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

Science.gov (United States)

Culick, F. E. C.

1972-01-01

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

Combustion response to acoustic perturbation in liquid rocket engines

Science.gov (United States)

Ghafourian, Akbar

An experimental study of the effect of acoustic perturbations on combustion behavior of a model liquid propellant rocket engine has been carried out. A pair of compression drivers were used to excite transverse and longitudinal acoustic fields at strengths of up to 156.6 dB and 159.5 dB respectively in the combustion chamber of the experimental rocket engine. Propellant simulants were injected into the combustion chamber through a single element shear coaxial injector. Water and air were used in cold flow studies and ethanol and oxygen-enriched air were used as fuel and oxidizer in reacting hot flow studies. In cold flow studies an imposed transverse acoustic field had a more pronounced effect on the spray pattern than a longitudinal acoustic fields. A transverse acoustic field widened the spray by as much as 33 percent and the plane of impingement of the spray with chamber walls moved up closer to the injection plane. The behavior was strongly influenced by the gas phase velocity but was less sensitive to changes in the liquid phase velocity. In reacting hot flow studies the effects of changes in equivalence ratio, excitation amplitude, excitation frequency, liquid and gas phase velocity and chamber pressure on the response of the injector to imposed high frequency transverse acoustic excitation were measured. Reducing the equivalence ratio from 7.4 to 3.8 increased the chamber pressure response to the imposed excitation at 3000 Hz. Increasing the excitation amplitude from 147 dB to 155.6 dB at 3000 Hz increased the chamber pressure response to the excitation. In the frequency range of 1240 Hz to 3220 Hz, an excitation frequency of 3000 Hz resulted in the largest response of the chamber pressure indicating the importance of fluid dynamic coupling. Increasing the liquid phase velocity from 9.2 m/sec to 22.7 m/sec, did not change the amplitude of the chamber pressure response to excitation. This implied the importance of local equivalence ratio and not the overall

Staged combustion with piston engine and turbine engine supercharger

Science.gov (United States)

Fischer, Larry E [Los Gatos, CA; Anderson, Brian L [Lodi, CA; O'Brien, Kevin C [San Ramon, CA

2011-11-01

A combustion engine method and system provides increased fuel efficiency and reduces polluting exhaust emissions by burning fuel in a two-stage combustion system. Fuel is combusted in a piston engine in a first stage producing piston engine exhaust gases. Fuel contained in the piston engine exhaust gases is combusted in a second stage turbine engine. Turbine engine exhaust gases are used to supercharge the piston engine.

Advanced Materials and Manufacturing for Low-Cost, High-Performance Liquid Rocket Combustion Chambers, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — Silicided niobium alloy (C103) combustion chambers have been used extensively in both NASA and DoD liquid rocket propulsion systems. Niobium alloys offer a good...

Experimental investigation of solid rocket motors for small sounding rockets

Science.gov (United States)

Suksila, Thada

2018-01-01

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

Design considerations for a pressure-driven multi-stage rocket

Science.gov (United States)

Sauerwein, Steven Craig

2002-01-01

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

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

OpenAIRE

David Greatrix

2015-01-01

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

Advances for laser ignition of internal combustion and rocket engines

International Nuclear Information System (INIS)

Schwarz, E.

2011-01-01

The scope of the PhD thesis presented here is the investigation of theoretical and practical aspects of laser-induced spark ignition and laser thermal ignition. Laser ignition systems are currently undergoing a rapidly development with growing intensity involving more and more research groups who mainly concentrate on the field of car and large combustion engines. This research is primarily driven by the engagement to meet the increasingly strict emission limits and by the intention to use the limited energy reserves more efficiently. For internal combustion engines, laser plasma-induced ignition will allow to combine the goals for legally required reductions of pollutant emissions and higher engine efficiencies. Also for rocket engines laser ignition turns out to be very attractive. A highly reliable ignition system like laser ignition would represent an option for introducing non-toxic propellants in order to replace highly toxic and carcinogenic hydrazine-based propellants commonly used in launch vehicle upper stages and satellites. The most important results on laser ignition and laser plasma generation, accomplished by the author and, in some respects, enriched by cooperation with colleagues are presented in the following. The emphasis of this thesis is placed on the following issues: - Two-color effects on laser plasma generation - Theoretical considerations about the focal volume concerning plasma generation - Plasma transmission experiments - Ignition experiments on laser-induced ignition - Ignition experiments on thermally-induced ignition - Feasibility study on laser ignition of rocket engines The purpose of the two-color laser plasma experiments is to investigate possible constructive interference effects of driving fields that are not monochromatic, but contain (second) harmonic radiation with respect to the goal of lowering the plasma generation threshold. Such effects have been found in a number of related processes, such as laser ablation or high

Development and Hot-fire Testing of Additively Manufactured Copper Combustion Chambers for Liquid Rocket Engine Applications

Science.gov (United States)

Gradl, Paul R.; Greene, Sandy Elam; Protz, Christopher S.; Ellis, David L.; Lerch, Bradley A.; Locci, Ivan E.

2017-01-01

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket

Science.gov (United States)

Dawson, Virginia P.; Bowles, Mark D.

2004-01-01

The Centaur is one of the most powerful rockets in the world. As an upper-stage rocket for the Atlas and Titan boosters it has been a reliable workhorse for NASA for over forty years and has played an essential role in many of NASA's adventures into space. In this CD-ROM you will be able to explore the Centaur's history in various rooms to this virtual museum. Visit the "Movie Theater" to enjoy several video documentaries on the Centaur. Enter the "Interview Booth" to hear and read interviews with scientists and engineers closely responsible for building and operating the rocket. Go to the "Photo Gallery" to look at numerous photos of the rocket throughout its history. Wander into the "Centaur Library" to read various primary documents of the Centaur program. Finally, stop by the "Observation Deck" to watch a virtual Centaur in flight.

Performance of a RBCC Engine in Rocket-Operation

Science.gov (United States)

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

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

Development and Hotfire Testing of Additively Manufactured Copper Combustion Chambers for Liquid Rocket Engine Applications

Science.gov (United States)

Gradl, Paul R.; Greene, Sandy; Protz, Chris

2017-01-01

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA’s Marshall Space Flight Center (MSFC) has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. MSFC’s efforts include a 4,000 pounds-force thrust liquid oxygen/methane (LOX/CH4) combustion chamber. Small thrust chambers for 1,200 pounds-force LOX/hydrogen (H2) applications have also been designed and fabricated with SLM GRCop-84. Similar chambers have also completed development with an Inconel 625 jacket bonded to the GRCop-84 material, evaluating direct metal deposition (DMD) laser- and arc-based techniques. The same technologies for these lower thrust applications are being applied to 25,000-35,000 pounds-force main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

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

Directory of Open Access Journals (Sweden)

David R. Greatrix

2009-03-01

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

Modeling of Uneven Flow and Electromagnetic Field Parameters in the Combustion Chamber of Liquid Rocket Engine with a Near-wall Layer Available

Directory of Open Access Journals (Sweden)

A. V. Rudinskii

2015-01-01

Full Text Available The paper concerns modeling of an uneven flow and electromagnetic field parameters in the combustion chamber of the liquid rocket engine with a near-wall layer available.The research objective was to evaluate quantitatively influence of changing model chamber mode of the liquid rocket engine on the electro-physical characteristics of the hydrocarbon fuel combustion by-products.The main method of research was based on development of a final element model of the flowing path of the rocket engine chamber and its adaptation to the boundary conditions.The paper presents a developed two-dimensional non-stationary mathematical model of electro-physical processes in the liquid rocket engine chamber using hydrocarbon fuel. The model takes into consideration the features of a gas-dynamic contour of the engine chamber and property of thermo-gas-dynamic characteristics of the ionized products of combustion of hydrocarbonic fuel. Distributions of magnetic field intensity and electric conductivity received and analyzed taking into account a low-temperature near-wall layer. Special attention is paid to comparison of obtained calculation values of the electric current, which is taken out from intrachamber space of the engine with earlier published data of other authors.

Numerical investigations of hybrid rocket engines

Science.gov (United States)

Betelin, V. B.; Kushnirenko, A. G.; Smirnov, N. N.; Nikitin, V. F.; Tyurenkova, V. V.; Stamov, L. I.

2018-03-01

Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction.

Robust Low Cost Liquid Rocket Combustion Chamber by Advanced Vacuum Plasma Process

Science.gov (United States)

Holmes, Richard; Elam, Sandra; Ellis, David L.; McKechnie, Timothy; Hickman, Robert; Rose, M. Franklin (Technical Monitor)

2001-01-01

Next-generation, regeneratively cooled rocket engines will require materials that can withstand high temperatures while retaining high thermal conductivity. Fabrication techniques must be cost efficient so that engine components can be manufactured within the constraints of shrinking budgets. Three technologies have been combined to produce an advanced liquid rocket engine combustion chamber at NASA-Marshall Space Flight Center (MSFC) using relatively low-cost, vacuum-plasma-spray (VPS) techniques. Copper alloy NARloy-Z was replaced with a new high performance Cu-8Cr-4Nb alloy developed by NASA-Glenn Research Center (GRC), which possesses excellent high-temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability. Functional gradient technology, developed building composite cartridges for space furnaces was incorporated to add oxidation resistant and thermal barrier coatings as an integral part of the hot wall of the liner during the VPS process. NiCrAlY, utilized to produce durable protective coating for the space shuttle high pressure fuel turbopump (BPFTP) turbine blades, was used as the functional gradient material coating (FGM). The FGM not only serves as a protection from oxidation or blanching, the main cause of engine failure, but also serves as a thermal barrier because of its lower thermal conductivity, reducing the temperature of the combustion liner 200 F, from 1000 F to 800 F producing longer life. The objective of this program was to develop and demonstrate the technology to fabricate high-performance, robust, inexpensive combustion chambers for advanced propulsion systems (such as Lockheed-Martin's VentureStar and NASA's Reusable Launch Vehicle, RLV) using the low-cost VPS process. VPS formed combustion chamber test articles have been formed with the FGM hot wall built in and hot fire tested, demonstrating for the first time a coating that will remain intact through the hot firing test, and with

Combustion and Magnetohydrodynamic Processes in Advanced Pulse Detonation Rocket Engines

Science.gov (United States)

Cole, Lord Kahil

A number of promising alternative rocket propulsion concepts have been developed over the past two decades that take advantage of unsteady combustion waves in order to produce thrust. These concepts include the Pulse Detonation Rocket Engine (PDRE), in which repetitive ignition, propagation, and reflection of detonations and shocks can create a high pressure chamber from which gases may be exhausted in a controlled manner. The Pulse Detonation Rocket Induced Magnetohydrodynamic Ejector (PDRIME) is a modification of the basic PDRE concept, developed by Cambier (1998), which has the potential for performance improvements based on magnetohydrodynamic (MHD) thrust augmentation. The PDRIME has the advantage of both low combustion chamber seeding pressure, per the PDRE concept, and efficient energy distribution in the system, per the rocket-induced MHD ejector (RIME) concept of Cole, et al. (1995). In the initial part of this thesis, we explore flow and performance characteristics of different configurations of the PDRIME, assuming quasi-one-dimensional transient flow and global representations of the effects of MHD phenomena on the gas dynamics. By utilizing high-order accurate solvers, we thus are able to investigate the fundamental physical processes associated with the PDRIME and PDRE concepts and identify potentially promising operating regimes. In the second part of this investigation, the detailed coupling of detonations and electric and magnetic fields are explored. First, a one-dimensional spark-ignited detonation with complex reaction kinetics is fully evaluated and the mechanisms for the different instabilities are analyzed. It is found that complex kinetics in addition to sufficient spatial resolution are required to be able to quantify high frequency as well as low frequency detonation instability modes. Armed with this quantitative understanding, we then examine the interaction of a propagating detonation and the applied MHD, both in one-dimensional and two

Emission and combustion characteristics of multiple stage diesel combustion; Nidan nensho ni yoru diesel kikan no nensho to haishutsubutsu tokusei

Energy Technology Data Exchange (ETDEWEB)

Hashizume, T; Miyamoto, T; Tsujimura, K [New A.C.E. Institute Co. Ltd., Tokyo (Japan); Kobayashi, S; Shimizu, K [Japan Automobile Research Institute, Tsukuba (Japan)

1997-10-01

A new concept of multiple stage diesel combustion was studied by means of engine test, combustion observation and numerical simulation, in order to reduce NOx emissions at high load conditions. With this concept, the premixed combustion occurs under the fuel lean conditions and the diffusion combustion occurs under the high temperature conditions. As seen in the result of combustion observation, a first stage combustion occurs with no luminous flame. A second stage combustion occurs with a luminous flame after very short ignition delay period. However the luminous flame is disappeared immediately. Because cylinder temperature is high, and hence soot oxidizes immediately. 5 refs., 11 figs., 1 tab.

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

Science.gov (United States)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

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

Experimental and numerical studies on two-stage combustion of biomass

Energy Technology Data Exchange (ETDEWEB)

Houshfar, Eshan

2012-07-01

In this thesis, two-stage combustion of biomass was experimentally/numerically investigated in a multifuel reactor. The following emissions issues have been the main focus of the work: 1- NOx and N2O 2- Unburnt species (CO and CxHy) 3- Corrosion related emissions.The study had a focus on two-stage combustion in order to reduce pollutant emissions (primarily NOx emissions). It is well known that pollutant emissions are very dependent on the process conditions such as temperature, reactant concentrations and residence times. On the other hand, emissions are also dependent on the fuel properties (moisture content, volatiles, alkali content, etc.). A detailed study of the important parameters with suitable biomass fuels in order to optimize the various process conditions was performed. Different experimental studies were carried out on biomass fuels in order to study the effect of fuel properties and combustion parameters on pollutant emissions. Process conditions typical for biomass combustion processes were studied. Advanced experimental equipment was used in these studies. The experiments showed the effects of staged air combustion, compared to non-staged combustion, on the emission levels clearly. A NOx reduction of up to 85% was reached with staged air combustion using demolition wood as fuel. An optimum primary excess air ratio of 0.8-0.95 was found as a minimizing parameter for the NOx emissions for staged air combustion. Air staging had, however, a negative effect on N2O emissions. Even though the trends showed a very small reduction in the NOx level as temperature increased for non-staged combustion, the effect of temperature was not significant for NOx and CxHy, neither in staged air combustion or non-staged combustion, while it had a great influence on the N2O and CO emissions, with decreasing levels with increasing temperature. Furthermore, flue gas recirculation (FGR) was used in combination with staged combustion to obtain an enhanced NOx reduction. The

Design and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

Science.gov (United States)

Garcia, C. P.; Medina, C. R.; Protz, C. S.; Kenny, R. J.; Kelly, G. W.; Casiano, M. J.; Hulka, J. R.; Richardson, B. R.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural

Combustion Stability Assessments of the Black Brant Solid Rocket Motor

Science.gov (United States)

Fischbach, Sean

2014-01-01

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

An X-ray Experiment with Two-Stage Korean Sounding Rocket

Directory of Open Access Journals (Sweden)

Uk-Won Nam

1998-12-01

Full Text Available The test result of the X-ray observation system is presented which have been developed at Korea Astronomy Observatory for 3 years (1995-1997. The instrument, which is composed of detector and signal processing parts, is designed for the future observations of compact X-ray sources. The performance of the instrument was tested by mounting on the two-stage Korean Sounding Rocket, which was launched from Taean rocket flight center on June 11 at 10:00 KST 1998. Telemetry data were received from individual parts of the instrument for 32 and 55.7 sec, respectively, since the launch of the rocket. In this paper, the result of the data analysis based on the telemetry data and discussion about the performance of the instrument is reported.

