust coefficient results, however, it is important to use the appropriate gas model assumptions to compute the ideal exit velocity. Otherwise the values obtained may be an overly optimistic estimate of nozzle performance.

  11. Solid Phase Characterization of Solids Recovered from Failed Sluicer Arm

    Cooke, Gary A. [Hanford Site (HNF), Richland, WA (United States)


    The Enclosure to this memo discusses the solid phase characterization of a solid sample that was retrieved from the single-shell Tank 241-C-111 extended reach sluicer #2. This sluicer, removed from riser #3 on September 25, 2014, was found to have approximately 0.4 gallons of solid tank waste adhering to the nozzle area.

  12. US Rocket Propulsion Industrial Base Health Metrics

    Doreswamy, Rajiv


    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

  13. Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

    Messing, Wesley E.


    Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

  14. Optimization of Construction of the rocket-assisted projectile

    Arkhipov Vladimir


    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.

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

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


    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.

  16. Simulation of a Downsized FDM Nozzle

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


    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...... validated. This kind of simulations is facing multiple problems connected to the description of the material properties with temperature and pressure dependency....

  17. Through an Annular Turbine Nozzle

    Rainer Kurz


    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.

  18. Rocket Flight Path

    Jamie Waters


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

  19. Combustion Instabilities In Solid Propellant Rocket Motors


    34 AIAA Paper No. 98{3218. Avalon, G., Ugurtas, B. Grisch, F. and Bresson , F. (2000) \\Numerical Computations and Visualization Tests of the Flow Inside...Burners and Related Devices," Prog. in Comb. Energy and Science, Vol. 19, No. 4, pp. 313{364. Roberts , A.K. and Brownlee, W.G. (1971) \\Nonlinear

  20. Solid Rocket Testing at AFRL (Briefing Charts)



  1. Rockets in World War I


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

  2. Characterization of nal powders for rocket propulsion

    Merotto, L.; Galfetti, L.; Colombo, G.; DeLuca, L. T.


    Nanosized metal powders are known to significantly improve both solid and hybrid rocket performance, but have some drawbacks in terms of cost, safety, and possible influence on propellant mechanical properties. Performance enhancement through nanosized metal or metal hydride addition to solid fuels is currently under investigation also for hybrid propulsion. Therefore, a preburning characterization of the powders used in solid propellant or fuel manufacturing is useful to assess their effects on the ballistic properties and engine performance. An investigation concerning the comparative characterization of several aluminum powders having different particle size, age, and coating is presented. Surface area, morphology, chemical species concentration and characteristics, surface passivation layers, surface and subsurface chemical composition, ignition temperature and ignition delay are investigated. The aim of this characterization is to experimentally assess the effect of the nAl powder properties on ballistic characteristics of solid fuels and solidrocket composite-propellant performance, showing an increase in terms of Is caused by the decrease of two-phase losses in solid and a possible significant rf increase in hybrid rockets.

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


    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...

  4. Advanced Tactical Booster Technologies: Applications for Long-Range Rocket Systems


    System HIMARS [3] which can employ the MGM- 140 Army Tactical Missile System (ATacMS) solid propellant missile [4] to achieve the required range...launcher. 15. SUBJECT TERMS solid rocket; optimisation; artillery 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT 18. NUMBER OF PAGES 19a...the volumetrically constrained environment of a land-based launcher. Keywords— solid rocket; optimisation; artillery I. INTRODUCTION The Australian

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

    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.

  6. Effect of adiabatic inhibitor on afterward-dome insulation ablation in segmented solid rocket motors%分段式固体发动机绝热环对后封头绝热层烧蚀影响分析

    王建儒; 何国强; 许团委; 李江; 李强


    The ablation effect of adiabatic inhibitor on afterward-dome insulation was discussed systematically under combustion chamber operating condition for a typical SSRM(segmented solid rocket motor).Both numerical simulation results and tested data show that the existence of adiabatic inhibitor accelerate the gas flow speed to a certain extent,which result in two phase flow gas concentration and speed enhanced in certain scope of afterward-dome of combustion chamber,and make the ablation quantity of afterward-dome increased significantly.Furthermore,with the numerical analysis of modified SSRM,the result of this investigation is to provide the motor design characteristics and the system engineering approach used to effectively suppress the insulation ablation of afterward-dome for SSRMs.%针对某典型分段式固体发动机燃烧室内部绝热环在发动机工作过程中对后封头绝热层的烧蚀影响进行了系统分析,通过数值模拟和已有试验数据证明,分段式固体发动机工作过程中,分段药柱端面绝热环的存在一定程度上加速了燃烧室内部燃气流动的速度,导致燃烧室后封头一定范围内两相流燃气浓度的增加,从而导致后封头烧蚀量的大幅增加.在此基础上,结合改进型分段式固体发动机的数值分析,提出了分段式固体发动机在有效拟制后封头绝热层烧蚀加剧方面的设计措施和建议.

  7. Condensed phase particle acceleration analysis for overload test ground of solid rocket motor%固体发动机地面过载试验凝相粒子加速度分析

    许团委; 田维平; 王建儒; 李强


    Aiming at solid rocket motor overload ground simulation test, the acceleration theoretical analysis for condensate phase particle was carried out. The results show that the particles are mainly affected by the Coriolis acceleration and gas drag role, but the role of the centrifugal acceleration is not dominant. Finally,3D two-phase flow calculations as well as the ground simulation overload test under three groups overload conditions were carried out. Numerical calculation results reveal that particle deposition zone maintain between 60°~80° along the first quadrant to the four quadrant direction;the rotation test results reflect that serious e-rosion site and ablation site maintain between 70°~90° along the first quadrant to the four quadrant direction. Two results coincide well to a certain extent,which verify that the theoretical analysis is reasonable adn correct.%针对固体火箭发动机地面模拟过载试验,开展了燃烧室凝相粒子所受加速度理论分析,认为粒子主要受哥氏加速度和气相阻力的作用,离心加速度的作用并不占优。最后,针对3组不同过载下的缩比发动机,分别开展了三维两相流数值计算并进行了地面旋转模拟过载试验,数值计算得到的粒子聚集区维持在第Ⅰ象限偏第Ⅳ象限60°~80°之间;旋转试验解剖后的结果反映了颗粒冲刷严重部位及烧蚀严重部位维持在第Ⅰ象限偏第Ⅳ象限70°~90°之间。两种结果具有一定吻合性,也验证了理论分析的合理性和正确性。

  8. 固体火箭发动机壳体脱黏缺陷的热波检测%Debond defect detection in shell of solid rocket motor by thermal wave nondestructive testing

    宋远佳; 张炜; 杨正伟; 田干


    Based on thermal wave nondestructive testing technique, debond defects between composites shell and insulation layer of solid rocket motor ( SRM) were inspected by numerical analysis and experiment. The thermal images were enhanced and segmented by subtracting background method and watershed algorithms respectively. The hot spots area and the time characteristics of thermal sequence diagram were applied to quantificationally estimate the defects size and depth. The sample was also inspected by acoustic-ultrasonic ( AU) method for comparing testing effect. The results show that thermal wave technique has the disadvantage of fast, which can accurately test defect within 20 s; A best inspection time is exited, which was estimated at 7 s in experiment of the three 5 mm-deep defects; defect depths could be calculated by the material parameters and the estimated defect sizes. And the bigger of the debond area, the easier to be detected; the thermal images were more suitable for quantificational analysis than that by AU method.%基于热波技术,对固体火箭发动机复合材料壳体/绝热层脱黏缺陷进行检测研究,采用去除背景和分水岭方法对热图进行降噪、增强和分割处理,利用表面热斑区域的面积对缺陷尺寸进行估算,根据热图序列的时间特征计算缺陷的深度,并与声-超声检测结果进行比较.结果表明,热波技术检测速度快,20 s可对缺陷准确定位,检测结果直观;5 mm深的3个缺陷最佳检测时间均为7s;缺陷深度通过材料参数和缺陷尺寸进行计算,且缺陷越大,误差越小;与声-超声技术相比,热波检测更适合进行定量分析.

  9. Simulating Modulated Thermography of Cladding Debond in Solid Rockets%固体火箭包覆层脱粘调制红外热波检测法的数值模拟

    郭兴旺; 李苒笙; 丁蒙蒙


    调制红外热波无损检测是一种可以检测材料内部缺陷的先进技术,在固体火箭发动机包覆层脱粘的诊断中有一定的应用前景.以有限元法对固体火箭发动机包覆层脱粘的调制红外热波检测法进行数值模拟,研究加热条件和结构对可检测性的影响.得出可检信息参数(表面过余温度幅值、相位和相位差)随热波激励条件(调制频率、热流强度)和结构参数(材料、缺陷大小、深度、厚度等)变化的规律.针对典型材料和结构尺寸,给出最佳调制频率、盲频以及相位差的预测值.调制红外热波检测法可用于固体火箭发动机包覆层脱粘的无损检测,检测条件可通过计算机仿真来优化.%Modulated infrared thermal wave nondestructive testing (IR NDT) is an advanced technique of internal flaw inspection,and has certain application prospect in detecting cladding debond in solid rocket motors (SRM). The numerical simulation of modulated IR NDT for cladding debond in SRM is carried out by using finite element method, and the influences of heating condition and structure on the detectability are studied. The evolutions of informative parameters (i.e. the amplitude, phase and phase difference of surface excessive temperature) versus heating condition (i.e. the modulation frequency and heat flow intensity) and structural parameters (i.e. materials, defect size, depth and thickness) are obtained. The optimal modulation frequency, blind frequency and phase difference for typical material and structural size are predicted. Modulated thermography can be used in detecting cladding debond in SRM, and the test condition can be optimized through computer simulation.

  10. Baking Soda and Vinegar Rockets

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


    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…

  11. Baking Soda and Vinegar Rockets

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


    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. Laval nozzles for cluster-jet targets

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


    Cluster-jet targets are highly suited as internal targets for storage ring experiments. Here the target beam itself is produced by the expansion of pre-cooled gases within fine Laval nozzles. With such targets high and constant target beam thicknesses can be achieved and adjusted continuously during operation. At the prototype cluster-jet target for the PANDA experiment, which was built up and set successfully into operation at the University of Muenster, density structures within the cluster beam directly behind the nozzle have been observed. Therefore, a tilting system was installed, allowing for an adjustment of the nozzle system relative to the experimental setup. With this installation target densities of more than 2 x 10{sup 15} atoms/cm{sup 2} at a distance of 2.1 m behind the nozzle were achieved. To study the impact of the Laval nozzle geometry on the beam structures and the achievable density, an improved nozzle production method was established. With this technique it is possible to produce with high efficiency fine micrometer-sized nozzles with variable geometries, e.g. different opening angles, opening diameters or lengths of the exit trumpet. The method for the production of Laval nozzles are presented, and new perspectives are discussed.

  13. Kinematic characteristics analysis of rolling ball joint socket nozzle%滚动球窝喷管运动性能分析

    刘文芝; 任毅斌; 刘仲民; 庞明思; 赵永忠


    The motion force,complexity deformation and immeasurability of the mechanism are key problems in the system de-sign and performance requirement,when the ball,concave sphere,convex sphere and supporting body are under the rolling body con-tact impacting,in the rolling ball joint socket nozzle of a certain solid propellant rocket motor. In order to solve the dynamic perform-ance problems of the system,Spalart-Allmaras turbulence model was used in the calculation of nozzle flow field to simulate the igni-tion test and obtain the system thrust. Coupled rigid and flexible multi-body dynamics model was established. Characteristics of mo-tion on the system parts,contact stress and deformation were computed and analyzed. Analysis results verify rationality of system de-sign and kinematic characteristics in the pre-research state. A practical and feasible way for design of solid propellant rocket motor thrust vector control system was provided.%某固体火箭发动机滚动球窝喷管运动过程中,其接头内部和阴球、阳球及支撑体在滚动体的接触碰撞作用下,运动及受力、变形状态复杂,无法试验测量,是系统结构设计和功能发挥的关键问题。为解决系统运动性能,模拟热试车试验,采用单方程湍流模型计算喷管内流场,得到系统推力;建立系统刚柔耦合多体动力学模型,计算分析系统及接头内部主要构件的运动规律、接触应力及变形。在预研阶段,以检验系统结构设计和运动性能的合理性,同时为机械类固体火箭发动机推力向量控制系统的研制提供更实用和可行的理论分析方法。

  14. Introduction to Rocket Propulsion


    Von Braun; 1966. 4. Introduction to Ordnance Technology; IHSP 76-129; 1976. 5. Physics; D. Halliday and R. Resnick ; 1963. 6. Physics Tells Why: Luke Sky- walker in Star Wars when he said "Don’t get cocky." We never plan for EVERYTHING, though we like to think we do. As we’ve said, rocket

  15. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  16. This "Is" Rocket Science!

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela


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

  17. The Relativistic Rocket

    Antippa, Adel F.


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

  18. Rocketing to the Skies


    ONE sunny morning,we startedfor Yanqi Lake,Huairou District,Beijing,to try“rocket bungy”,so farthe only facility for this sport inChina.On the way there,wequestioned our courage and heartendurance. Entering the gate we saw,towering over a banner saying,

  19. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

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

    Ferragut, N. J.


    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.

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

    Ferragut, N. J.


    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.

  2. High Energy Density Additives for Hybrid Fuel Rockets to Improve Performance and Enhance Safety

    Jaffe, Richard L.