High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner For Advanced Rocket Engines

Science.gov (United States)

Bhat, Biliyar N.; Ellis, David; Singh, Jogender

2014-01-01

Advanced high thermal conductivity materials research conducted at NASA Marshall Space Flight Center (MSFC) with state of the art combustion chamber liner material NARloy-Z showed that its thermal conductivity can be increased significantly by adding diamond particles and sintering it at high temperatures. For instance, NARloy-Z containing 40 vol. percent diamond particles, sintered at 975C to full density by using the Field assisted Sintering Technology (FAST) showed 69 percent higher thermal conductivity than baseline NARloy-Z. Furthermore, NARloy-Z-40vol. percent D is 30 percent lighter than NARloy-Z and hence the density normalized thermal conductivity is 140 percent better. These attributes will improve the performance and life of the advanced rocket engines significantly. By one estimate, increased thermal conductivity will directly translate into increased turbopump power up to 2X and increased chamber pressure for improved thrust and ISP, resulting in an expected 20 percent improvement in engine performance. Follow on research is now being conducted to demonstrate the benefits of this high thermal conductivity NARloy-Z-D composite for combustion chamber liner applications in advanced rocket engines. The work consists of a) Optimizing the chemistry and heat treatment for NARloy-Z-D composite, b) Developing design properties (thermal and mechanical) for the optimized NARloy-Z-D, c) Fabrication of net shape subscale combustion chamber liner, and d) Hot fire testing of the liner for performance. FAST is used for consolidating and sintering NARlo-Z-D. The subscale cylindrical liner with built in channels for coolant flow is also fabricated near net shape using the FAST process. The liner will be assembled into a test rig and hot fire tested in the MSFC test facility to determine performance. This paper describes the development of this novel high thermal conductivity NARloy-Z-D composite material, and the advanced net shape technology to fabricate the combustion

On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

Science.gov (United States)

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

2004-01-01

All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

Staged fluidized-bed combustion and filter system

International Nuclear Information System (INIS)

Mei, J.S.; Halow, J.S.

1994-01-01

A staged fluidized-bed combustion and filter system are described for substantially reducing the quantity of waste through the complete combustion into ash-type solids and gaseous products. The device has two fluidized-bed portions, the first primarily as a combustor/pyrolyzer bed, and the second as a combustor/filter bed. The two portions each have internal baffles to define stages so that material moving therein as fluidized beds travel in an extended route through those stages. Fluidization and movement is achieved by the introduction of gases into each stage through a directional nozzle. Gases produced in the combustor/pyrolyzer bed are permitted to travel into corresponding stages of the combustor/filter bed through screen filters that permit gas flow but inhibit solids flow. Any catalyst used in the combustor/filter bed is recycled. The two beds share a common wall to minimize total volume of the system. A slightly modified embodiment can be used for hot gas desulfurization and sorbent regeneration. Either side-by-side rectangular beds or concentric beds can be used. The system is particularly suited to the processing of radioactive and chemically hazardous waste. 10 figures

Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

Science.gov (United States)

Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

2016-01-01

This paper describes the process development for fabricating a high thermal conductivity NARloy-Z-Diamond composite (NARloy-Z-D) combustion chamber liner for application in advanced rocket engines. The fabrication process is challenging and this paper presents some details of these challenges and approaches used to address them. Prior research conducted at NASA-MSFC and Penn State had shown that NARloy-Z-40%D composite material has significantly higher thermal conductivity than the state of the art NARloy-Z alloy. Furthermore, NARloy-Z-40 %D is much lighter than NARloy-Z. These attributes help to improve the performance of the advanced rocket engines. Increased thermal conductivity will directly translate into increased turbopump power, increased chamber pressure for improved thrust and specific impulse. Early work on NARloy-Z-D composites used the Field Assisted Sintering Technology (FAST, Ref. 1, 2) for fabricating discs. NARloy-Z-D composites containing 10, 20 and 40vol% of high thermal conductivity diamond powder were investigated. Thermal conductivity (TC) data. TC increased with increasing diamond content and showed 50% improvement over pure copper at 40vol% diamond. This composition was selected for fabricating the combustion chamber liner using the FAST technique.

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

Energy Technology Data Exchange (ETDEWEB)

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

2007-07-01

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

The issue of ensuring the safe explosion of the spent orbital stages of a launch vehicle with propulsion rocket engine

Directory of Open Access Journals (Sweden)

Trushlyakov Valeriy I.

2017-01-01

Full Text Available A method for increasing the safe explosion of the spent orbital stages of a space launch vehicle (SLV with a propulsion rocket engine (PRE based on the gasification of unusable residues propellant and venting fuel tanks. For gasification and ventilation the hot gases used produced by combustion of the specially selected gas generating composition (GGC with a set of physical and chemical properties. Excluding the freezing of the drainage system on reset gasified products (residues propellant+pressurization gas+hot gases in the near-Earth space is achieved by selecting the physical-chemical characteristics of the GGC. Proposed steps to ensure rotation of gasified products due to dumping through the drainage system to ensure the most favorable conditions for propellant gasification residues. For example, a tank with liquid oxygen stays with the orbital spent second stage of the SLV “Zenit”, which shows the effectiveness of the proposed method.

Additively Manufactured Low Cost Upper Stage Combustion Chamber

Science.gov (United States)

Protz, Christopher; Cooper, Ken; Ellis, David; Fikes, John; Jones, Zachary; Kim, Tony; Medina, Cory; Taminger, Karen; Willingham, Derek

2016-01-01

Over the past two years NASA's Low Cost Upper Stage Propulsion (LCUSP) project has developed Additive Manufacturing (AM) technologies and design tools aimed at reducing the costs and manufacturing time of regeneratively cooled rocket engine components. High pressure/high temperature combustion chambers and nozzles must be regeneratively cooled to survive their operating environment, causing their design fabrication to be costly and time consuming due to the number of individual steps and different processes required. Under LCUSP, AM technologies in Sintered Laser Melting (SLM) GRCop-84 and Electron Beam Freeform Fabrication (EBF3) Inconel 625 have been significantly advanced, allowing the team to successfully fabricate a 25k-class regenerative chamber. Estimates of the costs and schedule of future builds indicate cost reductions and significant schedule reductions will be enabled by this technology. Characterization of the microstructural and mechanical properties of the SLM-produced GRCop-84, EBF3 Inconel 625 and the interface layer between the two has been performed and indicates the properties will meet the design requirements. The LCUSP chamber is to be tested with a previously demonstrated SLM injector in order to advance the Technology Readiness Level (TRL) and demonstrate the capability of the application of these processes. NASA is advancing these technologies to reduce cost and schedule for future engine applications and commercial needs.

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

OpenAIRE

S. Saha; D. Chakraborty

2016-01-01

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

Performances Study of a Hybrid Rocket Engine

Directory of Open Access Journals (Sweden)

Adrian-Nicolae BUTURACHE

2018-06-01

Full Text Available This paper presents a study which analyses the functioning and performances optimization of a hybrid rocket engine based on gaseous oxygen and polybutadiene polymer (HTPB. Calculations were performed with NASA CEA software in order to obtain the parameters resulted following the combustion process. Using these parameters, the main parameters of the hybrid rocket engine were optimized. Using the calculus previously stated, an experimental rocket engine producing 100 N of thrust was pre-dimensioned, followed by an optimization of the rocket engine as a function of several parameters. Having the geometry and the main parameters of the hybrid rocket engine combustion process, numerical simulations were performed in the CFX – ANSYS commercial software, which allowed visualizing the flow field and the jet expansion. Finally, the analytical calculus was validated through numerical simulations.

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

Science.gov (United States)

Ferragut, N. J.

1976-01-01

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

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

Directory of Open Access Journals (Sweden)

David Greatrix

2015-02-01

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

Combustion characteristics of the LO2/GCH4 fuel-rich preburners for staged combustion cycle rocket engines

Science.gov (United States)

Ono, Fumiei; Tamura, Hiroshi; Sakamoto, Hiroshi; Sasaki, Masaki

1991-09-01

The combustion characteristics of Liquid Oxygen (LO2)/Gaseous Methane (GCH4) fuel rich preburners were experimentally studied using subscale hardware. Three types of preburners with coaxial type propellant injection elements were designed and fabricated, and were used for hot fire testing. LO2 was used as oxidizer, and GCH4 at room temperature was used as fuel. The tests were conducted at chamber pressures ranging from 6.7 to 11.9 M Pa, and oxidizer to fuel ratios ranged from 0.16 to 0.42. The test results, which include combustion gas temperature T(sub c), characteristic velocity C(sup *) and soot adhesion data, are presented. The T(sub c) efficiency and the C(sup *) efficiency were found to be a function of oxidizer to fuel ratio and chamber pressure. These efficiencies are correlated by an empirical correlation parameter which accounts for the effects of oxidizer to fuel ratio and chamber pressure. The exhaust plumes were colorless and transparent under all tests conditions. There was some soot adhesion to the chamber wall, but no soot adhesion was observed on the main injector simulator orifices. Higher temperature igniter gas was required to ignite the main propellants of the preburner compared with that of the LO2/Gaseous Hydrogen (GH2) propellants combination.

Ash related behaviour in staged and non-staged combustion of biomass fuels and fuel mixtures

International Nuclear Information System (INIS)

Becidan, Michaël; Todorovic, Dusan; Skreiberg, Øyvind; Khalil, Roger A.; Backman, Rainer; Goile, Franziska; Skreiberg, Alexandra; Jovovic, Aleksandar; Sørum, Lars

2012-01-01

The fate of selected elements (with focus on the important players in corrosion i.e. Na, K, Pb, Zn, Cl and S) are investigated for three biomasses (wood, demolition wood and coffee waste) and six mixtures of these as pellets both with and without air staging in a laboratory reactor. In order to get a complete overview of the combustion products, both online and offline analytical methods are used. Information is collected about: flue gas composition, particle (fly ash) size distribution and composition, bottom ash composition and melting properties. The main findings are: (1) complex interactions are taking place between the mixed fuels during combustion; (2) the mode of occurrence of an element as well as the overall structure of the fuel are important for speciation; (3) the pelletisation process, by bringing chemical elements into intimate contact, may affect partitioning and speciation; (4) staging and mixing might simultaneously have positive and negative effects on operation; (5) staging affects the governing mechanisms of fly ash (aerosols) formation. -- Highlights: ► Complex interactions are taking place between the mixed fuels during combustion. ► The mode of occurrence of an element as well as the overall structure of the fuel are important for speciation. ► The pelletisation process, by bringing chemical elements into intimate contact, may affect partitioning and speciation. ► Staging and mixing might simultaneously have positive and negative effects on operation. ► Staging affects the governing mechanisms of fly ash (aerosols) formation.

Application of C/C composites to the combustion chamber of rocket engines. Part 1: Heating tests of C/C composites with high temperature combustion gases

Science.gov (United States)

Tadano, Makoto; Sato, Masahiro; Kuroda, Yukio; Kusaka, Kazuo; Ueda, Shuichi; Suemitsu, Takeshi; Hasegawa, Satoshi; Kude, Yukinori

1995-04-01

Carbon fiber reinforced carbon composite (C/C composite) has various superior properties, such as high specific strength, specific modulus, and fracture strength at high temperatures of more than 1800 K. Therefore, C/C composite is expected to be useful for many structural applications, such as combustion chambers of rocket engines and nose-cones of space-planes, but C/C composite lacks oxidation resistivity in high temperature environments. To meet the lifespan requirement for thermal barrier coatings, a ceramic coating has been employed in the hot-gas side wall. However, the main drawback to the use of C/C composite is the tendency for delamination to occur between the coating layer on the hot-gas side and the base materials on the cooling side during repeated thermal heating loads. To improve the thermal properties of the thermal barrier coating, five different types of 30-mm diameter C/C composite specimens constructed with functionally gradient materials (FGM's) and a modified matrix coating layer were fabricated. In this test, these specimens were exposed to the combustion gases of the rocket engine using nitrogen tetroxide (NTO) / monomethyl hydrazine (MMH) to evaluate the properties of thermal and erosive resistance on the thermal barrier coating after the heating test. It was observed that modified matrix and coating with FGM's are effective in improving the thermal properties of C/C composite.

Thermo-acoustic instabilities of high-frequency combustion in rocket engines; Instabilites thermo-acoustiques de combustion haute-frequence dans les moteurs fusees

Energy Technology Data Exchange (ETDEWEB)

Cheuret, F

2005-10-15

Rocket motors are confined environments where combustion occurs in extreme conditions. Combustion instabilities can occur at high frequencies; they are tied to the acoustic modes of the combustion chamber. A common research chamber, CRC, allows us to study the response of a turbulent two-phase flame to acoustic oscillations of low or high amplitudes. The chamber is characterised under cold conditions to obtain, in particular, the relative damping coefficient of acoustic oscillations. The structure and frequency of the modes are determined in the case where the chamber is coupled to a lateral cavity. We have used a powder gun to study the response to a forced acoustic excitation at high amplitude. The results guide us towards shorter flames. The injectors were then modified to study the combustion noise level as a function of injection conditions. The speed of the gas determines whether the flames are attached or lifted. The noise level of lifted flames is higher. That of attached flames is proportional to the Weber number. The shorter flames whose length is less than the radius of the CRC, necessary condition to obtain an effective coupling, are the most sensitive to acoustic perturbations. The use of a toothed wheel at different positions in the chamber allowed us to obtain informations on the origin of the thermo-acoustic coupling, main objective of this thesis. The flame is sensitive to pressure acoustic oscillations, with a quasi-zero response time. These observations suggest that under the conditions of the CRC, we observe essentially the response of chemical kinetics to pressure oscillations. (author)

Rocket + Science = Dialogue

Science.gov (United States)

Morris,Bruce; Sullivan, Greg; Burkey, Martin

2010-01-01

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

Smoke reduction using multiple stage diesel combustion; Nidan nensho ni yoru diesel kikan no smoke teigen koka

Energy Technology Data Exchange (ETDEWEB)

Tsujimura, K. [Chiba Institute of Technology, Chiba (Japan); Hashizume, T.; Miyamoto, T.; Akagawa, H.

1999-09-25

A new diesel combustion concept termed MULDIC (MULtiple stage Diesel Combustion), which can reduce NO{sub x} emissions at high load conditions, was studied by means of engine test, combustion observation and numerical simulation. The engine test showed that simultaneous reduction of NO{sub x} and smoke could be obtained with MULDIC operation. In-cylinder combustion observation was carried out in order to clarify the cause of the smoke reduction of the MULDIC. In these photographs, a first stage combustion sins observed with nonluminous flame. In a second stage combustion, a highly sooting luminous flame was observed because the ignition delay was short. However this luminous flame disappeared quickly. Computational results showed that, oxygen is distributed evenly in the cylinder in later portion of the second stage combustion. Moreover, since cylinder temperature was very high, most of the soot oxidized rapidly. Thus, low smoke combustion was obtained. (author)

Development of Explosive Ripper with Two-Stage Combustion

Science.gov (United States)

1974-10-01

inch pipe duct work, the width of this duct proved to be detrimental in marginally rippable material; the duct, instead of the penetrator tip, was...marginally rippable rock. ID. Operating Requirements 2. Fuel The two-stage combustion device is designed to operate using S A 42. the same diesel

Low Cost Upper Stage-Class Propulsion (LCUSP)

Science.gov (United States)

Vickers, John

2015-01-01

NASA is making space exploration more affordable and viable by developing and utilizing innovative manufacturing technologies. Technology development efforts at NASA in propulsion are committed to continuous innovation of design and manufacturing technologies for rocket engines in order to reduce the cost of NASA's journey to Mars. The Low Cost Upper Stage-Class Propulsion (LCUSP) effort will develop and utilize emerging Additive Manufacturing (AM) to significantly reduce the development time and cost for complex rocket propulsion hardware. Benefit of Additive Manufacturing (3-D Printing) Current rocket propulsion manufacturing techniques are costly and have lengthy development times. In order to fabricate rocket engines, numerous complex parts made of different materials are assembled in a way that allow the propellant to collect heat at the right places to drive the turbopump and simultaneously keep the thrust chamber from melting. The heat conditioned fuel and oxidizer come together and burn inside the combustion chamber to provide thrust. The efforts to make multiple parts precisely fit together and not leak after experiencing cryogenic temperatures on one-side and combustion temperatures on the other is quite challenging. Additive manufacturing has the potential to significantly reduce the time and cost of making rocket parts like the copper liner and Nickel-alloy jackets found in rocket combustion chambers where super-cold cryogenic propellants are heated and mixed to the extreme temperatures needed to propel rockets in space. The Selective Laser Melting (SLM) machine fuses 8,255 layers of copper powder to make a section of the chamber in 10 days. Machining an equivalent part and assembling it with welding and brazing techniques could take months to accomplish with potential failures or leaks that could require fixes. The design process is also enhanced since it does not require the 3D model to be converted to 2-D drawings. The design and fabrication process

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

Energy Technology Data Exchange (ETDEWEB)

Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

2015-06-15

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort.

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

International Nuclear Information System (INIS)

Labib, Satira; King, Jeffrey

2015-01-01

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort

Two-stage combustion for reducing pollutant emissions from gas turbine combustors

Science.gov (United States)

Clayton, R. M.; Lewis, D. H.

1981-01-01

Combustion and emission results are presented for a premix combustor fueled with admixtures of JP5 with neat H2 and of JP5 with simulated partial-oxidation product gas. The combustor was operated with inlet-air state conditions typical of cruise power for high performance aviation engines. Ultralow NOx, CO and HC emissions and extended lean burning limits were achieved simultaneously. Laboratory scale studies of the non-catalyzed rich-burning characteristics of several paraffin-series hydrocarbon fuels and of JP5 showed sooting limits at equivalence ratios of about 2.0 and that in order to achieve very rich sootless burning it is necessary to premix the reactants thoroughly and to use high levels of air preheat. The application of two-stage combustion for the reduction of fuel NOx was reviewed. An experimental combustor designed and constructed for two-stage combustion experiments is described.

Hybrid rocket motor testing at Nammo Raufoss A/S

Science.gov (United States)

Rønningen, Jan-Erik; Kubberud, Nils

2005-08-01

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

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

Science.gov (United States)

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

2016-10-01

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

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

Directory of Open Access Journals (Sweden)

Jichao Hu

2014-01-01

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

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

Directory of Open Access Journals (Sweden)

Kalomba Mboyi

2015-04-01

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

Combustion characteristics and influential factors of isooctane active-thermal atmosphere combustion assisted by two-stage reaction of n-heptane

Energy Technology Data Exchange (ETDEWEB)

Lu, Xingcai; Ji, Libin; Ma, Junjun; Zhou, Xiaoxin; Huang, Zhen [Key Lab. for Power Machinery and Engineering of MOE, Shanghai Jiao Tong University, 200240 Shanghai (China)

2011-02-15

This paper presents an experimental study on the isooctane active-thermal atmosphere combustion (ATAC) which is assisted by two-stage reaction of n-heptane. The active-thermal atmosphere is created by low- and high-temperature reactions of n-heptane which is injected at intake port, and isooctane is directly injected into combustion chamber near the top dead center. The effects of isooctane injection timing, active-thermal atmosphere intensity, overall equivalence ratio, and premixed ratio on combustion characteristics and emissions are investigated. The experimental results reveal that, the isooctane ignition and combustion can be classified to thermal atmosphere combustion, active atmosphere combustion, and active-thermal atmosphere combustion respectively according to the extent of n-heptane oxidation as well as effects of isooctane quenching and charge cooling. n-Heptane equivalence ratio, isooctane equivalence ratio and isooctane delivery advance angle are major control parameters. In one combustion cycle, the isooctane ignited and burned after those of n-heptane, and then this combustion phenomenon can also be named as dual-fuel sequential combustion (DFSC). The ignition timing of the overall combustion event is mainly determined by n-heptane equivalence ratio and can be controlled in flexibility by simultaneously adjusting isooctane equivalence ratio. The isooctane ignition regime, overall thermal efficiency, and NO{sub x} emissions show strong sensitivity to the fuel delivery advance angle between 20 CA BTDC and 25 CA BTDC. (author)

Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

Directory of Open Access Journals (Sweden)

Jaime Rubio Hervas

2012-01-01

Full Text Available The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

Staged, High-Pressure Oxy-Combustion Technology: Development and Scale-Up

Energy Technology Data Exchange (ETDEWEB)

Axelbaum, Richard; Xia, Fei; Gopan, Akshay; Kumfer, Benjamin

2014-09-30

Washington University in St. Louis and its project partners are developing a unique pressurized oxy-combustion process that aims to improve efficiency and costs by reducing the recycling of flue gas to near zero. Normally, in the absence of recycled flue gas or another inert gas, combustion of fuel and oxygen results in a dramatic increase in temperature of the combustion products and radiant energy, as compared to combustion in air. High heat flux to the boiler tubes may result in a tube surface temperatures that exceed safe operating limits. In the Staged Pressurized Oxy-Combustion (SPOC) process, this problem is addressed by staging the delivery of fuel and by novel combustion design that allows control of heat flux. In addition, the main mode of heat transfer to the steam cycle is by radiation, as opposed to convection. Therefore, the requirement for recycling large amounts of flue gas, for temperature control or to improve convective heat transfer, is eliminated, resulting in a reduction in auxiliary loads. The following report contains a detailed summary of scientific findings and accomplishments for the period of Oct. 1, 2013 to Sept 30, 2014. Results of ASPEN process and CFD modelling activities aimed at improving the SPOC process and boiler design are presented. The effects of combustion pressure and fuel moisture on the plant efficiency are discussed. Combustor pressure is found to have only a minor impact beyond 16 bar. For fuels with moisture content greater than approx 30%, e.g. coal/water slurries, the amount of latent heat of condensation exceeds that which can be utilized in the steam cycle and plant efficiency is reduced significantly. An improved boiler design is presented that achieves a more uniform heat flux profile. In addition, a fundamental study of radiation in high-temperature, high-pressure, particle-laden flows is summarized which provides a more complete understanding of heat transfer in these unusual conditions and to allow for

Experimental study on combustion modes and thrust performance of a staged-combustor of the scramjet with dual-strut

Science.gov (United States)

Yang, Qingchun; Chetehouna, Khaled; Gascoin, Nicolas; Bao, Wen

2016-05-01

To enable the scramjet operate in a wider flight Mach number, a staged-combustor with dual-strut is introduced to hold more heat release at low flight Mach conditions. The behavior of mode transition was examined using a direct-connect model scramjet experiment along with pressure measurements. The typical operating modes of the staged-combustor are analyzed. Fuel injection scheme has a significant effect on the combustor operating modes, particularly for the supersonic combustion mode. Thrust performances of the combustor with different combustion modes and fuel distributions are reported in this paper. The first-staged strut injection has a better engine performance in the operation of subsonic combustion mode. On the contrast, the second-staged strut injection has a better engine performance in the operation of supersonic combustion mode.

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

Energy Technology Data Exchange (ETDEWEB)

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

1989-01-01

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

LOX/Methane Regeneratively-Cooled Rocket Engine Development

Data.gov (United States)

National Aeronautics and Space Administration — The purpose of this project is to advance the technologies required to build a subcritical regeneratively cooled liquid oxygen/methane rocket combustion chamber for...

Combustion performance evaluation of air staging of palm oil blends.

Science.gov (United States)

Mohd Jaafar, Mohammad Nazri; Eldrainy, Yehia A; Mat Ali, Muhammad Faiser; Wan Omar, W Z; Mohd Hizam, Mohd Faizi Arif

2012-02-21

The problems of global warming and the unstable price of petroleum oils have led to a race to develop environmentally friendly biofuels, such as palm oil or ethanol derived from corn and sugar cane. Biofuels are a potential replacement for fossil fuel, since they are renewable and environmentally friendly. This paper evaluates the combustion performance and emission characteristics of Refined, Bleached, and Deodorized Palm Oil (RBDPO)/diesel blends B5, B10, B15, B20, and B25 by volume, using an industrial oil burner with and without secondary air. Wall temperature profiles along the combustion chamber axis were measured using a series of thermocouples fitted axially on the combustion chamber wall, and emissions released were measured using a gas analyzer. The results show that RBDPO blend B25 produced the maximum emission reduction of 56.9% of CO, 74.7% of NOx, 68.5% of SO(2), and 77.5% of UHC compared to petroleum diesel, while air staging (secondary air) in most cases reduces the emissions further. However, increasing concentrations of RBDPO in the blends also reduced the energy released from the combustion. The maximum wall temperature reduction was 62.7% for B25 at the exit of the combustion chamber.

Computational Fluid Dynamics (CFD) applications in rocket propulsion analysis and design

Science.gov (United States)

Mcconnaughey, P. K.; Garcia, R.; Griffin, L. W.; Ruf, J. H.

1993-01-01

Computational Fluid Dynamics (CFD) has been used in recent applications to affect subcomponent designs in liquid propulsion rocket engines. This paper elucidates three such applications for turbine stage, pump stage, and combustor chamber geometries. Details of these applications include the development of a high turning airfoil for a gas generator (GG) powered, liquid oxygen (LOX) turbopump, single-stage turbine using CFD as an integral part of the design process. CFD application to pump stage design has emphasized analysis of inducers, impellers, and diffuser/volute sections. Improvements in pump stage impeller discharge flow uniformity have been seen through CFD optimization on coarse grid models. In the area of combustor design, recent CFD analysis of a film cooled ablating combustion chamber has been used to quantify the interaction between film cooling rate, chamber wall contraction angle, and geometry and their effects of these quantities on local wall temperature. The results are currently guiding combustion chamber design and coolant flow rate for an upcoming subcomponent test. Critical aspects of successful integration of CFD into the design cycle includes a close-coupling of CFD and design organizations, quick turnaround of parametric analyses once a baseline CFD benchmark has been established, and the use of CFD methodology and approaches that address pertinent design issues. In this latter area, some problem details can be simplified while retaining key physical aspects to maintain analytical integrity.

Additive Manufacturing for Affordable Rocket Engines

Science.gov (United States)

West, Brian; Robertson, Elizabeth; Osborne, Robin; Calvert, Marty

2016-01-01

Additive manufacturing (also known as 3D printing) technology has the potential to drastically reduce costs and lead times associated with the development of complex liquid rocket engine systems. NASA is using 3D printing to manufacture rocket engine components including augmented spark igniters, injectors, turbopumps, and valves. NASA is advancing the process to certify these components for flight. Success Story: MSFC has been developing rocket 3D-printing technology using the Selective Laser Melting (SLM) process. Over the last several years, NASA has built and tested several injectors and combustion chambers. Recently, MSFC has 3D printed an augmented spark igniter for potential use the RS-25 engines that will be used on the Space Launch System. The new design is expected to reduce the cost of the igniter by a factor of four. MSFC has also 3D printed and tested a liquid hydrogen turbopump for potential use on an Upper Stage Engine. Additive manufacturing of the turbopump resulted in a 45% part count reduction. To understanding how the 3D printed parts perform and to certify them for flight, MSFC built a breadboard liquid rocket engine using additive manufactured components including injectors, turbomachinery, and valves. The liquid rocket engine was tested seven times in 2016 using liquid oxygen and liquid hydrogen. In addition to exposing the hardware to harsh environments, engineers learned to design for the new manufacturing technique, taking advantage of its capabilities and gaining awareness of its limitations. Benefit: The 3D-printing technology promises reduced cost and schedule for rocket engines. Cost is a function of complexity, and the most complicated features provide the largest opportunities for cost reductions. This is especially true where brazes or welds can be eliminated. The drastic reduction in part count achievable with 3D printing creates a waterfall effect that reduces the number of processes and drawings, decreases the amount of touch

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

Science.gov (United States)

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

2016-01-01

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

COMBUSTION STAGE NUMERICAL ANALYSIS OF A MARINE ENGINE

Directory of Open Access Journals (Sweden)

DOREL DUMITRU VELCEA

2016-06-01

Full Text Available The primary goal of engine design is to maximize each efficiency factor, in order to extract the most power from the least amount of fuel. In terms of fluid dynamics, the volumetric and combustion efficiency are dependent on the fluid dynamics in the engine manifolds and cylinders. Cold flow analysis involves modeling the airflow in the transient engine cycle without reactions. The goal is to capture the mixture formation process by accurately accounting for the interaction of moving geometry with the fluid dynamics of the induction process. The changing characteristics of the air flow jet that tumbles into the cylinder with swirl via intake valves and the exhaust jet through the exhaust valves as they open and close can be determined, along with the turbulence production from swirl and tumble due to compression and squish. The target of this paper was to show how, by using the reverse engineering techniques, one may replicate and simulate the functioning conditions and parameters of an existing marine engine. The departing information were rather scarce in terms of real processes taking place in the combustion stage, but at the end we managed to have a full picture of the main parameters evolution during the combustion phase inside this existing marine engine

Staged combustion - main method for suppressing nitrogen oxides in pulverized-coal fired boilers

Energy Technology Data Exchange (ETDEWEB)

Kotler, V.R. (Vsesoyuznyi Teplotekhnicheskii Institut (USSR))

1989-08-01

Describes principles behind staged combustion, which is based on organizing furnace operations so that only part of the air from the fuel is taken into the furnace. The remaining air, which is needed for combustion, is fed as a tertiary blast jet into the intermediate flame zone. Following inflammation and combustion of the volatile matter, the oxygen concentration in the flame drops sharply causing a retardation of the oxidation reactions forming NO and an intensification of the reactions causing the nitrogen-containing radicals NH{sub i} and CN to be converted into N{sub 2}. When the reducing agents CO, H{sub 2} and CH{sub 4} are present in certain flame zones, even the nitrogen oxide is reduced to N{sub 2}. The NO concentrations in the flame are reduced until the jet of tertiary air is introduced. Discusses with reference to practice in the USA and Western Europe how to achieve maximum effect of this method for different types of boiler and presents the results of observations of the introduction of staged combustion to the BKZ-210-140 boiler burning Kuznetsk gassy coal. 5 refs.

Modified Regression Rate Formula of PMMA Combustion by a Single Plane Impinging Jet

Directory of Open Access Journals (Sweden)

Tsuneyoshi Matsuoka

2017-01-01

Full Text Available A modified regression rate formula for the uppermost stage of CAMUI-type hybrid rocket motor is proposed in this study. Assuming a quasi-steady, one-dimensional, an energy balance against a control volume near the fuel surface is considered. Accordingly, the regression rate formula which can calculate the local regression rate by the quenching distance between the flame and the regression surface is derived. An experimental setup which simulates the combustion phenomenon involved in the uppermost stage of a CAMUI-type hybrid rocket motor was constructed and the burning tests with various flow velocities and impinging distances were performed. A PMMA slab of 20 mm height, 60 mm width, and 20 mm thickness was chosen as a sample specimen and pure oxygen and O2/N2 mixture (50/50 vol.% were employed as the oxidizers. The time-averaged regression rate along the fuel surface was measured by a laser displacement sensor. The quenching distance during the combustion event was also identified from the observation. The comparison between the purely experimental and calculated values showed good agreement, although a large systematic error was expected due to the difficulty in accurately identifying the quenching distance.

Staged, High-Pressure Oxy-Combustion Technology: Development and Scale-Up

Energy Technology Data Exchange (ETDEWEB)

Axelbaum, Richard [Washington Univ., St. Louis, MO (United States); Kumfer, Benjamin [Washington Univ., St. Louis, MO (United States); Gopan, Akshay [Washington Univ., St. Louis, MO (United States); Yang, Zhiwei [Washington Univ., St. Louis, MO (United States); Phillips, Jeff [Electric Power Research Inst. (EPRI), Palo Alto, CA (United States); Pint, Bruce [Oak Ridge National Lab. (ORNL), Oak Ridge, TN (United States)

2017-12-29

The immediate need for a high efficiency, low cost carbon capture process has prompted the recent development of pressurized oxy-combustion. With a greater combustion pressure the dew point of the flue gas is increased, allowing for effective integration of the latent heat of flue gas moisture into the Rankine cycle. This increases the net plant efficiency and reduces costs. A novel, transformational process, named Staged, Pressurized Oxy-Combustion (SPOC), achieves additional step changes in efficiency and cost reduction by significantly reducing the recycle of flue gas. The research and development activities conducted under Phases I and II of this project (FE0009702) include: SPOC power plant cost and performance modeling, CFD-assisted design of pressurized SPOC boilers, theoretical analysis of radiant heat transfer and ash deposition, boiler materials corrosion testing, construction of a 100 kWth POC test facility, and experimental testing. The results of this project have advanced the technology readiness level (TRL) of the SPOC technology from 1 to 5.

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

International Nuclear Information System (INIS)

Corvin, P.; Steinhilber, H.

1981-01-01

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

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

Science.gov (United States)

Heitkotter, Robert H

1956-01-01

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

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

Science.gov (United States)

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

2017-09-01

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

Study of Liquid Breakup Process in Solid Rocket Motor Nozzle

Science.gov (United States)

2016-02-16

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

Aluminum Agglomeration and Trajectory in Solid Rocket Motors

National Research Council Canada - National Science Library

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

2007-01-01

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

Emission characteristics of multiple stage diesel combustion. Effect of exhaust gas recirculation; Nidan nensho diesel kikan no haishutsubutsu tokusei. EGR no eikyo

Energy Technology Data Exchange (ETDEWEB)

Hashizume, T.; Miyamoto, T.; Akagawa, H.; Tsujimura, K. [New A.C.E. Institute Co. Ltd., Tokyo (Japan)

1998-05-01

For an aim to reduce NOx emission from diesel engines, it has become possible to realize it with smoke emission maintained at low levels by taking the following steps: initial combustion is carried out as lean pre-mixed combustion by adopting early fuel injection; the fuel is injected again after completion of this combustion; and EGR is combined with two-stage combustion which performs diffusion combustion under high temperature atmosphere. When a large quantity of EGR is used, cylinder temperature drops to have ignition timing delayed in the first stage, serving for improving fuel consumption. The problem of increase in smoke generation is solved by optimizing the injection timing at the second stage to suppress smoke generation increase, resulting in realization of lower NOx emission. By completing the second-stage fuel injection before ignition of the first-stage injection, it was possible to realize further lower NOx emission. Smoke increase due to higher EGR ratio was suppressed by pre-mixing both fuels injected in the first and second stages, although this is a high load operation. In addition, oxygen concentration and cylinder temperature were reduced, the gas pre-mixture was homogenized, and combustion velocity was suppressed by delaying the angle of ignition timing. This made low smoke combustion at {lambda} = 1 possible even in compressed ignition combustion. 8 refs., 12 figs., 1 tab.