    We propose a conceptual study of prototype strained hydrocarbon molecules as high energy density additives for hybrid rocket fuels to boost the performance of these rockets without compromising safety and reliability. Use of these additives could extend the range of applications for which hybrid rockets become an attractive alternative to conventional solid or liquid fuel rockets. The objectives of the study were to confirm and quantify the high enthalpy of these strained molecules and to assess improvement in rocket performance that would be expected if these additives were blended with conventional fuels. We confirmed the chemical properties (including enthalpy) of these additives. However, the predicted improvement in rocket performance was too small to make this a useful strategy for boosting hybrid rocket performance.

  3. Flow and Noise from Septa Nozzles

    Zaman, K. B. M. Q.; Bridges, J. E.


    Flow and noise fields are explored for the concept of distributed propulsion. A model-scale experiment is performed with an 8:1 aspect ratio rectangular nozzle that is divided into six passages by five septa. The septa geometries are created by placing plastic inserts within the nozzle. It is found that the noise radiation from the septa nozzle can be significantly lower than that from the baseline rectangular nozzle. The reduction of noise is inferred to be due to the introduction of streamwise vortices in the flow. The streamwise vortices are produced by secondary flow within each passage. Thus, the geometry of the internal passages of the septa nozzle can have a large influence. The flow evolution is profoundly affected by slight changes in the geometry. These conclusions are reached by mostly experimental results of the flowfield aided by brief numerical simulations.

  4. External Cylindrical Nozzle with Controlled Vacuum

    V. N. Pil'gunov


    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

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

    Jichao Hu


    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.

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

    Hu, Jichao; Chang, Juntao; Bao, Wen


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

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

    Thorpe, Douglas G.


    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.

  8. Jet noise suppression by porous plug nozzles

    Bauer, A. B.; Kibens, V.; Wlezien, R. W.


    Jet noise suppression data presented earlier by Maestrello for porous plug nozzles were supplemented by the testing of a family of nozzles having an equivalent throat diameter of 11.77 cm. Two circular reference nozzles and eight plug nozzles having radius ratios of either 0.53 or 0.80 were tested at total pressure ratios of 1.60 to 4.00. Data were taken both with and without a forward motion or coannular flow jet, and some tests were made with a heated jet. Jet thrust was measured. The data were analyzed to show the effects of suppressor geometry on nozzle propulsive efficiency and jet noise. Aerodynamic testing of the nozzles was carried out in order to study the physical features that lead to the noise suppression. The aerodynamic flow phenomena were examined by the use of high speed shadowgraph cinematography, still shadowgraphs, extensive static pressure probe measurements, and two component laser Doppler velocimeter studies. The different measurement techniques correlated well with each other and demonstrated that the porous plug changes the shock cell structure of a standard nozzle into a series of smaller, periodic cell structures without strong shock waves. These structures become smaller in dimension and have reduced pressure variations as either the plug diameter or the porosity is increased, changes that also reduce the jet noise and decrease thrust efficiency.

  9. Fastrac Nozzle Design, Performance and Development

    Peters, Warren; Rogers, Pat; Lawrence, Tim; Davis, Darrell; DAgostino, Mark; Brown, Andy


    With the goal of lowering the cost of payload to orbit, NASA/MSFC (Marshall Space Flight Center) researched ways to decrease the complexity and cost of an engine system and its components for a small two-stage booster vehicle. The composite nozzle for this Fastrac Engine was designed, built and tested by MSFC with fabrication support and engineering from Thiokol-SEHO (Science and Engineering Huntsville Operation). The Fastrac nozzle uses materials, fabrication processes and design features that are inexpensive, simple and easily manufactured. As the low cost nozzle (and injector) design matured through the subscale tests and into full scale hot fire testing, X-34 chose the Fastrac engine for the propulsion plant for the X-34. Modifications were made to nozzle design in order to meet the new flight requirements. The nozzle design has evolved through subscale testing and manufacturing demonstrations to full CFD (Computational Fluid Dynamics), thermal, thermomechanical and dynamic analysis and the required component and engine system tests to validate the design. The Fastrac nozzle is now in final development hot fire testing and has successfully accumulated 66 hot fire tests and 1804 seconds on 18 different nozzles.

  10. Liquid rocket engine injectors

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


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

  11. Liquid Rocket Engine Testing


    booster rocket engines • 6000-10000 psia capabilities – Can use gaseous nitrogen, helium, or hydrogen to pressurize propellant tanks 9Distribution A...Approved for Public Release; Distribution Unlimited. PA Clearance 16493 Simplified Test Stand Layout Oxidizer  TankFuel  Tank High  Pressure   Gas (GN2...requires large, complex facilities to deliver propellant at the proper pressure , temperature, and flow rates • The enormous energies involved

  12. Inverse dynamic energy management for multi-constrained depleted shutdown of solid rocket%固体火箭多约束耗尽关机的动态逆能量管理方法

    张志健; 王小虎


    To solve the energy management problem of a solid rocket with multi-constraining depleted shutdown,a novel inverse dynamic energy management (IEM) method was proposed.Firstly,a novel model based on excess velocity capability was estab-lished,and its performance was studied.Secondly,a closed-loop IEM method with two specific realizations was proposed, and the constrained and unconstrained convergence condition was discussed.Finally,the IEM's performance,energy manage results and the difference with general energy management ( GEM) and spline energy management ( SEM) were verified by computer simulation, which demonstrates a strong robust,high control precision, low attitude maneuver,small terminal angle of IEM.Indeed,IEM can suc-cessfully manage 2%~83.3% energy under a sort of fixed parameters.As the ratio of excess energy is below 41.3%, the terminal ve-locity is less than 0.25 m/s,the flight-path angle is less than 0.014°,the attitude angular velocity is less than 0.25°/s,and the angle of attack is less than 5° .%针对具有速度控制能力的固体火箭多约束耗尽关机问题,建立以多余视速度增量为状态量的新型能量管理模型,分析了该模型能量管理动态特性,提出一种闭环动态逆能量管理( IEM)方式,得到了无约束和有约束关机时能量管理的收敛条件,并据此设计了IEM的2种具体实现。最后以数学仿真验证了该方法的特性、能量管理效果,分析了与其他2种闭环能量管理方法---通用能量管理(GEM)、样条能量管理(SEM)的异同。结果表明,IEM方法鲁棒性强、精度高、姿态变化缓慢,能量管理范围大,一组固定的参数实现了2.1%~83.3%的能量管理;当能耗小于41.3%时,关机点速度偏差小于0.25 m/s,速度倾角小于0.014°,姿态角速度小于0.25°/s,攻角小于5°。

  13. Method of cooling gas only nozzle fuel tip

    Bechtel, William Theodore; Fitts, David Orus; DeLeonardo, Guy Wayne


    A diffusion flame nozzle gas tip is provided to convert a dual fuel nozzle to a gas only nozzle. The nozle tip diverts compressor discharge air from the passage feeding the diffusion nozzle air swirl vanes to a region vacated by removal of the dual fuel components, so that the diverted compressor discharge air can flow to and through effusion holes in the end cap plate of the nozzle tip. In a preferred embodiment, the nozzle gas tip defines a cavity for receiving the compressor discharge air from a peripheral passage of the nozzle for flow through the effusion openings defined in the end cap plate.

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

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


    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.

  15. Design of a new type vapor recovery system nozzle

    Fu, S. H.; Cao, G. J.; Zhang, D. S.


    To settle the problem of low-efficiency recovery for Vapor recovery system nozzle, this paper advances a purely mechanical structure of the self-sealing refueling VRS nozzle. The structure, operating principle and controlled process of the nozzle is given. And an application of the nozzle is discussed. All indicated that the nozzle has a reasonable structure, can fuel and vapor recovery simultaneous start and stop. And thus improve the recovery efficiency and reduce oil leakage.

  16. Calibrating feedwater flow nozzles in-situ

    Caudill, M. [Tri-State Generation and Transmission, Inc., Montrose, CA (United States); Diaz-Tous, I.; Murphy, S.; Leggett, M.; Crandall, C. [ENCOR-AMERICA, Inc., Mountain View, CA (United States)


    This paper presents a new method for in-situ calibration of feedwater flow nozzles wherein feedwater flow is determined indirectly by performing a high accuracy heat balance around the highest-pressure feedwater heater. It is often difficult to reliably measure feedwater flow. Over the life of a power plant, the feedwater nozzle can accumulate deposits, erode, or suffer other damage that can render the original nozzle calibration inaccurate. Recalibration of installed feedwater flow nozzles is expensive and time consuming. Traditionally, the nozzle is cut out of the piping and sent to a laboratory for recalibration, which can be an especially difficult, expensive, and time-consuming task when involving high pressure feedwater lines. ENCOR-AMERICA, INC. has developed an accurate and cost-effective method of calibrating feedwater nozzles in-situ as previously reported at the 1994 EPRI Heat Rate Improvement Conference. In this method, feedwater flow and differential pressure across the nozzle are measured concurrently. The feedwater flow is determined indirectly by performing a heat balance around the highest-pressure feedwater heater. Extraction steam to the feedwater heater is measured by use of a high accuracy turbine flowmeter. The meters used have been calibrated at an independent laboratory with a primary or secondary device traceable to the NIST. In this paper, a new variation on the above method is reported. The new approach measures the heater drains and vent flows instead of the extraction steam flow. Test theory and instrumentation will be discussed. Results of in-situ feedwater nozzle calibration tests performed at two units owned by Tri-State Generation and Transmission Association will be presented.

  17. Nuclear Thermal Rocket Propulsion Systems


    NUCLEAR THERMAL ROCKET PROPULSION SYSTEMS, IAA WHITE PAPER PARIS, FRANCE, MARCH 2005 Lt Col Timothy J. Lawrence U.S. Air Force Academy...YYYY) 18-03-2005 2. REPORT TYPE White Paper 3. DATES COVERED (From - To) 18 Mar 2005 4. TITLE AND SUBTITLE NUCLEAR THERMAL ROCKET PROPULSION...reduce radiation exposure, is to have a high energy system like a nuclear thermal rocket that can get the payload to the destination in the fastest

  18. Rocket Assembly and Checkout Facility

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

  19. Rocket-Powered Parachutes Rescue Entire Planes


    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.

  20. Rocket + Science = Dialogue

    Morris,Bruce; Sullivan, Greg; Burkey, Martin


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

  1. Pengembangan Rancangan Nozzle Waterjet untuk Meningkatkan Kecepatan Renang pada Tank BMP-3F (Infantry Fighting Vehicle

    Rozzaqi Anata


    Full Text Available Negara Kepulauan Republik Indonesia (NKRI memiliki wilayah  perairan yang luas, sehingga pertahanan negara di sektor perairan menjadi lebih dirapatkan. Strategi yang dibentuk adalah dengan memproduksi dan membeli kendaraan tempur. Salah satu kendaraan yang dibeli adalah tank amphibi BMP-3F buatan rusia. Kendaraan tank ini ketika dioperasikan di perairan hanya mencapai kecepatan 10 km/h, oleh karena itu akan dilakukan pengembangan perancangan nozzle waterjet untuk dapat meningkatkan kecepatan renang dari tank BMP-3F. Sehingga dilakukan beberapa modifikasi dari variasi nozzle yang akan dianalisa menggunakan SolidWorks yakni variasi diameter nozzle dari kondisi awal 140 mm hingga menjadi 110 mm, serta perbedaan sudut nozzle yang nantinya akan membentuk cone, dari 10 hingga 40, serta penambahan ulir pada sisi outlet water jet. Dari hasil analisa data dan perhitungan diperoleh untuk hasil thrust tertinggi dengan bentuk nozzle cone variasi 40 menghasilkan thrust sebesar 146,347 kN dengan kecepatan renang meningkat sebesar 89% dari kecepatan awal yakni menjadi 10,017 knot pada saat thrust deduction factor sebesar 0,3076.

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

    Ingemar Skoog, A.


    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.




    Full Text Available Liquid rocket engines have variety of propellant combinations which produces very high specific impulses. It is due to this fact; very high heat fluxes are incident on the combustion chamber and the nozzle walls. In order to deal with these heat fluxes, a wide range of cooling techniques have been employed, out of which a combination of film cooling and regenerative cooling promises to be the most effective one. The present study involves the numerical analysis of combustion in a typical film cooled cryogenic rocket engine thrust chamber considering the combustion of the fuel, heat transfer through the chamber walls and the fluid flow simultaneously. Analysis was done for a typical rocket engine thrust chamber with a single coaxial injector which uses gaseous hydrogen as the fuel and liquid oxygen as the oxidizer.

  4. Cycle Trades for Nuclear Thermal Rocket Propulsion Systems

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


    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.

  5. Mars Rocket Propulsion System

    Zubrin, Robert; Harber, Dan; Nabors, Sammy


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

  6. Nozzle assembly for an earth boring drill bit

    Madigan, J. A.


    A nozzle assembly for an earth boring drill bit of the type adapted to receive drilling fluid under pressure and having a nozzle bore in the bottom thereof positioned closely adjacent the well bore bottom when the bit is in engagement therewith with the bore having inner and outer portions. The nozzle assembly comprises a generally cylindrical nozzle member of abrasion and erosion resistant material, selected from a plurality of such members, each being of the same outer diameter but having passaging therein of different cross-sectional area. The nozzle member is adapted to be fitted in the inner portion of the nozzle bore in sealing relationship therewith for forming a first seal for the nozzle assembly. The nozzle assembly further comprises a locknut, separate from the nozzle member, for detachbably securing the nozzle member in the nozzle bore, formed at least in part of an abrasion and erosion resistant material. The locknut has a threaded side wall engageable with the outer portion of the nozzle bore, and an aperture therethrough for enabling a stream of drilling fluid from the nozzle member to flow therethrough and being so configured in section as to receive a tool for turning the lockout to install it in and remove it from the nozzle bore.