Combustion of coal gas fuels in a staged combustor

Science.gov (United States)

Rosfjord, T. J.; Mcvey, J. B.; Sederquist, R. A.; Schultz, D. F.

1982-01-01

Gaseous fuels produced from coal resources generally have heating values much lower than natural gas; the low heating value could result in unstable or inefficient combustion. Coal gas fuels may contain ammonia which if oxidized in an uncontrolled manner could result in unacceptable nitrogen oxide exhaust emission levels. Previous investigations indicate that staged, rich-lean combustion represents a desirable approach to achieve stable, efficient, low nitrogen oxide emission operation for coal-derived liquid fuels contaning up to 0.8-wt pct nitrogen. An experimental program was conducted to determine whether this fuel tolerance can be extended to include coal-derived gaseous fuels. The results of tests with three nitrogen-free fuels having heating values of 100, 250, and 350 Btu/scf and a 250 Btu/scf heating value doped to contain 0.7 pct ammonia are presented.

Theoretical Acoustic Absorber Design Approach for LOX/LCH4 Pintle Injector Rocket Engines

Science.gov (United States)

Candelaria, Jonathan

Liquid rocket engines, or LREs, have served a key role in space exploration efforts. One current effort involves the utilization of liquid oxygen (LOX) and liquid methane (LCH4) LREs to explore Mars with in-situ resource utilization for propellant production. This on-site production of propellant will allow for greater payload allocation instead of fuel to travel to the Mars surface, and refueling of propellants to travel back to Earth. More useable mass yields a greater benefit to cost ratio. The University of Texas at El Paso's (UTEP) Center for Space Exploration and Technology Research Center (cSETR) aims to further advance these methane propulsion systems with the development of two liquid methane - liquid oxygen propellant combination rocket engines. The design of rocket engines, specifically liquid rocket engines, is complex in that many variables are present that must be taken into consideration in the design. A problem that occurs in almost every rocket engine development program is combustion instability, or oscillatory combustion. It can result in the destruction of the rocket, subsequent destruction of the vehicle and compromise the mission. These combustion oscillations can vary in frequency from 100 to 20,000 Hz or more, with varying effects, and occur from different coupling phenomena. It is important to understand the effects of combustion instability, its physical manifestations, how to identify the instabilities, and how to mitigate or dampen them. Linear theory methods have been developed to provide a mathematical understanding of the low- to mid-range instabilities. Nonlinear theory is more complex and difficult to analyze mathematically, therefore no general analytical method that yields a solution exists. With limited resources, time, and the advice of our NASA mentors, a data driven experimental approach utilizing quarter wave acoustic dampener cavities was designed. This thesis outlines the methodology behind the design of an acoustic

Linear stability analysis in a solid-propellant rocket motor

Energy Technology Data Exchange (ETDEWEB)

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

1995-10-01

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

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

Energy Technology Data Exchange (ETDEWEB)

Porter, F.S. E-mail: frederick.s.porter@gsfc.nasa.gov; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T

2000-04-07

The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight.

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

International Nuclear Information System (INIS)

Porter, F.S.; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T.

2000-01-01

The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight

Influence of atomization quality modulation on flame dynamics in a hypergolic rocket engine

Directory of Open Access Journals (Sweden)

Moritz Schulze

2016-09-01

Full Text Available For the numerical evaluation of the thermoacoustic stability of rocket engines often hybrid methods are applied, which separate the computation of wave propagation in the combustor from the analysis of the flame response to acoustic perturbations. Closure requires a thermoacoustic feedback model which provides the heat release fluctuation in the source term of the employed wave transport equations. The influence of the acoustic fluctuations in the combustion chamber on the heat release fluctuations from the modulation of the atomization of the propellants in a hypergolic upper stage rocket engine is studied. Numerical modeling of a single injector provides the time mean reacting flow field. A network of transfer functions representing all aspects relevant for the feedback model is presented. Analytical models for the injector admittances and for the atomization transfer functions are provided. The dynamics of evaporation and combustion are studied numerically and the numerical results are analyzed. An analytical approximation of the computed flame transfer function is combined with the analytical models for the injector and the atomization quality to derive the feedback model for the wave propagation code. The evaluation of this model on the basis of the Rayleigh index reveals the thermoacoustic driving potential originating from the fluctuating spray quality.

Regarding the perturbed operating process of DB propellant rocket motor at extreme initial grain temperatures

Directory of Open Access Journals (Sweden)

Ioan ION

2012-03-01

Full Text Available Despite many decades of study, the combustion instability of several DB propellants is still of particular concern, especially at extreme grain temperature conditions of rocket motor operating. The purpose of the first part of the paper is to give an overview of our main experimental results on combustion instabilities and pressure oscillations in DB propellant segmented grain rocket motors (SPRM-01, large L/D ratio, working at extreme initial grain temperatures. Thus, we recorded some particular pressure-time traces with significant perturbed pressure signal that was FFT analysed. An updated mathematical model incorporating transient frequency-dependent combustion response, in conjunction with pressure-dependent burning, is applied to investigate and predict the DB propellant combustion instability phenomenon. The susceptibility of the tested motor SPRM-01 with DB propellant to get a perturbed working and to go unstable with pressure was evidenced and this risk has to be evaluated. In the last part of our paper we evaluated the influence of recorded perturbed thrust on the rocket behaviour on the trajectory. The study revealed that at firing-table initial conditions, this kind of perturbed motor operating may not lead to an unstable rocket flight, but the ballistic parameters would be influenced in an unacceptable manner.

Two-step rocket engine bipropellant valve concept

Science.gov (United States)

Capps, J. E.; Ferguson, R. E.; Pohl, H. O.

1969-01-01

Initiating combustion of altitude control rocket engines in a precombustion chamber of ductile material reduces high pressure surges generated by hypergolic propellants. Two-step bipropellant valve concepts control initial propellant flow into precombustion chamber and subsequent full flow into main chamber.

Analysis of Chemical Reaction Kinetics Behavior of Nitrogen Oxide During Air-staged Combustion in Pulverized Boiler

Directory of Open Access Journals (Sweden)

Jun-Xia Zhang

2016-03-01

Full Text Available Because the air-staged combustion technology is one of the key technologies with low investment running costs and high emission reduction efficiency for the pulverized boiler, it is important to reveal the chemical reaction kinetics mechanism for developing various technologies of nitrogen oxide reduction emissions. At the present work, a three-dimensional mesh model of the large-scale four corner tangentially fired boiler furnace is established with the GAMBIT pre-processing of the FLUENT software. The partial turbulent premixed and diffusion flame was simulated for the air-staged combustion processing. Parameters distributions for the air-staged and no the air-staged were obtained, including in-furnace flow field, temperature field and nitrogen oxide concentration field. The results show that the air-staged has more regular velocity field, higher velocity of flue gas, higher turbulence intensity and more uniform temperature of flue gas. In addition, a lower negative pressure zone and lower O2 concentration zone is formed in the main combustion zone, which is conducive to the NO of fuel type reduced to N2, enhanced the effect of NOx reduction. Copyright © 2016 BCREC GROUP. All rights reserved Received: 5th November 2015; Revised: 14th January 2016; Accepted: 16th January 2016  How to Cite: Zhang, J.X., Zhang, J.F. (2016. Analysis of Chemical Reaction Kinetics Behavior of Nitrogen Oxide During Air-staged Combustion in Pulverized Boiler. Bulletin of Chemical Reaction Engineering & Catalysis, 11 (1: 100-108. (doi:10.9767/bcrec.11.1.431.100-108 Permalink/DOI: http://dx.doi.org/10.9767/bcrec.11.1.431.100-108

How Does Rocket Propulsion Work? The most common answer to ...

Indian Academy of Sciences (India)

internal combustion engines. The fuel or propellant is stored in the fuel tank. Here we will consider liquid hydrogen as the fuel. For the combustion to take place in outer space or in the absence of atmospheric oxygen the rocket carries along an oxidizer; here we will consider liquid oxygen as the oxidizer. The oxidizer or in.

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

Science.gov (United States)

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

2017-06-01

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

Parameters Affecting the Erosive Burning of Solid Rocket Motor

OpenAIRE

Abdelaziz Almostafa; Guozhu Liang; Elsayed Anwer

2018-01-01

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

Liquid Rocket Engine Testing Overview

Science.gov (United States)

Rahman, Shamim

2005-01-01

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

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

Science.gov (United States)

1972-01-01

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

Lean-rich axial stage combustion in a can-annular gas turbine engine

Science.gov (United States)

Laster, Walter R.; Szedlacsek, Peter

2016-06-14

An apparatus and method for lean/rich combustion in a gas turbine engine (10), which includes a combustor (12), a transition (14) and a combustor extender (16) that is positioned between the combustor (12) and the transition (14) to connect the combustor (12) to the transition (14). Openings (18) are formed along an outer surface (20) of the combustor extender (16). The gas turbine (10) also includes a fuel manifold (28) to extend along the outer surface (20) of the combustor extender (16), with fuel nozzles (30) to align with the respective openings (18). A method (200) for axial stage combustion in the gas turbine engine (10) is also presented.

Aerodynamics and flow characterisation of multistage rockets

Science.gov (United States)

Srinivas, G.; Prakash, M. V. S.

2017-05-01

The main objective of this paper is to conduct a systematic flow analysis on single, double and multistage rockets using ANSYS software. Today non-air breathing propulsion is increasing dramatically for the enhancement of space exploration. The rocket propulsion is playing vital role in carrying the payload to the destination. Day to day rocket aerodynamic performance and flow characterization analysis has becoming challenging task to the researchers. Taking this task as motivation a systematic literature is conducted to achieve better aerodynamic and flow characterization on various rocket models. The analyses on rocket models are very little especially in numerical side and experimental area. Each rocket stage analysis conducted for different Mach numbers and having different flow varying angle of attacks for finding the critical efficiency performance parameters like pressure, density and velocity. After successful completion of the analysis the research reveals that flow around the rocket body for Mach number 4 and 5 best suitable for designed payload. Another major objective of this paper is to bring best aerodynamics flow characterizations in both aero and mechanical features. This paper also brings feature prospectus of rocket stage technology in the field of aerodynamic design.

Multivariable optimization of liquid rocket engines using particle swarm algorithms

Science.gov (United States)

Jones, Daniel Ray

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

A study of air breathing rockets. 3: Supersonic mode combustors

Science.gov (United States)

Masuya, G.; Chinzel, N.; Kudo, K.; Murakami, A.; Komuro, T.; Ishii, S.

An experimental study was made on supersonic mode combustors of an air breathing rocket engine. Supersonic streams of room-temperature air and hot fuel-rich rocket exhaust were coaxially mixed and burned in a concially diverging duct of 2 deg half-angle. The effect of air inlet Mach number and excess air ratio was investigated. Axial wall pressure distribution was measured to calculate one dimensional change of Mach number and stagnation temperature. Calculated results showed that supersonic combustion occurred in the duct. At the exit of the duct, gas sampling and Pitot pressure measurement was made, from which radial distributions of various properties were deduced. The distribution of mass fraction of elements from rocket exhaust showed poor mixing performance in the supersonic mode combustors compared with the previously investigated cylindrical subsonic mode combustors. Secondary combustion efficiency correlated well with the centerline mixing parameter, but not with Annushkin's non-dimensional combustor length. No major effect of air inlet Mach number or excess air ratio was seen within the range of conditions under which the experiment was conducted.

High speed diagnostics for characterization of oxygen / hydrogen rocket injector flowfields

Science.gov (United States)

Locke, Justin M.

Temporally-resolved diagnostics are needed to characterize the highly-turbulent flowfields of rocket engine combustors. Two different high speed diagnostic techniques have been applied successfully to study the combustion and mixing characteristics of single-element shear-coaxial injector flowfield in a rocket combustor. Tunable diode laser absorption spectroscopy (TDLAS) is applied to make in-situ measurements with combusting gas-gas propellants, and high speed imaging incorporating backlighting is applied to combusting and non-combusting liquid-gas and gas-gas propellants. Tunable diode laser absorption spectroscopy has been used to make path-integrated temperature and H2O mole fraction measurements in a gaseous oxygen / gaseous hydrogen uni-element rocket chamber with hot background flow. Four mixture conditions were studied at a nominal chamber pressure of 115 psia. Near infrared diode lasers were utilized to target rovibrational transitions of water vapor, which is created through the combustion processes. Both direct absorption spectroscopy and wavelength modulation spectroscopy with second harmonic normalized by first harmonic (1f-normalized WMS-2 f) techniques were applied, with the harmonic detection technique found to yield the best results. Centerline measurements were made at two axial locations, in the near-injector region and far-field region further downstream. Time-resolved measurements of temperature and H2O mole fraction in the rocket chamber are presented. The TDLAS measurements shows clear differences between the near-injector and downstream measurement locations. Mean path-integrated temperatures and H2O mole fractions in the near-injector region are lower than the downstream measurement location. Fluctuations in path-integrated temperature and H2O mole fraction in the downstream location are significantly greater than the near-injector region. This suggests increased turbulence and larger-scale mixing processes are occurring in the downstream

Review on pressure swirl injector in liquid rocket engine

Science.gov (United States)

Kang, Zhongtao; Wang, Zhen-guo; Li, Qinglian; Cheng, Peng

2018-04-01

The pressure swirl injector with tangential inlet ports is widely used in liquid rocket engine. Commonly, this type of pressure swirl injector consists of tangential inlet ports, a swirl chamber, a converging spin chamber, and a discharge orifice. The atomization of the liquid propellants includes the formation of liquid film, primary breakup and secondary atomization. And the back pressure and temperature in the combustion chamber could have great influence on the atomization of the injector. What's more, when the combustion instability occurs, the pressure oscillation could further affects the atomization process. This paper reviewed the primary atomization and the performance of the pressure swirl injector, which include the formation of the conical liquid film, the breakup and atomization characteristics of the conical liquid film, the effects of the rocket engine environment, and the response of the injector and atomization on the pressure oscillation.

Hybrid Approach for Modeling Chemical Kinetics and Turbulence Effects on Combustion-Instability, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Combustion instabilities pose a significant technical risk in the development of liquid and solid rocket motors. Much of the effort in modeling combustion...

Star-grain rocket motor - nonsteady internal ballistics

Energy Technology Data Exchange (ETDEWEB)

Loncaric, S.; Greatrix, D.R.; Fawaz, Z. [Ryerson University, Dept. of Aerospace Engineering, Toronto (Canada)

2004-01-01

The nonsteady internal ballistics of a star-grain solid-propellant rocket motor are investigated through a numerical simulation model that incorporates both the internal flow and surrounding structure. The effects of structural vibration on burning rate augmentation and wave development in nonsteady operation are demonstrated. The amount of damping plays a role in influencing the predicted axial combustion instability symptoms of the motor. The variation in oscillation frequencies about a given star grain section periphery, and along the grain with different levels of burn-back, also influences the means by which the local acceleration drives the combustion and flow behaviour. (authors)

Longitudinal acoustic instabilities in slender solid propellant rockets : linear analysis

OpenAIRE

García Schafer, Juan Esteban; Liñán Martínez, Amable

2001-01-01

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

Development of high performance hybrid rocket fuels

Science.gov (United States)

Zaseck, Christopher R.