  7. What fuel for a rocket?

    Miranda, E N


    Elementary concepts from general physics and thermodynamics have been used to analyze rocket propulsion. Making some reasonable assumptions, an expression for the exit velocity of the gases is found. From that expression one can conclude what are the desired properties for a rocket fuel.

  8. Rocket launchers as passive controllers

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


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

  9. Pulse Detonation Rocket MHD Power Experiment

    Litchford, Ron J.; Cook, Stephen (Technical Monitor)


    A pulse detonation research engine (MSFC (Marshall Space Flight Center) Model PDRE (Pulse Detonation Rocket Engine) G-2) has been developed for the purpose of examining integrated propulsion and magnetohydrodynamic power generation applications. The engine is based on a rectangular cross-section tube coupled to a converging-diverging nozzle, which is in turn attached to a segmented Faraday channel. As part of the shakedown testing activity, the pressure wave was interrogated along the length of the engine while running on hydrogen/oxygen propellants. Rapid transition to detonation wave propagation was insured through the use of a short Schelkin spiral near the head of the engine. The measured detonation wave velocities were in excess of 2500 m/s in agreement with the theoretical C-J velocity. The engine was first tested in a straight tube configuration without a nozzle, and the time resolved thrust was measured simultaneously with the head-end pressure. Similar measurements were made with the converging-diverging nozzle attached. The time correlation of the thrust and head-end pressure data was found to be excellent. The major purpose of the converging-diverging nozzle was to configure the engine for driving an MHD generator for the direct production of electrical power. Additional tests were therefore necessary in which seed (cesium-hydroxide dissolved in methanol) was directly injected into the engine as a spray. The exhaust plume was then interrogated with a microwave interferometer in an attempt to characterize the plasma conditions, and emission spectroscopy measurements were also acquired. Data reduction efforts indicate that the plasma exhaust is very highly ionized, although there is some uncertainty at this time as to the relative abundance of negative OH ions. The emission spectroscopy data provided some indication of the species in the exhaust as well as a measurement of temperature. A 24-electrode-pair segmented Faraday channel and 0.6 Tesla permanent

  10. Pulse Detonation Rocket MHD Power Experiment

    Litchford, Ron J.; Cook, Stephen (Technical Monitor)


    A pulse detonation research engine (MSFC (Marshall Space Flight Center) Model PDRE (Pulse Detonation Rocket Engine) G-2) has been developed for the purpose of examining integrated propulsion and magnetohydrodynamic power generation applications. The engine is based on a rectangular cross-section tube coupled to a converging-diverging nozzle, which is in turn attached to a segmented Faraday channel. As part of the shakedown testing activity, the pressure wave was interrogated along the length of the engine while running on hydrogen/oxygen propellants. Rapid transition to detonation wave propagation was insured through the use of a short Schelkin spiral near the head of the engine. The measured detonation wave velocities were in excess of 2500 m/s in agreement with the theoretical C-J velocity. The engine was first tested in a straight tube configuration without a nozzle, and the time resolved thrust was measured simultaneously with the head-end pressure. Similar measurements were made with the converging-diverging nozzle attached. The time correlation of the thrust and head-end pressure data was found to be excellent. The major purpose of the converging-diverging nozzle was to configure the engine for driving an MHD generator for the direct production of electrical power. Additional tests were therefore necessary in which seed (cesium-hydroxide dissolved in methanol) was directly injected into the engine as a spray. The exhaust plume was then interrogated with a microwave interferometer in an attempt to characterize the plasma conditions, and emission spectroscopy measurements were also acquired. Data reduction efforts indicate that the plasma exhaust is very highly ionized, although there is some uncertainty at this time as to the relative abundance of negative OH ions. The emission spectroscopy data provided some indication of the species in the exhaust as well as a measurement of temperature. A 24-electrode-pair segmented Faraday channel and 0.6 Tesla permanent

  11. Rocket Plume Scaling for Orion Wind Tunnel Testing

    Brauckmann, Gregory J.; Greathouse, James S.; White, Molly E.


    A wind tunnel test program was undertaken to assess the jet interaction effects caused by the various solid rocket motors used on the Orion Launch Abort Vehicle (LAV). These interactions of the external flowfield and the various rocket plumes can cause localized aerodynamic disturbances yielding significant and highly non-linear control amplifications and attenuations. This paper discusses the scaling methodologies used to model the flight plumes in the wind tunnel using cold air as the simulant gas. Comparisons of predicted flight, predicted wind tunnel, and measured wind tunnel forces-and-moments and plume flowfields are made to assess the effectiveness of the selected scaling methodologies.

  12. Upper Stage Engine Composite Nozzle Extensions

    Valentine, Peter G.; Allen, Lee R.; Gradl, Paul R.; Greene, Sandra E.; Sullivan, Brian J.; Weller, Leslie J.; Koenig, John R.; Cuneo, Jacques C.; Thompson, James; Brown, Aaron; hide


    Carbon-carbon (C-C) composite nozzle extensions are of interest for use on a variety of launch vehicle upper stage engines and in-space propulsion systems. The C-C nozzle extension technology and test capabilities being developed are intended to support National Aeronautics and Space Administration (NASA) and United States Air Force (USAF) requirements, as well as broader industry needs. Recent and on-going efforts at the Marshall Space Flight Center (MSFC) are aimed at both (a) further developing the technology and databases for nozzle extensions fabricated from specific CC materials, and (b) developing and demonstrating low-cost capabilities for testing composite nozzle extensions. At present, materials development work is concentrating on developing a database for lyocell-based C-C that can be used for upper stage engine nozzle extension design, modeling, and analysis efforts. Lyocell-based C-C behaves in a manner similar to rayon-based CC, but does not have the environmental issues associated with the use of rayon. Future work will also further investigate technology and database gaps and needs for more-established polyacrylonitrile- (PAN-) based C-C's. As a low-cost means of being able to rapidly test and screen nozzle extension materials and structures, MSFC has recently established and demonstrated a test rig at MSFC's Test Stand (TS) 115 for testing subscale nozzle extensions with 3.5-inch inside diameters at the attachment plane. Test durations of up to 120 seconds have been demonstrated using oxygen/hydrogen propellants. Other propellant combinations, including the use of hydrocarbon fuels, can be used if desired. Another test capability being developed will allow the testing of larger nozzle extensions (13.5- inch inside diameters at the attachment plane) in environments more similar to those of actual oxygen/hydrogen upper stage engines. Two C-C nozzle extensions (one lyocell-based, one PAN-based) have been fabricated for testing with the larger

  13. Integrity of the Plasma Magnetic Nozzle

    Gerwin, Richard A.


    This report examines the physics governing certain aspects of plasma propellant flow through a magnetic nozzle, specifically the integrity of the interface between the plasma and the nozzle s magnetic field. The injection of 100s of eV plasma into a magnetic flux nozzle that converts thermal energy into directed thrust is fundamental to enabling 10 000s of seconds specific impulse and 10s of kW/kg specific power piloted interplanetary propulsion. An expression for the initial thickness of the interface is derived and found to be approx.10(exp -2) m. An algorithm is reviewed and applied to compare classical resistivity to gradient-driven microturbulent (anomalous) resistivity, in terms of the spatial rate and time integral of resistive interface broadening, which can then be related to the geometry of the nozzle. An algorithm characterizing plasma temperature, density, and velocity dependencies is derived and found to be comparable to classical resistivity at local plasma temperatures of approx. 200 eV. Macroscopic flute-mode instabilities in regions of "adverse magnetic curvature" are discussed; a growth rate formula is derived and found to be one to two e-foldings of the most unstable Rayleigh-Taylor (RT) mode. After establishing the necessity of incorporating the Hall effect into Ohm s law (allowing full Hall current to flow and concomitant plasma rotation), a critical nozzle length expression is derived in which the interface thickness is limited to about 1 ion gyroradius.

  14. Compressible vortex loops: Effect of nozzle geometry

    Zare-Behtash, H. [School of MACE, University of Manchester, M60 1QD (United Kingdom)], E-mail:; Kontis, K. [School of MACE, University of Manchester, M60 1QD (United Kingdom)], E-mail:; Gongora-Orozco, N. [School of MACE, University of Manchester, M60 1QD (United Kingdom); Takayama, K. [Tohoku University, Shock Wave Research Centre, Sendai 980-8577 (Japan)


    Vortex loops are fundamental building blocks of supersonic free jets. Isolating them allows for an easier study and better understanding of such flows. The present study looks at the behaviour of compressible vortex loops of different shapes, generated due to the diffraction of a shock wave from a shock tube with different exit nozzle geometries. These include a 15 mm diameter circular nozzle, two elliptical nozzles with minor to major axis ratios of 0.4 and 0.6, a 30 x 30 mm square nozzle, and finally two exotic nozzles resembling a pair of lips with minor to major axis ratios of 0.2 and 0.5. The experiments were performed for diaphragm pressure ratios of P{sub 4}/P{sub 1}=4, 8, and 12, with P{sub 4} and P{sub 1} being the pressures within the high pressure and low pressure compartments of the shock tube, respectively. High-speed schlieren photography as well as PIV measurements of both stream-wise and head-on flows have been conducted.

  15. Nuclear Rocket Engine Reactor

    Lanin, Anatoly


    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.

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

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


    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.

  17. Discharge coefficient of small sonic nozzles

    Yin Zhao-Qin


    Full Text Available The purpose of this investigation is to understand flow characteristics in mini/micro sonic nozzles, in order to precisely measure and control miniscule flowrates. Experimental and numerical simulation methods have been used to study critical flow Venturi nozzles. The results show that the nozzle’s size and shape influence gas flow characteristics which leading the boundary layer thickness to change, and then impact on the discharge coefficient. With the diameter of sonic nozzle throat decreasing, the discharge coefficient reduces. The maximum discharge coefficient exits in the condition of the inlet surface radius being double the throat diameter. The longer the diffuser section, the smaller the discharge coefficient becomes. Diffuser angle affects the discharge coefficient slightly.

  18. New inlet nozzle assembly: C Reactor

    Calkin, J.F.


    The use of self-supported fuel elements in ribless Zircaloy-2 tubes at C-Reactor requires some inlet nozzle modification to allow charging of the larger overall diameter fuel pieces. A new nozzle assembly has been developed (by Equipment Development Operation -- IPD) which will allow use of the new fuel pieces and at the same time increase the reliability of the header-to-tube piping and reduce pumping power losses. Flow test data were requested for the new assembly and the results of these tests are presented herein. This report also presents a comparison of the header to tube energy losses for the various reactor inlet nozzle assemblies which are currently used on the Hanford production reactors.

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

    Cummings, Chase B.


    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.

  20. Effect of gaseous and solid simulated jet plumes on an 040A space shuttle launch configuration at m=1.6 to 2.2

    Dods, J. B., Jr.; Brownson, J. J.; Kassner, D. L.; Blackwell, K. L.; Decker, J. P.; Roberts, B. B.


    The effect of plume-induced flow separation and aspiration effects due to operation of both orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the Space Shuttle Vehicle is determined. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold-gas jet supplied by an auxiliary 200-atm air supply system and solid-body plume simulators. The aerodynamic effects produced by these two simulation procedures are compared. The parameters most significantly affected by the jet plumes are pitching moment, elevon control effectiveness, axial force, and orbiter wing loads. The solid rocket motor (SRM) plumes have the largest effect on the aerodynamic characteristics. The effect of the orbiter plumes in combination with the SRM plumes is also significant. Variations in the nozzle design parameters and configuration changes can reduce the jet plume-induced aerodynamic effects.

  1. Combustor nozzles in gas turbine engines

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


    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.

  2. Experimental investigation of flow through planar double divergent nozzles

    Arora, Rajat; Vaidyanathan, Aravind


    Dual bell nozzle is one of the feasible and cost effective techniques for altitude adaptation. Planar double divergent nozzle with a rectangular cross section was designed for two different NPR's to simulate and investigate the flow regimes similar to those inside the dual bell nozzle. Measurements involved flow visualization using Schlieren technique and wall static pressure measurements. The flow transition between the two nozzles at the respective inflection points and the formation of recirculation region due to flow separation was analyzed in detail. Cold flow tests were performed on the double divergent nozzle in the over-expanded conditions to study the shock wave characteristics. The results obtained from the two independent double divergent nozzles were compared with those obtained from a single divergent nozzle of the same area ratio. From the experiments it was observed that inflection angle played a key role in defining the type of shock structures existing inside the double divergent nozzles.

  3. Submerged entry nozzle clogging during continuous casting of Al-killed steel

    F. Tehovnik


    Full Text Available Nozzle clogging is a common problem in the production of continuously cast Al-killed steels. Clogging occurs when there are solid inclusions in molten steel at casting temperatures. SENs (Submerged entry nozzles from continuous casting of Al-killed low alloy steel grades with increased content of sulfur (0,020 to 0,035 % S were examined. The examinations revealed that the deposits are mainly alumina based, with spinel and sulfur inclusions and some entrapped steel melt. It was concluded that the process of clogging begins when the steel melt infiltrates the refractory and removes the protective zirconia surface, thus allowing the adhesion of fine solid aluminates, which form the deposits.

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

    Ioan RUSU


    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.

  5. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Greene, William D. (Inventor)


    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

  6. Actively Cooled Ceramic Composite Nozzle Material Project

    National Aeronautics and Space Administration — For Next Generation Launch Vehicles (NGLV), Either a Rocket-based or Turbine-based Combined Cycle (RBCC or TBCC) engine will power the Next Generation Launch Vehicle...