In this document I discuss paraffin fuel combustion and investigate the effects of additives on paraffin entrainment and regression. In general, hybrid rockets offer an economical and safe alternative to standard liquid and solid rockets. However, slow polymeric fuel regression and low combustion efficiency have limited the commercial use of hybrid rockets. Paraffin is a fast burning fuel that has received significant attention in the 2000's and 2010's as a replacement for standard fuels. Paraffin regresses three to four times faster than polymeric fuels due to the entrainment of a surface melt layer. However, further regression rate enhancement over the base paraffin fuel is necessary for widespread hybrid rocket adoption. I use a small scale opposed flow burner to investigate the effect of additives on the combustion of paraffin. Standard additives such as aluminum combust above the flame zone where sufficient oxidizer levels are present. As a result no heat is generated below the flame itself. In small scale opposed burner experiments the effect of limited heat feedback is apparent. Aluminum in particular does not improve the regression of paraffin in the opposed burner. The lack of heat feedback from additive combustion limits the applicability of the opposed burner. In turn, the results obtained in the opposed burner with metal additive loaded hybrid fuels do not match results from hybrid rocket experiments. In addition, nano-scale aluminum increases melt layer viscosity and greatly slows the regression of paraffin in the opposed flow burner. However, the reactive additives improve the regression rate of paraffin in the opposed burner where standard metals do not. At 5 wt.% mechanically activated titanium and carbon (Ti-C) improves the regression rate of paraffin by 47% in the opposed burner. The mechanically activated Ti C likely reacts in or near the melt layer and provides heat feedback below the flame region that results in faster opposed burner regression

Not just rocket science

Energy Technology Data Exchange (ETDEWEB)

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

2007-10-15

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

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

Science.gov (United States)

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

2010-01-01

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

Current and Future Critical Issues in Rocket Propulsion Systems

Science.gov (United States)

Navaz, Homayun K.; Dix, Jeff C.

1998-01-01

The objective of this research was to tackle several problems that are currently of great importance to NASA. In a liquid rocket engine several complex processes take place that are not thoroughly understood. Droplet evaporation, turbulence, finite rate chemistry, instability, and injection/atomization phenomena are some of the critical issues being encountered in a liquid rocket engine environment. Pulse Detonation Engines (PDE) performance, combustion chamber instability analysis, 60K motor flowfield pattern from hydrocarbon fuel combustion, and 3D flowfield analysis for the Combined Cycle engine were of special interest to NASA. During the summer of 1997, we made an attempt to generate computational results for all of the above problems and shed some light on understanding some of the complex physical phenomena. For this purpose, the Liquid Thrust Chamber Performance (LTCP) code, mainly designed for liquid rocket engine applications, was utilized. The following test cases were considered: (1) Characterization of a detonation wave in a Pulse Detonation Tube; (2) 60K Motor wall temperature studies; (3) Propagation of a pressure pulse in a combustion chamber (under single and two-phase flow conditions); (4) Transonic region flowfield analysis affected by viscous effects; (5) Exploring the viscous differences between a smooth and a corrugated wall; and (6) 3D thrust chamber flowfield analysis of the Combined Cycle engine. It was shown that the LTCP-2D and LTCP-3D codes are capable of solving complex and stiff conservation equations for gaseous and droplet phases in a very robust and efficient manner. These codes can be run on a workstation and personal computers (PC's).

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

Science.gov (United States)

Westra, Douglas G.; West, Jeffrey S.

2014-01-01

NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

Parameters Affecting the Erosive Burning of Solid Rocket Motor

Directory of Open Access Journals (Sweden)

Abdelaziz Almostafa

2018-01-01

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

Controllable Solid Propulsion Combustion and Acoustic Knowledge Base Improvements

Science.gov (United States)

McCauley, Rachel; Fischbach, Sean; Fredrick, Robert

2012-01-01

Controllable solid propulsion systems have distinctive combustion and acoustic environments that require enhanced testing and analysis techniques to progress this new technology from development to production. In a hot gas valve actuating system, the movement of the pintle through the hot gas exhibits complex acoustic disturbances and flow characteristics that can amplify induced pressure loads that can damage or detonate the rocket motor. The geometry of a controllable solid propulsion gas chamber can set up unique unsteady flow which can feed acoustic oscillations patterns that require characterization. Research in this area aids in the understanding of how best to design, test, and analyze future controllable solid rocket motors using the lessons learned from past government programs as well as university research and testing. This survey paper will give the reader a better understanding of the potentially amplifying affects propagated by a controllable solid rocket motor system and the knowledge of the tools current available to address these acoustic disturbances in a preliminary design. Finally the paper will supply lessons learned from past experiences which will allow the reader to come away with understanding of what steps need to be taken when developing a controllable solid rocket propulsion system. The focus of this survey will be on testing and analysis work published by solid rocket programs and from combustion and acoustic books, conference papers, journal articles, and additionally from subject matter experts dealing currently with controllable solid rocket acoustic analysis.

Optimization of the rocket mode trajectory in a rocket based combined cycle (RBCC) engine powered SSTO vehicle

Science.gov (United States)

Foster, Richard W.

1989-07-01

The application of rocket-based combined cycle (RBCC) engines to booster-stage propulsion, in combination with all-rocket second stages in orbital-ascent missions, has been studied since the mid-1960s; attention is presently given to the case of the 'ejector scramjet' RBCC configuration's application to SSTO vehicles. While total mass delivered to initial orbit is optimized at Mach 20, payload delivery capability to initial orbit optimizes at Mach 17, primarily due to the reduction of hydrogen fuel tankage structure, insulation, and thermal protection system weights.

A reliability as an independent variable (RAIV) methodology for optimizing test planning for liquid rocket engines

Science.gov (United States)

Strunz, Richard; Herrmann, Jeffrey W.

2011-12-01

The hot fire test strategy for liquid rocket engines has always been a concern of space industry and agency alike because no recognized standard exists. Previous hot fire test plans focused on the verification of performance requirements but did not explicitly include reliability as a dimensioning variable. The stakeholders are, however, concerned about a hot fire test strategy that balances reliability, schedule, and affordability. A multiple criteria test planning model is presented that provides a framework to optimize the hot fire test strategy with respect to stakeholder concerns. The Staged Combustion Rocket Engine Demonstrator, a program of the European Space Agency, is used as example to provide the quantitative answer to the claim that a reduced thrust scale demonstrator is cost beneficial for a subsequent flight engine development. Scalability aspects of major subsystems are considered in the prior information definition inside the Bayesian framework. The model is also applied to assess the impact of an increase of the demonstrated reliability level on schedule and affordability.

Subsonic Glideback Rocket Demonstrator Flight Testing

Science.gov (United States)

DeTurris, Dianne J.; Foster, Trevor J.; Barthel, Paul E.; Macy, Daniel J.; Droney, Christopher K.; Talay, Theodore A. (Technical Monitor)

2001-01-01

For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

Mixing and combustion enhancement of Turbocharged Solid Propellant Ramjet

Science.gov (United States)

Liu, Shichang; Li, Jiang; Zhu, Gen; Wang, Wei; Liu, Yang

2018-02-01

Turbocharged Solid Propellant Ramjet is a new concept engine that combines the advantages of both solid rocket ramjet and Air Turbo Rocket, with a wide operation envelope and high performance. There are three streams of the air, turbine-driving gas and augment gas to mix and combust in the afterburner, and the coaxial intake mode of the afterburner is disadvantageous to the mixing and combustion. Therefore, it is necessary to carry out mixing and combustion enhancement research. In this study, the numerical model of Turbocharged Solid Propellant Ramjet three-dimensional combustion flow field is established, and the numerical simulation of the mixing and combustion enhancement scheme is conducted from the aspects of head region intake mode to injection method in afterburner. The results show that by driving the compressed air to deflect inward and the turbine-driving gas to maintain strong rotation, radial and tangential momentum exchange of the two streams can be enhanced, thereby improving the efficiency of mixing and combustion in the afterburner. The method of injecting augment gas in the transverse direction and making sure the injection location is as close as possible to the head region is beneficial to improve the combustion efficiency. The outer combustion flow field of the afterburner is an oxidizer-rich environment, while the inner is a fuel-rich environment. To improve the efficiency of mixing and combustion, it is necessary to control the injection velocity of the augment gas to keep it in the oxygen-rich zone of the outer region. The numerical simulation for different flight conditions shows that the optimal mixing and combustion enhancement scheme can obtain high combustion efficiency and have excellent applicability in a wide working range.

Experimental study of a staged combustion system for stationary gas turbine applications

Science.gov (United States)

Lamont, Warren G.

Two optically accessible experimental test rigs were designed and constructed to investigate a staged or distributed combustion system for stationary gas turbine applications. The test rigs were fuelled with natural gas and featured two combustion zones: the main combustion zone (MCZ) and the secondary combustion zone (SCZ). The MCZ is a swirl stabilized dump combustor and the SCZ, which is axially downstream from the MCZ, is formed by a transverse jet injecting a premixed fuel/air mixture into the vitiated stream. After installing and commissioning the test rig, an emission survey was conducted to investigate the SCZ conditions, equivalence ratio and momentum ratio, that produce low NOx emissions and give a higher temperature rise before a simulated high pressure turbine than firing only the MCZ. The emission survey found several operating conditions that show the benefit of combustion staging. These beneficial conditions had an SCZ equivalence ratio between 0.41 and 1.12. The data from the emission survey was then used to create an artificial neural network (ANN). The ANN used a multi-layer feed-forward network architecture and was trained with experimental data using the backpropagation training algorithm. The ANN was then used to create performance maps and optimum operational regions were sought. Lastly, optical diagnostics were used to obtain information on the nature of the SCZ reactive jet. The diagnostics included high speed CH* chemiluminescence, OH planar laser induced fluorescence (PLIF) and dual-pump coherent anti-Stokes Raman scattering (CARS). The chemiluminescence and PLIF were used to qualitatively determine the size and shape of the transverse jet reaction zone. Dual-pump CARS was used to quantitatively determine the temperature and H2/N2 concentration ratio profile at the mid-plane of the transverse jet. Dual-pump CARS data was collected for four operating conditions but only one is presented in this dissertation. For the condition presented, the

Mean Flow Augmented Acoustics in Rocket Systems

Science.gov (United States)

Fischbach, Sean R.

2014-01-01

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

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

Science.gov (United States)

Williams, R. W. (Compiler)

1996-01-01

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

Overall evaluation of combustion and NO(x) emissions for a down-fired 600 MW(e) supercritical boiler with multiple injection and multiple staging.

Science.gov (United States)

Kuang, Min; Li, Zhengqi; Liu, Chunlong; Zhu, Qunyi

2013-05-07

To achieve significant reductions in NOx emissions and to eliminate strongly asymmetric combustion found in down-fired boilers, a deep-air-staging combustion technology was trialed in a down-fired 600 MWe supercritical utility boiler. By performing industrial-sized measurements taken of gas temperatures and species concentrations in the near wing-wall region, carbon in fly ash and NOx emissions at various settings, effects of overfire air (OFA) and staged-air damper openings on combustion characteristics, and NOx emissions within the furnace were experimentally determined. With increasing the OFA damper opening, both fluctuations in NOx emissions and carbon in fly ash were initially slightly over OFA damper openings of 0-40% but then lengthened dramatically in openings of 40-70% (i.e., NOx emissions reduced sharply accompanied by an apparent increase in carbon in fly ash). Decreasing the staged-air declination angle clearly increased the combustible loss but slightly influenced NOx emissions. In comparison with OFA, the staged-air influence on combustion and NOx emissions was clearly weaker. Only at a high OFA damper opening of 50%, the staged-air effect was relatively clear, i.e., enlarging the staged-air damper opening decreased carbon in fly ash and slightly raised NOx emissions. By sharply opening the OFA damper to deepen the air-staging conditions, although NOx emissions could finally reduce to 503 mg/m(3) at 6% O2 (i.e., an ultralow NOx level for down-fired furnaces), carbon in fly ash jumped sharply to 15.10%. For economical and environment-friendly boiler operations, an optimal damper opening combination (i.e., 60%, 50%, and 50% for secondary air, staged-air, and OFA damper openings, respectively) was recommended for the furnace, at which carbon in fly ash and NOx emissions attained levels of about 10% and 850 mg/m(3) at 6% O2, respectively.

Computational Fluid Dynamic Modeling of Rocket Based Combined Cycle Engine Flowfields

Science.gov (United States)

Daines, Russell L.; Merkle, Charles L.

1994-01-01

Computational Fluid Dynamic techniques are used to study the flowfield of a fixed geometry Rocket Based Combined Cycle engine operating in rocket ejector mode. Heat addition resulting from the combustion of injected fuel causes the subsonic engine flow to choke and go supersonic in the slightly divergent combustor-mixer section. Reacting flow computations are undertaken to predict the characteristics of solutions where the heat addition is determined by the flowfield. Here, adaptive gridding is used to improve resolution in the shear layers. Results show that the sonic speed is reached in the unheated portions of the flow first, while the heated portions become supersonic later. Comparison with results from another code show reasonable agreement. The coupled solutions show that the character of the combustion-based thermal choking phenomenon can be controlled reasonably well such that there is opportunity to optimize the length and expansion ratio of the combustor-mixer.

Two-stage combustion, a new concept in reducing alkali related operational problem; Tvaastegsfoerbraenning, ett nytt koncept att minska alkalirelaterad driftproblematik

Energy Technology Data Exchange (ETDEWEB)

Hjoernhede, Anders; Hermansson, Sven; Seemann, Martin; Alamia, Alberto; Larsson, Anton; Davidsson, Kent; Niklasson, Fredrik; Atongka Tchoffor, Placid; Jones, Frida

2012-07-01

A possible way to reduce alkali related problems during thermal conversion of fuels into heat and power is to thermally split the fuel into a relatively alkali-lean gas stream and a relatively alkali-rich char stream. The alkali-lean gas stream could then be combusted and used for high-temperature purpose, e.g. superheating of steam. The char residue could be combusted for heating of e.g. steam tubes at inferior temperatures. The purpose of this project has been to study and demonstrate such two-stage combustion, with separation of the early devolatilization from alkali-rich fuels for steam superheating. Within the project, it has been investigated if this type of thermal conversion is thermo-economically feasible. Furthermore, it has been investigated in laboratory and pilot scale trials, which kinds of fuels, temperature levels and fuel residence times that are required to achieve a flue gas with significantly lower concentrations of alkaloids, compared to single-stage conversion. The results from the thermo-economical simulations show that it is possible to arrange the heat exchanger surfaces in a power plant for thermal two-stage conversion. Compared to traditional one-stage combustion, the electricity efficiency could be increased from 35 % to more than 38 %. However, the results from combustion, gasification and pyrolysis in lab, together with gasification trials in the Chalmers pilot plant, show that the investigated woody fuels (bark and wood pellets) do not seem to be feasible for two-stage conversion. These fuels emit alkaloids exclusively in conjunction with the devolatilization in opposite to the desired effect. On the other hand, straw is a potential fuel candidate for two-stage conversion, since straw emits alkaloids during both devolatilization and char conversion. Furthermore, the results show that increased temperature generally causes increased alkali release. From the results it is suggested that the propensity of a fuel to release alkali during

Effect of the Reburning Zone Stoichiometry on the Nox Concentration at the Three-Stage Combustion of Pulverized Coal

Directory of Open Access Journals (Sweden)

Chernetskaya Nelya

2016-01-01

Full Text Available Numerical study of heat and mass transfer taking into account the combustion of coal particles in the furnace at the three-stage combustion of pulverized coal was performed. Analysis of the reburning zone stoichiometry on the concentration of nitrogen oxides at the furnace outlet was made. The values of excess air in the primary and reburning combustion zones, providing for the concentration of nitrogen oxides at the furnace outlet is not more than 350 mg/m3 and unburned carbon not more than 1 % when burning coal with a high content of nitrogen were established.

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

Science.gov (United States)

Zhang, Shuai; Hu, Fan; Zhang, Weihua

2016-10-01

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

Performance and technical feasibility comparison of reusable launch systems: A synthesis of the ESA winged launcher studies

Science.gov (United States)

Berry, W.; Grallert, H.

1996-02-01

The paper presents a synthesis of the performance and technical feasibility assessment of 7 reusable launcher types, comprising 13 different vehicles, studied by European Industry for ESA in the ESA Winged Launcher Study in the period January 1988 to May 1994. The vehicles comprised single-stage-to-orbit (SSTO) and two-stage-to-orbit (TSTO) vehicles, propelled by either air-breathing/rocket propulsion or entirely by rocket propulsion. The results showed that an SSTO vehicle of the HOTOL-type, propelled by subsonic combustion air-breathing/rocket engines could barely deliver the specified payload mass and was aerodynamically unstable; that a TSTO vehicle of the Saenger type, employing subsonic combustion airbreathing propulsion in its first stage and rocket propulsion in its second stage, could readily deliver the specified payload mass and was found to be technically feasible and versatile; that an SSTO vehicle of the NASP type, propelled by supersonic combustion airbreathing/rocket propulsion was able to deliver a reduced payload mass, was very complex and required very advanced technologies; that an air-launched rocket propelled vehicle of the Interim HOTOL type, although technically feasible, could deliver only a reduced payload mass, being constrained by the lifting capability of the carrier airplane; that three different, entirely rocket-propelled vehicles could deliver the specified payload mass, were technically feasible but required relatively advanced technologies.

Investigation of the cooling film distribution in liquid rocket engine

Directory of Open Access Journals (Sweden)

Luís Antonio Silva

2011-05-01

Full Text Available This study presents the results of the investigation of a cooling method widely used in the combustion chambers, which is called cooling film, and it is applied to a liquid rocket engine that uses as propellants liquid oxygen and kerosene. Starting from an engine cooling, whose film is formed through the fuel spray guns positioned on the periphery of the injection system, the film was experimentally examined, it is formed by liquid that seeped through the inner wall of the combustion chamber. The parameter used for validation and refinement of the theoretical penetration of the film was cooling, as this parameter is of paramount importance to obtain an efficient thermal protection inside the combustion chamber. Cold tests confirmed a penetrating cold enough cooling of the film for the length of the combustion chamber of the studied engine.

Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article

Science.gov (United States)

Hulka, J. R.; Protz, C. S.; Garcia, C. P.; Casiano, M. J.; Parton, J. A.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for

Low NOx combustion technologies for high-temperature natural gas combustion

International Nuclear Information System (INIS)

Flamme, Michael

1999-01-01

Because of the high process temperature which is required for some processes like glass melting and the high temperature to which the combustion air is preheated, NOx emission are extremely high. Even at these high temperatures, NOx emissions could be reduced drastically by using advanced combustion techniques such as staged combustion or flame-less oxidation, as experimental work has shown. In the case of oxy-fuel combustion, the NOx emission are also very high if conventional burners are used. The new combustion techniques achieve similar NOx reductions. (author)

Affordable Development and Demonstration of a Small Nuclear Thermal Rocket (NTR) Engine and Stage: How Small Is Big Enough?

Science.gov (United States)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2016-01-01

89-centimeters) -long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 thermal megawatts of power. To reduce the cost of the FTD mission, a simple one-burn lunar flyby mission was considered to reduce the liquid hydrogen (LH2) propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids (NEA), and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

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

Science.gov (United States)

Williams, R. W. (Compiler)

1996-01-01

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

Analysis of pressure blips in aft-finocyl solid rocket motor

Science.gov (United States)

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

2016-07-01

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

IEA low NOx combustion project Stage III. Low NOx combustion and sorbent injection demonstration projects. V.2

International Nuclear Information System (INIS)

Payne, R.

1991-03-01

This report summarizes the main results from an IES project concerning the demonstration of low-NO x combustion and sorbent injection as techniques for the control of NO x and SO x emissions from pulverized coal fired utility boilers. The project has built upon information generated in two previous stages of activity, where NO x and SO x control processes were evaluated at both fundamental and pilot-scales. The concept for this stage of the project was for a unique collaboration, where the participating countries (Canada, Denmark and Sweden, together with the United States) have pooled information from full scale boiler demonstrations of low-NO x burner and sorbent injection technologies, and have jointly contributed to establishing a common basis for data evaluation. Demonstration testing was successfully carried out on five wall-fired commercial boiler systems which ranged in size from a 20 MW thermal input boiler used for district heating, up to a 300 MW electric utility boiler. All of these units were fired on high-volatile bituminous coals with sulfur contents ranging from 0.6-3.2 percent. At each site the existing burners were either modified or replaced to provide for low-NO x combustion, and provisions were made to inject calcium based sorbent materials into the furnace space for SO 2 emission control. The results of sorbent injection testing showed moderate levels of SO 2 removal which ranged from approximately 15 to 55 percent at an injected calcium to sulfur molar ratio to 2.0 and with boiler operation at nominal full load. Sulfur capture was found to depend upon the combined effects of parameters such as: sorbent type and reactivity; peak sorbent temperature; coal sulfur content; and the thermal characteristics of the boilers. (8 refs., 58 figs., 6 tabs.)

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

Science.gov (United States)

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

2012-01-01

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

Lightweight Ultrahigh Temperature CMC-Lined C/C Combustion Chambers, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — NASA and DoD are seeking high-performance, lightweight liquid rocket combustion chambers with future performance goals that cannot be achieved using state-of-the-art...

Polymer Combustion as a Basis for Hybrid Propulsion: A Comprehensive Review and New Numerical Approaches

Directory of Open Access Journals (Sweden)

Vasily Novozhilov

2011-10-01

Full Text Available Hybrid Propulsion is an attractive alternative to conventional liquid and solid rocket motors. This is an active area of research and technological developments. Potential wide application of Hybrid Engines opens the possibility for safer and more flexible space vehicle launching and manoeuvring. The present paper discusses fundamental combustion issues related to further development of Hybrid Rockets. The emphasis is made on the two aspects: (1 properties of potential polymeric fuels, and their modification, and (2 implementation of comprehensive CFD models for combustion in Hybrid Engines. Fundamentals of polymeric fuel combustion are discussed. Further, steps necessary to accurately describe their burning behaviour by means of CFD models are investigated. Final part of the paper presents results of preliminary CFD simulations of fuel burning process in Hybrid Engine using a simplified set-up.

THREE-STAGE COMBUSTION (REBURNING) ON A FULL SCALE OPERATING BOILER IN THE U.S.S.R.

Science.gov (United States)

The report gives results of a program to complete preliminary design of a three- stage combustion (reburn) system for nitrogen oxide (NOx) emissions control on an operating boiler in the U. S.S. R. he program to design the reburn system consisted of five tasks: visiting the Ladyz...

Evaluation of undeveloped rocket engine cycle applications to advanced transportation

Science.gov (United States)

1990-01-01

Undeveloped pump-fed, liquid propellant rocket engine cycles were assessed and evaluated for application to Next Manned Transportation System (NMTS) vehicles, which would include the evolving Space Transportation System (STS Evolution), the Personnel Launch System (PLS), and the Advanced Manned Launch System (AMLS). Undeveloped engine cycles selected for further analysis had potential for increased reliability, more maintainability, reduced cost, and improved (or possibly level) performance when compared to the existing SSME and proposed STME engines. The split expander (SX) cycle, the full flow staged combustion (FFSC) cycle, and a hybrid version of the FFSC, which has a LOX expander drive for the LOX pump, were selected for definition and analysis. Technology requirements and issues were identified and analyses of vehicle systems weight deltas using the SX and FFSC cycles in AMLS vehicles were performed. A strawman schedule and cost estimate for FFSC subsystem technology developments and integrated engine system demonstration was also provided.

Large Liquid Rocket Testing: Strategies and Challenges

Science.gov (United States)

Rahman, Shamim A.; Hebert, Bartt J.

2005-01-01

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

Infrared signature modelling of a rocket jet plume - comparison with flight measurements

International Nuclear Information System (INIS)

Rialland, V; Perez, P; Roblin, A; Guy, A; Gueyffier, D; Smithson, T

2016-01-01

The infrared signature modelling of rocket plumes is a challenging problem involving rocket geometry, propellant composition, combustion modelling, trajectory calculations, fluid mechanics, atmosphere modelling, calculation of gas and particles radiative properties and of radiative transfer through the atmosphere. This paper presents ONERA simulation tools chained together to achieve infrared signature prediction, and the comparison of the estimated and measured signatures of an in-flight rocket plume. We consider the case of a solid rocket motor with aluminized propellant, the Black Brant sounding rocket. The calculation case reproduces the conditions of an experimental rocket launch, performed at White Sands in 1997, for which we obtained high quality infrared signature data sets from DRDC Valcartier. The jet plume is calculated using an in-house CFD software called CEDRE. The plume infrared signature is then computed on the spectral interval 1900-5000 cm -1 with a step of 5 cm -1 . The models and their hypotheses are presented and discussed. Then the resulting plume properties, radiance and spectra are detailed. Finally, the estimated infrared signature is compared with the spectral imaging measurements. The discrepancies are analyzed and discussed. (paper)

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

Science.gov (United States)

Yen, Yi-Hsin

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

The second stage of a Titan II rocket is lifted for mating at the launch tower, Vandenberg AFB

Science.gov (United States)

2000-01-01

At the launch tower, Vandenberg Air Force Base, Calif., the second stage of a Titan II rocket is lifted to vertical. The Titan will power the launch of a National Oceanic and Atmospheric Administration (NOAA-L) satellite scheduled no earlier than Sept. 12. NOAA-L is part of the Polar-Orbiting Operational Environmental Satellite (POES) program that provides atmospheric measurements of temperature, humidity, ozone and cloud images, tracking weather patterns that affect the global weather and climate. Hybrid rocket engine research program at Ryerson University

Energy Technology Data Exchange (ETDEWEB)

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

2007-07-01

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

This Is Rocket Science!

Science.gov (United States)

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-09-01

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

Combustion Stability Verification for the Thrust Chamber Assembly of J-2X Developmental Engines 10001, 10002, and 10003

Science.gov (United States)

Morgan, C. J.; Hulka, J. R.; Casiano, M. J.; Kenny, R. J.; Hinerman, T. D.; Scholten, N.

2015-01-01

The J-2X engine, a liquid oxygen/liquid hydrogen propellant rocket engine available for future use on the upper stage of the Space Launch System vehicle, has completed testing of three developmental engines at NASA Stennis Space Center. Twenty-one tests of engine E10001 were conducted from June 2011 through September 2012, thirteen tests of the engine E10002 were conducted from February 2013 through September 2013, and twelve tests of engine E10003 were conducted from November 2013 to April 2014. Verification of combustion stability of the thrust chamber assembly was conducted by perturbing each of the three developmental engines. The primary mechanism for combustion stability verification was examining the response caused by an artificial perturbation (bomb) in the main combustion chamber, i.e., dynamic combustion stability rating. No dynamic instabilities were observed in the TCA, although a few conditions were not bombed. Additional requirements, included to guard against spontaneous instability or rough combustion, were also investigated. Under certain conditions, discrete responses were observed in the dynamic pressure data. The discrete responses were of low amplitude and posed minimal risk to safe engine operability. Rough combustion analyses showed that all three engines met requirements for broad-banded frequency oscillations. Start and shutdown transient chug oscillations were also examined to assess the overall stability characteristics, with no major issues observed.

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

Directory of Open Access Journals (Sweden)

Si-Yuan Chen

2016-01-01

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

Friction Stir Welding of GR-Cop 84 for Combustion Chamber Liners

Science.gov (United States)

Russell, Carolyn K.; Carter, Robert; Ellis, David L.; Goudy, Richard

2004-01-01

GRCop-84 is a copper-chromium-niobium alloy developed by the Glenn Research Center for liquid rocket engine combustion chamber liners. GRCop-84 exhibits superior properties over conventional copper-base alloys in a liquid hydrogen-oxygen operating environment. The Next Generation Launch Technology program has funded a program to demonstrate scale-up production capabilities of GR-Cop 84 to levels suitable for main combustion chamber production for the prototype rocket engine. This paper describes a novel method of manufacturing the main combustion chamber liner. The process consists of several steps: extrude the GR-Cop 84 powder into billets, roll the billets into plates, bump form the plates into cylinder halves and friction stir weld the halves into a cylinder. The cylinder is then metal spun formed to near net liner dimensions followed by finish machining to the final configuration. This paper describes the friction stir weld process development including tooling and non-destructive inspection techniques, culminating in the successful production of a liner preform completed through spin forming.

Rocket Based Combined Cycle (RBCC) engine inlet

Science.gov (United States)

2004-01-01

Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

Fundamentals of aircraft and rocket propulsion

CERN Document Server

El-Sayed, Ahmed F

2016-01-01

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

Combustion/particle sizing experiments at the Naval Postgraduate School Combustion Research Laboratory

Science.gov (United States)

Powers, John; Netzer, David

1987-01-01

Particle behavior in combustion processes is an active research area at NPS. Currently, four research efforts are being conducted: (1) There is a long standing need to better understand the soot production and combustion processes in gas turbine combustors, both from a concern for improved engine life and to minimize exhaust particulates. Soot emissions are strongly effected by fuel composition and additives; (2) A more recent need for particle sizing/behavior measurements is in the combustor of a solid fuel ramjet which uses a metallized fuel. High speed motion pictures are being used to study rather large burning particles; (3) In solid propellant rocket motors, metals are used to improve specific impulse and/or to provide damping for combustion pressure oscillations. Particle sizing experiments are being conducted using diode arrays to measure the light intensity as a function of scattering angle; (4) Once a good quality hologram is attained, a need exists for obtaining the particle distributions from hologram in a short period of time. A Quantimet 720 Image Analyzer is being used to reconstruct images.

H2-O2 supercritical combustion modeling using a CFD code

Directory of Open Access Journals (Sweden)

Benarous Abdallah

2009-01-01

Full Text Available The characteristics of propellant injection, mixing, and combustion have a profound effect on liquid rocket engine performance. The necessity of raising rocket engines performance requires a combustion chamber operation often in a supercritical regime. A supercritical combustion model based on a one-phase multi-components approach is developed and tested on a non-premixed H2-O2 flame configuration. A two equations turbulence model is used for describing the jet dynamics where a limited Pope correction is added to account for the oxidant spreading rate. Transport properties of the mixture are calculated using extended high pressure forms of the mixing rules. An equilibrium chemistry scheme is adopted in this combustion case, with both algebraic and stochastic expressions for the chemistry/turbulence coupling. The model was incorporated into a computational fluid dynamics commercial code (Fluent 6.2.16. The validity of the present model was investigated by comparing predictions of temperature, species mass fractions, recirculation zones and visible flame length to the experimental data measured on the Mascotte test rig. The results were confronted also with advanced code simulations. It appears that the agreement between the results was fairly good in the chamber regions situated downstream the near injection zone.

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

Science.gov (United States)

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

2014-11-01

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

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

International Nuclear Information System (INIS)

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

2014-01-01

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

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

Science.gov (United States)

Foster, Richard W.

1992-01-01

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

X-ray Radiography Measurements of Shear Coaxial Rocket Injectors

Science.gov (United States)

2013-05-07

Shear coaxial jets can be found in a number of combustion devices – Turbofan engine exhaust , air blast furnaces, and liquid rocket engines ...water and gaseous nitro-gen as propellant simulants at atmospheric backpressure , the effect of momentum flux ratio and mass flux ratio, are...the effect of momentum flux ratio, mass flux ratio and post thickness on the liquid mass distribution – Use quantitative centerline profiles to

Rapid Manufacture of Combustion Chambers Using Ductile, High Strength MMCs (1000-803), Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Triton Systems, Inc. (Triton) proposes to develop a cost-effective manufacturing approach to fabricate combustion chambers for a rocket technology demonstrator...

Experimental study on the influence of oxygen content in the combustion air on the combustion characteristics

International Nuclear Information System (INIS)

Bělohradský, Petr; Skryja, Pavel; Hudák, Igor

2014-01-01

This study was focused on the experimental investigation of the very promising combustion technology called as the oxygen-enhanced combustion (OEC), which uses the oxidant containing higher proportion of oxygen than in the atmospheric air, i.e. more than 21%. The work investigated and compared the characteristics of two OEC methods, namely the premix enrichment and air-oxy/fuel combustion, when the overall oxygen concentration was varied from 21% to 46%. The combustion tests were performed with the experimental two-gas-staged burner of low-NO x type at the burner thermal input of 750 kW for two combustion regimes – one-staged and two-staged combustion. The oxygen concentration in the flue gas was maintained in the neighborhood of 3% vol. (on dry basis). The aim of tests was to assess the impact of the oxidant composition, type of OEC method and fuel-staging on the characteristic combustion parameters in detail. The investigated parameters included the concentration of nitrogen oxides (NO x ) in the flue gas, flue gas temperature, heat flux to the combustion chamber wall, and lastly the stability, shape and dimensions of flame. It was observed that NO x emission is significantly lower when the air-oxy/fuel method is used compared to the premix enrichment method. Moreover, when the fuel was staged, NO x emission was below 120 mg/Nm 3 at all investigated oxygen flow rates. Increasing oxygen concentration resulted in higher heating intensity due to higher concentrations of CO 2 and H 2 O. The available heat at 46% O 2 was higher by 20% compared with that at 21% O 2 . - Highlights: • Premix-enrichment and air-oxy/fuel combustion methods were experimentally studied. • NO x increased sharply as oxygen concentration increased during PE tests. • NO x was below 120 mg/Nm 3 for all investigated oxygen flow rates in AO tests. • Radiative heat transfer was enhanced ca. 20% as O 2 concentration was increased. • OEC flames were observed stable, more luminous and

Investigation of Cooling Water Injection into Supersonic Rocket Engine Exhaust

Science.gov (United States)

Jones, Hansen; Jeansonne, Christopher; Menon, Shyam

2017-11-01

Water spray cooling of the exhaust plume from a rocket undergoing static testing is critical in preventing thermal wear of the test stand structure, and suppressing the acoustic noise signature. A scaled test facility has been developed that utilizes non-intrusive diagnostic techniques including Focusing Color Schlieren (FCS) and Phase Doppler Particle Anemometry (PDPA) to examine the interaction of a pressure-fed water jet with a supersonic flow of compressed air. FCS is used to visually assess the interaction of the water jet with the strong density gradients in the supersonic air flow. PDPA is used in conjunction to gain statistical information regarding water droplet size and velocity as the jet is broken up. Measurement results, along with numerical simulations and jet penetration models are used to explain the observed phenomena. Following the cold flow testing campaign a scaled hybrid rocket engine will be constructed to continue tests in a combusting flow environment similar to that generated by the rocket engines tested at NASA facilities. LaSPACE.