  7. Noise of Embedded High Aspect Ratio Nozzles

    Bridges, James E.


    A family of high aspect ratio nozzles were designed to provide a parametric database of canonical embedded propulsion concepts. Nozzle throat geometries with aspect ratios of 2:1, 4:1, and 8:1 were chosen, all with convergent nozzle areas. The transition from the typical round duct to the rectangular nozzle was designed very carefully to produce a flow at the nozzle exit that was uniform and free from swirl. Once the basic rectangular nozzles were designed, external features common to embedded propulsion systems were added: extended lower lip (a.k.a. bevel, aft deck), differing sidewalls, and chevrons. For the latter detailed Reynolds-averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) simulations were made to predict the thrust performance and to optimize parameters such as bevel length, and chevron penetration and azimuthal curvature. Seventeen of these nozzles were fabricated at a scale providing a 2.13 inch diameter equivalent area throat." ! The seventeen nozzles were tested for far-field noise and a few data were presented here on the effect of aspect ratio, bevel length, and chevron count and penetration. The sound field of the 2:1 aspect ratio rectangular jet was very nearly axisymmetric, but the 4:1 and 8:1 were not, the noise on their minor axes being louder than the major axes. Adding bevel length increased the noise of these nozzles, especially on their minor axes, both toward the long and short sides of the beveled nozzle. Chevrons were only added to the 2:1 rectangular jet. Adding 4 chevrons per wide side produced some decrease at aft angles, but increased the high frequency noise at right angles to the jet flow. This trend increased with increasing chevron penetration. Doubling the number of chevrons while maintaining their penetration decreased these effects. Empirical models of the parametric effect of these nozzles were constructed and quantify the trends stated above." Because it is the objective of the Supersonics Project that

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

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


    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

  9. British used Congreve Rockets to Attack Napoleon


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

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

    Chunkyraj Kh


    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.

  11. Internal performance characteristics of vectored axisymmetric ejector nozzles

    Lamb, Milton


    A series of vectoring axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at NASA-Langley Research Center. These ejector nozzles used convergent-divergent nozzles as the primary nozzles. The model geometric variables investigated were primary nozzle throat area, primary nozzle expansion ratio, effective ejector expansion ratio (ratio of shroud exit area to primary nozzle throat area), ratio of minimum ejector area to primary nozzle throat area, ratio of ejector upper slot height to lower slot height (measured on the vertical centerline), and thrust vector angle. The primary nozzle pressure ratio was varied from 2.0 to 10.0 depending upon primary nozzle throat area. The corrected ejector-to-primary nozzle weight-flow ratio was varied from 0 (no secondary flow) to approximately 0.21 (21 percent of primary weight-flow rate) depending on ejector nozzle configuration. In addition to the internal performance and pumping characteristics, static pressures were obtained on the shroud walls.

  12. Alternate Propellant Thermal Rocket Project

    National Aeronautics and Space Administration — The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by...

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


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

    Bakkiyaraj, R.


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

  15. Aggregate breakup in a contracting nozzle.

    Soos, Miroslav; Ehrl, Lyonel; Bäbler, Matthäus U; Morbidelli, Massimo


    The breakup of dense aggregates in an extensional flow was investigated experimentally. The flow was realized by pumping the suspension containing the aggregates through a contracting nozzle. Variation of the cluster mass distribution during the breakage process was measured by small-angle light scattering. Because of the large size of primary particles and the dense aggregate structure image analysis was used to determine the shape and structure of the produced fragments. It was found, that neither aggregate structure, characterized by a fractal dimension d(f) = 2.7, nor shape, characterized by an average aspect ratio equal to 1.5, was affected by breakage. Several passes through the nozzle were required to reach the steady state. This is explained by the radial variation of the hydrodynamic stresses at the nozzle entrance, characterized through computational fluid dynamics, which implies that only the fraction of aggregates whose strength is smaller than the local hydrodynamic stress is broken during one pass through the nozzle. Scaling of the steady-state aggregate size as a function of the hydrodynamic stress was used to determine the aggregate strength.

  16. Orbiter Water Dump Nozzles Redesign Lessons Learned

    Rotter, Hank


    Hank Rotter, NASA Technical Fellow for Environmental Control and Life Support System, will provide the causes and lessons learned for the two Space Shuttle Orbiter water dump icicles that formed on the side of the Orbiter. He will present the root causes and the criticality of these icicles, along with the redesign of the water dump nozzles and lessons learned during the redesign phase.

  17. New atomization nozzle for spray drying

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


    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.

  18. Microalgal cell disruption via ultrasonic nozzle spraying.

    Wang, M; Yuan, W


    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.

  19. New atomization nozzle for spray drying

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


    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.

  20. Fabrication of Microglass Nozzle for Microdroplet Jetting

    Dan Xie


    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.

  1. Not just rocket science

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


    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.

  2. Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode

    Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.


    The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations

  3. Experimental and CFD analysis of nozzle position of subsonic ejector

    Xilai ZHANG; Shiping JIN; Suyi HUANG; Guoqing TIAN


    The influence of nozzle position on the performance of an ejector was analyzed qualitatively with free jet flow model. Experimental investigations and computational fluid dynamics (CFD) analysis of the nozzle position of the subsonic ejector were also conducted. The results show that there is an optimum nozzle position for the ejector. The ejecting coefficient reaches its maximum when the nozzle is positioned at the optimum and decreases when deviating. Moreover, the nozzle position of an ejector is not a fixed value, but is influenced greatly by the flow parameters. Considering the complexity of the ejector, CFD is reckoned as a useful tool in the design of ejectors.

  4. Computational simulation of liquid fuel rocket injectors

    Landrum, D. Brian


    A major component of any liquid propellant rocket is the propellant injection system. Issues of interest include the degree of liquid vaporization and its impact on the combustion process, the pressure and temperature fields in the combustion chamber, and the cooling of the injector face and chamber walls. The Finite Difference Navier-Stokes (FDNS) code is a primary computational tool used in the MSFC Computational Fluid Dynamics Branch. The branch has dedicated a significant amount of resources to development of this code for prediction of both liquid and solid fuel rocket performance. The FDNS code is currently being upgraded to include the capability to model liquid/gas multi-phase flows for fuel injection simulation. An important aspect of this effort is benchmarking the code capabilities to predict existing experimental injection data. The objective of this MSFC/ASEE Summer Faculty Fellowship term was to evaluate the capabilities of the modified FDNS code to predict flow fields with liquid injection. Comparisons were made between code predictions and existing experimental data. A significant portion of the effort included a search for appropriate validation data. Also, code simulation deficiencies were identified.

  5. Standard Molded Composite Rocket Pyrogen Igniter - A progress report

    Lucy, M. H.