Internal Heterogeneous Processes in Aluminum Combustion

Science.gov (United States)

Dreizin, E. L.

1999-01-01

This paper discusses the aluminum particle combustion mechanism which has been expanded by inclusion of gas dissolution processes and ensuing internal phase transformations. This mechanism is proposed based on recent normal and microgravity experiments with particles formed and ignited in a pulsed micro-arc. Recent experimental findings on the three stages observed in Al particle combustion in air and shows the burning particle radiation, trajectory (streak), smoke cloud shapes, and quenched particle interiors are summarized. During stage I, the radiation trace is smooth and the particle flame is spherically symmetric. The temperature measured using a three-color pyrometer is close to 3000 K. Because it exceeds the aluminum boiling point (2730 K), this temperature most likely characterizes the vapor phase flame zone rather than the aluminum surface. The dissolved oxygen content within particles quenched during stage I was below the detection sensitivity (about 1 atomic %) for Wavelength Dispersive Spectroscopy (WDS). After an increase in the radiation intensity (and simultaneous decrease in the measured color temperature from about 3000 to 2800 K) indicative of the transition to stage II combustion, the internal compositions of the quenched particles change. Both oxygen-rich (approx. 10 atomic %) and oxygen-lean (combustion behavior and the evolution of its internal composition, the change from the spherically symmetric to asymmetric flame shape occurring upon the transition from stage I to stage II combustion could not be understood based only on the fact that dissolved oxygen is detected in the particles. The connection between the two phenomena appeared even less significant because in earlier aluminum combustion studies carried in O2/Ar mixtures, flame asymmetry was not observed as opposed to experiments in air or O2/CO mixtures. It has been proposed that the presence of other gases, i.e., hydrogen, or nitrogen causes the change in the combustion regime.

Ablative Material Testing at Lewis Rocket Lab

Science.gov (United States)

1997-01-01

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

Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

Science.gov (United States)

Elliott, T. S.; Majdalani, J.

2014-11-01

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

Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

International Nuclear Information System (INIS)

Elliott, T S; Majdalani, J

2014-01-01

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

Online dynamic flight optimisation applied to guidance of a variable-flow ducted rocket

NARCIS (Netherlands)

Halswijk, W.H.C.

2009-01-01

The Variable-Flow Ducted Rocket (VFDR) is a type of ramjet that can control the fuel mass flow to the combustion chamber. It combines the high efficiency at high-speed of ramjets with the throttlability of turbofans, and this makes VFDR propulsion an excellent choice for high speed, long range

Metallic Hydrogen: A Game Changing Rocket Propellant

Science.gov (United States)

Silvera, Isaac F.

2016-01-01

The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

Two-stage Lagrangian modeling of ignition processes in ignition quality tester and constant volume combustion chambers

KAUST Repository

Alfazazi, Adamu

2016-08-10

The ignition characteristics of isooctane and n-heptane in an ignition quality tester (IQT) were simulated using a two-stage Lagrangian (TSL) model, which is a zero-dimensional (0-D) reactor network method. The TSL model was also used to simulate the ignition delay of n-dodecane and n-heptane in a constant volume combustion chamber (CVCC), which is archived in the engine combustion network (ECN) library (http://www.ca.sandia.gov/ecn). A detailed chemical kinetic model for gasoline surrogates from the Lawrence Livermore National Laboratory (LLNL) was utilized for the simulation of n-heptane and isooctane. Additional simulations were performed using an optimized gasoline surrogate mechanism from RWTH Aachen University. Validations of the simulated data were also performed with experimental results from an IQT at KAUST. For simulation of n-dodecane in the CVCC, two n-dodecane kinetic models from the literature were utilized. The primary aim of this study is to test the ability of TSL to replicate ignition timings in the IQT and the CVCC. The agreement between the model and the experiment is acceptable except for isooctane in the IQT and n-heptane and n-dodecane in the CVCC. The ability of the simulations to replicate observable trends in ignition delay times with regard to changes in ambient temperature and pressure allows the model to provide insights into the reactions contributing towards ignition. Thus, the TSL model was further employed to investigate the physical and chemical processes responsible for controlling the overall ignition under various conditions. The effects of exothermicity, ambient pressure, and ambient oxygen concentration on first stage ignition were also studied. Increasing ambient pressure and oxygen concentration was found to shorten the overall ignition delay time, but does not affect the timing of the first stage ignition. Additionally, the temperature at the end of the first stage ignition was found to increase at higher ambient pressure

Numerical simulation of divergent rocket-based-combined-cycle performances under the flight condition of Mach 3

Science.gov (United States)

Cui, Peng; Xu, WanWu; Li, Qinglian

2018-01-01

Currently, the upper operating limit of the turbine engine is Mach 2+, and the lower limit of the dual-mode scramjet is Mach 4. Therefore no single power systems can operate within the range between Mach 2 + and Mach 4. By using ejector rockets, Rocket-based-combined-cycle can work well in the above scope. As the key component of Rocket-based-combined-cycle, the ejector rocket has significant influence on Rocket-based-combined-cycle performance. Research on the influence of rocket parameters on Rocket-based-combined-cycle in the speed range of Mach 2 + to Mach 4 is scarce. In the present study, influences of Mach number and total pressure of the ejector rocket on Rocket-based-combined-cycle were analyzed numerically. Due to the significant effects of the flight conditions and the Rocket-based-combined-cycle configuration on Rocket-based-combined-cycle performances, flight altitude, flight Mach number, and divergence ratio were also considered. The simulation results indicate that matching lower altitude with higher flight Mach numbers can increase Rocket-based-combined-cycle thrust. For another thing, with an increase of the divergent ratio, the effect of the divergent configuration will strengthen and there is a limit on the divergent ratio. When the divergent ratio is greater than the limit, the effect of divergent configuration will gradually exceed that of combustion on supersonic flows. Further increases in the divergent ratio will decrease Rocket-based-combined-cycle thrust.

Code Validation of CFD Heat Transfer Models for Liquid Rocket Engine Combustion Devices

National Research Council Canada - National Science Library

Coy, E. B

2007-01-01

.... The design of the rig and its capabilities are described. A second objective of the test rig is to provide CFD validation data under conditions relevant to liquid rocket engine thrust chambers...

Agglomerates, smoke oxide particles, and carbon inclusions in condensed combustion products of an aluminized GAP-based propellant

Science.gov (United States)

Ao, Wen; Liu, Peijin; Yang, Wenjing

2016-12-01

In solid propellants, aluminum is widely used to improve the performance, however the condensed combustion products especially the large agglomerates generated from aluminum combustion significantly affect the combustion and internal flow inside the solid rocket motor. To clarify the properties of the condensed combustion products of aluminized propellants, a constant-pressure quench vessel was adopted to collect the combustion products. The morphology and chemical compositions of the collected products, were then studied by using scanning electron microscopy coupled with energy dispersive (SEM-EDS) method. Various structures have been observed in the condensed combustion products. Apart from the typical agglomerates or smoke oxide particles observed before, new structures including the smoke oxide clusters, irregular agglomerates and carbon-inclusions are discovered and investigated. Smoke oxide particles have the highest amount in the products. The highly dispersed oxide particle is spherical with very smooth surface and is on the order of 1-2 μm, but due to the high temperature and long residence time, these small particles will aggregate into smoke oxide clusters which are much larger than the initial particles. Three types of spherical agglomerates have been found. As the ambient gas temperature is much higher than the boiling point of Al2O3, the condensation layer inside which the aluminum drop is burning would evaporate quickly, which result in the fact that few "hollow agglomerates" has been found compared to "cap agglomerates" and "solid agglomerates". Irregular agglomerates usually larger than spherical agglomerates. The formation of irregular agglomerates likely happens by three stages: deformation of spherical aluminum drops; combination of particles with various shape; finally production of irregular agglomerates. EDS results show the ratio of O to Al on the surface of agglomerates is lower in comparison to smoke oxide particles. C and O account for

Investigating co-combustion characteristics of bamboo and wood.

Science.gov (United States)

Liang, Fang; Wang, Ruijuan; Jiang, Changle; Yang, Xiaomeng; Zhang, Tao; Hu, Wanhe; Mi, Bingbing; Liu, Zhijia

2017-11-01

To investigate co-combustion characteristics of bamboo and wood, moso bamboo and masson pine were torrefied and mixed with different blend ratios. The combustion process was examined by thermogravimetric analyzer (TGA). The results showed the combustion process of samples included volatile emission and oxidation combustion as well as char combustion. The main mass loss of biomass blends occurred at volatile emission and oxidation combustion stage, while that of torrefied biomass occurred at char combustion stage. With the increase of bamboo content, characteristic temperatures decreased. Compared with untreated biomass, torrefied biomass had a higher initial and burnout temperature. With the increase of heating rates, combustion process of samples shifted to higher temperatures. Compared with non-isothermal models, activation energy obtained from isothermal model was lower. The result is helpful to promote development of co-combustion of bamboo and masson pine wastes. Copyright © 2017 Elsevier Ltd. All rights reserved.

Development of the Hawk/Nike Hawk sounding rocket vehicles

Science.gov (United States)

Flowers, B. J.

1976-01-01

A new sounding rocket family, the Hawk and Nike-Hawk Vehicles, have been developed, flight tested and added to the NASA Sounding Rocket Vehicle Stable. The Hawk is a single-stage vehicle that will carry 35.6 cm diameter payloads weighing 45.5 kg to 91 kg to altitudes of 78 km to 56 km, respectively. The two-stage Nike-Hawk will carry payloads weighing 68 kg to 136 kg to altitudes of 118 km to 113 km, respectively. Both vehicles utilize the XM22E8 Hawk rocket motor which is available in large numbers as a surplus item from the U.S. Army. The Hawk fin and tail can hardware were designed in-house. The Nike tail can and fin hardware are surplus Nike-Ajax booster hardware. Development objectives were to provide a vehicle family with a larger diameter, larger volume payload capability than the Nike-Apache and Nike-Tomahawk vehicles at comparable cost. Both vehicles performed nominally in flight tests.

The effects of staged gas combustion on NO{sub x} formation; L'effet de la combustion etagee sur la formation de NO{sub x}

Energy Technology Data Exchange (ETDEWEB)

Wopera, A. [University of Miskolc (Hungary); Sandor, P.; Sevcsik, M. [Dunaferr Co. (Hungary)

2000-07-01

The use of energy involves air pollution unavoidably. As a result of the ongoing change of the structure of the fuels used in Hungary, the decisive majority of the heat treatment and annealing furnaces in our industrial companies operate with natural gas. With natural gas heating if the firing control is proper, only one important air pollutant, the nitrogen monoxide comes into being. NO{sub x} emission of firing processes may be reduced by staged fuel combustion. One of the traditional activities in the field of energetics is the utilization of the energy content of waste fuels. The recovery of the energy content of waste gases, which can be used as fuels, should be examined from both a heat technology and an environment protection point of view. There is the possibility of technological heat utilization when it is mixed with gases of high thermal value. Its utilization as combination gas also makes it necessary to examine how the waste gases change the polluting effects of natural gas. Our research work studied the NO{sub x} formation at different rate of mixing of natural gas and waste gases and the effect of secondary combustion chamber at a two-stage burner. (authors)

Gas-Generator Augmented Expander Cycle Rocket Engine

Science.gov (United States)

Greene, William D. (Inventor)

2011-01-01

An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.

Integral performance optimum design for multistage solid propellant rocket motors

Energy Technology Data Exchange (ETDEWEB)

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

1989-04-01

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

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

Science.gov (United States)

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

2018-05-01

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

Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Science.gov (United States)

Morris, Christopher I.

2005-01-01

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

Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems

Science.gov (United States)

Ryzhkov, V.; Lapshin, E.

2018-01-01

The paper presents an algorithm for designing liquid small thrust rocket engine, the process of which consists of five aggregated stages with feedback. Three stages of the algorithm provide engineering support for design, and two stages - the actual engine design. A distinctive feature of the proposed approach is a deep study of the main technical solutions at the stage of engineering analysis and interaction with the created knowledge (data) base, which accelerates the process and provides enhanced design quality. The using multifunctional graphic package Siemens NX allows to obtain the final product -rocket engine and a set of design documentation in a fairly short time; the engine design does not require a long experimental development.

Characterising Combustion in Diesel Engines : Using parameterised finite stage cylinder process models

NARCIS (Netherlands)

Ding, Y.

2011-01-01

Characterising combustion of diesel engines is not only necessary when researching the instantaneous combustion phenomena but also when investigating the change of the combustion process under variable engine operating conditions. An effective way to achieve this goal is to parameterize the

Sounding rocket experiments during the IMS period at Syowa Station, Antarctica

International Nuclear Information System (INIS)

Hirasawa, T.; Nagata, T.

1979-01-01

During IMS Period, 19 sounding rockets were launched into auroras at various stages of polar substorms from Syowa Station (Geomag. lat. = -69.6 0 , Geomag. log. = 77.1 0 ), Antarctica. Through the successful rocket flights, the significant physical quantities in auroras were obtained: 19 profiles of electron density and temperature, 11 energy spectra of precipitating electrons, 15 frequency spectra of VLF and HF plasma waves and 4 vertical profiles of electric and magnetic fields. These rocket data have been analyzed and compared with the coordinated ground-based observation data for studies of polar substorms. (author)

Initial risk assessment for a single stage to orbit nuclear thermal rocket

Energy Technology Data Exchange (ETDEWEB)

Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

2015-06-15

Highlights: • The risks posed by the surface launch of a nuclear thermal rocket are considered. • Radiation exposure at the public viewing distance is insignificant. • Production of fission products and actinides during launch is limited. • The production of activated argon around the rocket may be a significant concern. - Abstract: In order to consider the possibility of a nuclear thermal rocket (NTR) ground launch, it is necessary to evaluate the risks from such a launch. This includes analysis of the radiation dose rate around the rocket, determining the rate of activation of the materials near the launch, and considering the radionuclides present in the core after the launch. This paper evaluates the potential risk of the NTR ground launch for a range of payloads from 1 to 15 metric tons (MT) using three NTR reactor cores (40, 80, and 120 cm in length) designed in a previous study, based on data produced by MCNP5 and MCNPX models. At the same power level, the 40 cm core length reactor results in the lowest radiation dose rate of the three reactors. Radiation dose rates decrease to background levels 3.5 km from the launch site. After a 1-year decay time, all of the activated materials produced by an NTR launch would be classified as Class A low-level waste. The activation of air produces significant amounts of argon-41 and nitrogen-16 within 100 m of the launch. The derived air concentration (DAC) ratio of the activation products decays to less than unity within 2 days, with only argon-41 remaining. After 10 min of full power operation, the 120 cm core for a 15 MT payload contains 2.5 × 10{sup 13}, 1.4 × 10{sup 12} and 1.5 × 10{sup 12} Bq of {sup 131}I, {sup 137}Cs, and {sup 90}Sr, respectively. The decay heat after shutdown increases with increasing reactor power with a maximum decay heat of 108 kW immediately after shutdown for the 15 MT payload.

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

Science.gov (United States)

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

1989-01-01

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

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

Science.gov (United States)

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

1996-11-01

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

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

Science.gov (United States)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

Data.gov (United States)

National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

Reduction of SO sub 2 emission from a fluidized-bed under staged combustion by coarse limestone

International Nuclear Information System (INIS)

Khan, W.Z.; Gibbbs, B.M.