    The pyrogen igniter has the function to furnish a controlled, high temperature, high pressure gas to ignite solid propellant surfaces in a rocket motor. Present pyrogens consist of numerous inert components. The Standard Molded Pyrogen Igniter (SMPI) consists of three basic parts, a cap with several integrally molded features, an ignition pellet retainer plate, and a tube with additional integrally molded features. A description is presented of an investigation which indicates that the SMPI concept is a viable approach to the design and manufacture of pyrogen igniters for solid propellant rocket motors. For some applications, combining the structural and thermal properties of molded composites can result in the manufacture of lighter assemblies at considerable cost reduction. It is demonstrated that high strength, thin walled tubes with high length to diameter ratios can be fabricated from reinforced plastic molding compound using the displacement compression process.

  6. Fluid Flow in Continuous Casting Mold with a Configured Nozzle

    王镭; 沈厚发; 柳百成


    The influence of a configured nozzle on the turbulent fluid flow in a continuous casting mold was investigated using the simulation program Visual Cast, which used the finite difference method and the SIMPLER algorithm. CAD software was used to construct the complicated nozzle in the calculational region. The simulation accuracy was validated by comparison with the classic driven cavity flow problem. The simulation results agree well with water modeling experiments. The simulations show that the velocity distribution at the nozzle port is uneven and the jet faces downward more than the nozzle outlet. Simulations with a configured nozzle and the inlet velocity at the nozzle entrance give precise results and overcome the traditional difficulty in determining the nozzle outlet velocity.

  7. An open cycle gas core fusion rocket for space exploration

    Kammash, T.; Godfrey, T.

    A nuclear propulsion system that utilizes fusion reactions to heat a plasma in a magnetically confined device is examined as a potential rocket. It makes use of a high density plasma in a magnetic mirror geometry with a collision mean free path much shorter than its length. Under these conditions the plasma behaves like a fluid with confinement properties dictated by gasdynamic laws. Accordingly, the plasma escape from the device is analogous to the flow of a gas into vacuum from a vessel with a hole. Such a system is capable of producing a very high specific impulse albeit at modest thrust. One approach for enhancing the thrust is to use an auxiliary hydrogen propellant that could be regeneratively heated before it is introduced into the reactor chamber. As is flows past the fusion plasma it will be further heated by the radiation (bremsstrahlung and synchrotron) emanating from the plasma, and upon emergence from the nozzle it will generate the desired thrust. The system thus functions much like an open cycle gas core rocket with very attractive propulsive capabilities. In this paper we present the underlying physics principles of such a concept and assess its capability by applying the results to a round trip mission to Mars. It is shown that the propulsion parameters exceed those of a gas core fission reactor and without many of major hydrodynamic problems confronted by the latter.

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

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


    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.

  9. The mixing of solid propellant by an artificial muscle actuator

    岩崎, 祥大; 伴, 遼介; 吉浜, 舜; 中村, 太郎; 羽生, 宏人; Iwasaki, Akihiro; Ban, Ryosuke; Yoshihama, Shun; Nakamura, Taro; Habu, Hiroto


    This research aims to reduce the cost of the solid rocket motor production, mainly solid propellant. The production process of the solid rocket propellant are usually employed the multi-batch mixing. However, this study using a peristaltic pump as a mixer will lead to the continuous process. The pump system can mix the powder materials for propellant and we consider that it will make the slurry of the solid propellant efficiently by the mechanism of the fluid dynamics in the pump.

  10. Study of Cavitation/Vaporization in Liquid Rocket Thruster Injectors


    Caveny, L. H., and Summerfield, M., Aluminized Solid Propellants Burning in a Rocket Motor Flowfield, AIAA Journal, Vol. 16, No. 7, 1978, pp. 736-739. [2...the swirl chamber, and the pulsator and manifold are made of 304 stainless steel . Figure 1: Nomenclature defined for swirl injector. Table 1...Wayne, NJ, 2009 [17] MATLAB, Matrix Laboratory, Software Package, R2009a, The MathWorks, Natick, MA, 2009. [18] Coleman, H.W. and Steele , W.G

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

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


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

  12. Rocket Engine Innovations Advance Clean Energy


    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.

  13. Integrated model of a composite propellant rocket

    Miccio, Francesco


    The combustion of composite solid propellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

  14. Additive Manufacturing a Liquid Hydrogen Rocket Engine

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


    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.

  15. Identification of Noise Sources During Rocket Engine Test Firings and a Rocket Launch Using a Microphone Phased-Array

    Panda, Jayanta; Mosher, Robert N.; Porter, Barry J.


    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.

  16. Dynamic mechanical analysis of double base rocket propellants

    Marcin Cegła


    Full Text Available The article presents dynamic mechanical analysis (DMA for solid rocket propellants testing. Principles of operation and measured values are briefly described. The authors refer to the previous research of PTFE material and literature data providing information about proper experimental conditions and influence of measurement frequency, load amplitude, and heating rate on the results of DMA tests. The experimental results of solid double-base rocket propellant testing obtained on the N Netzsch DMA 242 device are presented. Mechanical properties such as the dynamic storage modulus E´, the dynamic loss modulus E˝ and tan(δ were measured within temperature range from (–120°C to (+90°C at the heating rate of 1 K/min. The test sample was subjected to a dual cantilever multi-frequency test. Special attention was paid to determination of the glass transition temperature of the tested propellant in reference to the NATO standardization agreement 4540 as well as influence of the measurement frequency on the glass transition.[b]Keywords[/b]: Dynamic mechanical analysis, solid rocket propellants, glass transition temperature

  17. Solid Propellant Grain Structural Integrity Analysis


    The structural properties of solid propellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.

  18. Theoretical prediction of regression rates in swirl-injection hybrid rocket engines

    Ozawa, K.; Shimada, T.


    The authors theoretically and analytically predict what times regression rates of swirl injection hybrid rocket engines increase higher than the axial injection ones by estimating heat flux from boundary layer combustion to the fuel port. The schematic of engines is assumed as ones whose oxidizer is injected from the opposite side of the nozzle such as ones of Yuasa et al. propose. To simplify the estimation, we assume some hypotheses such as three-dimensional (3D) axisymmetric flows have been assumed. The results of this prediction method are largely consistent with Yuasa's experiments data in the range of high swirl numbers.

  19. Computer Modeling of a Rotating Detonation Engine in a Rocket Configuration


    coefficient CP Specific heat capacity at constant pressure ( J kg−K ) CS Nozzle stream thrust coefficient D Detonation wave speed in laboratory frame-of...greater than the detonation fuel-to-air ratio, the ratio of specific heats and gas constant at station c3.4 are calculated using Eq. 75 and Eq. 76...COMPUTER MODELING OF A ROTATING DETONATION ENGINE IN A ROCKET CONFIGURATION THESIS Nihar N. Shah, 1st Lt, USAF AFIT-ENY-MS-15-M-230 DEPARTMENT OF THE

  20. Rocket Science 101 Interactive Educational Program

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald


    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.