1998-01-01

A study was performed to investigate the effect of course limestone on the reduction of SO sub 2 emission from a fluidized-bed under staged combustion. The limestone of 1.2-2.5 mm size was premixed with fine coal of 1-3 mm size and fed under bed. The staging levels, bed height, and Ca/S ratio was fixed at 70:30, 30 cm and 3:1, respectively. A maximum of around 50% reduction in SO sub 2 emissions was achieved at both excess air of 20 and 40% and at 830 deg. C. bed temperature. The SO sub 2 emissions were very sensitive to bed temperature. The course limestone was found better in desulfurization efficiency at lower temperature than fine limestone. (author)

An Assessment of Combustion Dynamics in a Low-Nox, Second-Generation Swirl-Venturi Lean Direct Injection Combustion Concept

Science.gov (United States)

Tacina, K. M.; Chang, C. T.; Lee, P.; Mongia, H.; Podboy, D. P.; Dam, B.

2015-01-01

Dynamic pressure measurements were taken during flame-tube emissions testing of three second-generation swirl-venturi lean direct injection (SV-LDI) combustor configurations. These measurements show that combustion dynamics were typically small. However, a small number of points showed high combustion dynamics, with peak-to-peak dynamic pressure fluctuations above 0.5 psi. High combustion dynamics occurred at low inlet temperatures in all three SV-LDI configurations, so combustion dynamics were explored further at low temperature conditions. A point with greater than 1.5 psi peak-to-peak dynamic pressure fluctuations was identified at an inlet temperature of 450!F, a pressure of 100 psia, an air pressure drop of 3%, and an overall equivalence ratio of 0.35. This is an off design condition: the temperature and pressure are typical of 7% power conditions, but the equivalence ratio is high. At this condition, the combustion dynamics depended strongly on the fuel staging. Combustion dynamics could be reduced significantly without changing the overall equivalence ratio by shifting the fuel distribution between stages. Shifting the fuel distribution also decreased NOx emissions.

Cyclic combustion driven shocks in a sompressible vortex as a source of MHD power

Energy Technology Data Exchange (ETDEWEB)

Jenkins, J. M.; Swithenbank, J.

1963-04-15

Theory and experimental data are presented on the properties of cyclic, combustion-driven shocks in a compressible vortex. In this system, part of the chemical energy is converted directly into kinetic energy. Because the temperatures encountered in these combustions approach those obtained in detonations, it is suggested that the cycliccombustion system might be suitable for the production of magnetohydrodynamic power. The experimental data presented are based on observations of rocket motor characteristics. Effects of instabilities are discussed. (T.F.H.)

Unsteady Motions in Combustion Chambers for Propulsion Systems

Science.gov (United States)

2006-12-01

active in wind musical instruments. In all such cases, °ow separation is involved, followed by instability of a shear layer and formation of vortices... musical instruments. The idea that vortices might be responsible for oscillations in a solid propellant rocket seems to have been proposed ¯rst by...Combustion Oscillation in a Ducted Premixed Flame," Inst. Mech. Engineers, Int. Conf. on Comb. in Eng., Oxford, pp. 85{94. Campos -Delgado, D.V., Scheuermans

Assessing Spontaneous Combustion Instability with Nonlinear Time Series Analysis

Science.gov (United States)

Eberhart, C. J.; Casiano, M. J.

2015-01-01

Considerable interest lies in the ability to characterize the onset of spontaneous instabilities within liquid propellant rocket engine (LPRE) combustion devices. Linear techniques, such as fast Fourier transforms, various correlation parameters, and critical damping parameters, have been used at great length for over fifty years. Recently, nonlinear time series methods have been applied to deduce information pertaining to instability incipiency hidden in seemingly stochastic combustion noise. A technique commonly used in biological sciences known as the Multifractal Detrended Fluctuation Analysis has been extended to the combustion dynamics field, and is introduced here as a data analysis approach complementary to linear ones. Advancing, a modified technique is leveraged to extract artifacts of impending combustion instability that present themselves a priori growth to limit cycle amplitudes. Analysis is demonstrated on data from J-2X gas generator testing during which a distinct spontaneous instability was observed. Comparisons are made to previous work wherein the data were characterized using linear approaches. Verification of the technique is performed by examining idealized signals and comparing two separate, independently developed tools.

Rocket science

International Nuclear Information System (INIS)

Upson Sandra

2011-01-01

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

Air-Powered Rockets.

Science.gov (United States)

Rodriguez, Charley; Raynovic, Jim

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

Fundamental rocket injector/spray programs at the Phillips Laboratory

Science.gov (United States)

Talley, D. G.

1993-11-01

The performance and stability of liquid rocket engines is determined to a large degree by atomization, mixing, and combustion processes. Control over these processes is exerted through the design of the injector. Injectors in liquid rocket engines are called upon to perform many functions. They must first of all mix the propellants to provide suitable performance in the shortest possible length. For main injectors, this is driven by the tradeoff between the combustion chamber performance, stability, efficiency, and its weight and cost. In gas generators and preburners, however, it is also driven by the possibility of damage to downstream components, for example piping and turbine blades. This can occur if unburned fuel and oxidant later react to create hot spots. Weight and cost considerations require that the injector design be simple and lightweight. For reusable engines, the injectors must also be durable and easily maintained. Suitable atomization and mixing must be produced with as small a pressure drop as possible, so that the size and weight of pressure vessels and turbomachinery can be minimized. However, the pressure drop must not be so small as to promote feed system coupled instabilities. Another important function of the injectors is to ensure that the injector face plate and the chamber and nozzle walls are not damaged. Typically this requires reducing the heat transfer to an acceptable level and also keeping unburned oxygen from chemically attacking the walls, particularly in reusable engines. Therefore the mixing distribution is often tailored to be fuel-rich near the walls. Wall heat transfer can become catastrophically damaging in the presence of acoustic instabilities, so the injector must prevent these from occurring at all costs. In addition to acoustic stability (but coupled with it), injectors must also be kinetically stable. That is, the flame itself must maintain ignition in the combustion chamber. This is not typically a problem with main

Wavelength modulation spectroscopy near 5 μm for carbon monoxide sensing in a high-pressure kerosene-fueled liquid rocket combustor

Science.gov (United States)

Lee, Daniel D.; Bendana, Fabio A.; Schumaker, S. Alexander; Spearrin, R. Mitchell

2018-05-01

A laser absorption sensor was developed for carbon monoxide (CO) sensing in high-pressure, fuel-rich combustion gases associated with the internal conditions of hydrocarbon-fueled liquid bipropellant rockets. An absorption feature near 4.98 μm, comprised primarily of two rovibrational lines from the P-branch of the fundamental band, was selected to minimize temperature sensitivity and spectral interference with other combustion gas species at the extreme temperatures (> 3000 K) and pressures (> 50 atm) in the combustion chamber environment. A scanned wavelength modulation spectroscopy technique (1 f-normalized 2 f detection) is utilized to infer species concentration from CO absorption, and mitigate the influence of non-absorption transmission losses and noise associated with the harsh sooting combustor environment. To implement the sensing strategy, a continuous-wave distributed-feedback (DFB) quantum cascade laser (QCL) was coupled to a hollow-core optical fiber for remote mid-infrared light delivery to the test article, with high-bandwidth light detection by a direct-mounted photovoltaic detector. The method was demonstrated to measure time-resolved CO mole fraction over a range of oxidizer-to-fuel ratios and pressures (20-70 atm) in a single-element-injector RP-2-GOx rocket combustor.

Acoustic Pressure Oscillations Induced in I-Burner

Science.gov (United States)

Matsui, Kiyoshi

Iwama et al. invented the I-burner to investigate acoustic combustion instability in solid-propellant rockets (Proceedings of ICT Conference, 1994, pp. 26-1 26-14). Longitudinal pressure oscillations were induced in the combustion chamber of a thick-walled rocket by combustion of a stepped-perforation grain (I-burner). These oscillations were studied here experimentally. Two I-burners with an internal diameter of 80 mm and a length of 1208 mm or 2240 mm were made. The grain had stepped perforations (20 and 42 mm in diameter and 657 and 160 mm in length, respectively). Longitudinal pressure oscillations always occur in two stages when an HTPB (hydroxyl-terminated polybutadiene)/AP (ammonium perchlorate)/aluminum-powder propellant burns (54 tests; the highest average pressure in the combustion chamber was 9.5 29 MPa), but no oscillations occur when an HTPB/AP propellant burns (29 tests). The pressure oscillations are essentially linear, but dissipation adds a nonlinear nature to them. In the first stage, the amplitudes are small and the first wave group predominates. In the next stage, the amplitudes are large and many wave groups are present. The change in the grain form accompanying the combustion affects the pressure oscillations.

Experimental determination of plume properties in full-scale hydrogen-oxygen rockets

Science.gov (United States)

Brown, D. G.; Limbaugh, C. C.; Zaccardi, V. A.; Eskridge, R.

1989-01-01

An IR emission/absorption technique for determining radial profiles of static temperature and species partial pressure for cylindrically symmetric combustion gases typical of the effluent of turbine engines and liquid-propellant rockets is described. In the technique, the IR plume radiance and absorption is measured using a 1 x 256-element platinum silicide detector array which is filtered to obtain plume emission measurements in the H2O band near 3.0 microns. A minicomputer is employed to control data acquisition and reduction.

Preliminary Results on the Effects of Distributed Aluminum Combustion Upon Acoustic Growth Rates in a Rijke Burner

OpenAIRE

Newbold, Brian R.

1998-01-01

Distributed particle combustion in solid propellant rocket motors may be a significant cause of acoustic combustion instability. A Rijke burner has been developed as a tool to investigate the phenomenon. Previous improvements and characterization of the upright burner lead to the addition of a particle injection flame. The injector flame increases the burner's acoustic driving by about 10% which is proportional to the injector's additional 2 g/min of gas. Frequency remained fairly constant fo...

Consequences of sludge composition on combustion performance derived from thermogravimetry analysis

Energy Technology Data Exchange (ETDEWEB)

Li, Meiyan; Xiao, Benyi; Wang, Xu; Liu, Junxin, E-mail: jxliu@rcees.ac.cn

2015-01-15

Highlights: • Volatiles, particularly proteins, play a key role in sludge combustion. • Sludge combustion performance varies with different sludge organic concentrations. • Carbohydrates significantly affect the combustion rate in the second stage. • Combustion performance of digested sludge is more negative compared with others. - Abstract: Wastewater treatment plants produce millions of tons of sewage sludge. Sewage sludge is recognized as a promising feedstock for power generation via combustion and can be used for energy crisis adaption. We aimed to investigate the quantitative effects of various sludge characteristics on the overall sludge combustion process performance. Different types of sewage sludge were derived from numerous wastewater treatment plants in Beijing for further thermogravimetric analysis. Thermogravimetric–differential thermogravimetric curves were used to compare the performance of the studied samples. Proximate analytical data, organic compositions, elementary composition, and calorific value of the samples were determined. The relationship between combustion performance and sludge composition was also investigated. Results showed that the performance of sludge combustion was significantly affected by the concentration of protein, which is the main component of volatiles. Carbohydrates and lipids were not correlated with combustion performance, unlike protein. Overall, combustion performance varied with different sludge organic composition. The combustion rate of carbohydrates was higher than those of protein and lipid, and carbohydrate weight loss mainly occurred during the second stage (175–300 °C). Carbohydrates have a substantial effect on the rate of system combustion during the second stage considering the specific combustion feature. Additionally, the combustion performance of digested sewage sludge is more negative than the others.

Consequences of sludge composition on combustion performance derived from thermogravimetry analysis

International Nuclear Information System (INIS)

Li, Meiyan; Xiao, Benyi; Wang, Xu; Liu, Junxin

2015-01-01

Highlights: • Volatiles, particularly proteins, play a key role in sludge combustion. • Sludge combustion performance varies with different sludge organic concentrations. • Carbohydrates significantly affect the combustion rate in the second stage. • Combustion performance of digested sludge is more negative compared with others. - Abstract: Wastewater treatment plants produce millions of tons of sewage sludge. Sewage sludge is recognized as a promising feedstock for power generation via combustion and can be used for energy crisis adaption. We aimed to investigate the quantitative effects of various sludge characteristics on the overall sludge combustion process performance. Different types of sewage sludge were derived from numerous wastewater treatment plants in Beijing for further thermogravimetric analysis. Thermogravimetric–differential thermogravimetric curves were used to compare the performance of the studied samples. Proximate analytical data, organic compositions, elementary composition, and calorific value of the samples were determined. The relationship between combustion performance and sludge composition was also investigated. Results showed that the performance of sludge combustion was significantly affected by the concentration of protein, which is the main component of volatiles. Carbohydrates and lipids were not correlated with combustion performance, unlike protein. Overall, combustion performance varied with different sludge organic composition. The combustion rate of carbohydrates was higher than those of protein and lipid, and carbohydrate weight loss mainly occurred during the second stage (175–300 °C). Carbohydrates have a substantial effect on the rate of system combustion during the second stage considering the specific combustion feature. Additionally, the combustion performance of digested sewage sludge is more negative than the others

Role of additives in combustion waves and effect on stable combustion limit of double-base propellants

Energy Technology Data Exchange (ETDEWEB)

Kubota, N [Japan Defence Agency, Tachikawa. 3. Research Center

1978-12-01

The effect of additives on the flame structures and the burning rates of double-base propellants have been examined by means of photographic observations and temperature profile measurements. The additives used for this study are lead salicylate (PbSa, 2%), nickel (Ni, 1%), ammonium perchlorate (AP, 30%), and cyclotetramethylene tetranitramine (HMX, 30%). The addition of PbSa increases the burning rate, but does not improve the flame temperature characteristics. The addition of Ni increases the flame temperature significantly at pressures below 30 atm. The Ni acts as a catalyst to promote the flame reaction while it does not act as a burning rate modifier. The additions of AP and HMX increase the thermal performance of the propellant system, however, the HMX does not improve the stable combustion limit of the rocket motor at low pressures. The addition of Ni or AP is found to increase the flame temperature at pressures below 30 atm, and the stable combustion limits is lowered to below 3 atm.

Cryogenic rocket engine development at Delft aerospace rocket engineering

NARCIS (Netherlands)

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

2016-01-01

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

Real gas CFD simulations of hydrogen/oxygen supercritical combustion

Science.gov (United States)

Pohl, S.; Jarczyk, M.; Pfitzner, M.; Rogg, B.

2013-03-01

A comprehensive numerical framework has been established to simulate reacting flows under conditions typically encountered in rocket combustion chambers. The model implemented into the commercial CFD Code ANSYS CFX includes appropriate real gas relations based on the volume-corrected Peng-Robinson (PR) equation of state (EOS) for the flow field and a real gas extension of the laminar flamelet combustion model. The results indicate that the real gas relations have a considerably larger impact on the flow field than on the detailed flame structure. Generally, a realistic flame shape could be achieved for the real gas approach compared to experimental data from the Mascotte test rig V03 operated at ONERA when the differential diffusion processes were only considered within the flame zone.

Some Calculated Research Results of the Working Process Parameters of the Low Thrust Rocket Engine Operating on Gaseous Oxygen-Hydrogen Fuel

Science.gov (United States)

Ryzhkov, V.; Morozov, I.

2018-01-01

The paper presents the calculating results of the combustion products parameters in the tract of the low thrust rocket engine with thrust P ∼ 100 N. The article contains the following data: streamlines, distribution of total temperature parameter in the longitudinal section of the engine chamber, static temperature distribution in the cross section of the engine chamber, velocity distribution of the combustion products in the outlet section of the engine nozzle, static temperature near the inner wall of the engine. The presented parameters allow to estimate the efficiency of the mixture formation processes, flow of combustion products in the engine chamber and to estimate the thermal state of the structure.

On the combustion mechanisms of ZrH2 in double-base propellant.

Science.gov (United States)

Yang, Yanjing; Zhao, Fengqi; Yuan, Zhifeng; Wang, Ying; An, Ting; Chen, Xueli; Xuan, Chunlei; Zhang, Jiankan

2017-12-13

Metal hydrides are regarded as a series of promising hydrogen-supplying fuel for solid rocket propellants. Their effects on the energetic and combustion performances of propellants are closely related to their reaction mechanisms. Here we report a first attempt to determine the reaction mechanism of ZrH 2 , a high-density metal hydride, in the combustion of a double-base propellant to evaluate its potential as a fuel. ZrH 2 is determined to possess good resistance to oxidation by nitrocellulose and nitroglycerine. Thus its combustion starts with dehydrogenation to generate H 2 and metallic Zr. Subsequently, the newly formed Zr and H 2 participate in the combustion and, especially, Zr melts and then combusts on the burning surface which favors the heat feedback to the propellant. This phenomenon is completely different from the combustion behavior of the traditional fuel Al, where the Al particles are ejected off the burning surface of the propellant to get into the luminous flame zone to burn. The findings in this work validate the potential of ZrH 2 as a hydrogen-supplying fuel for double-base propellants.

Combustion Dynamics and Stability Modeling of a Liquid Oxygen/RP-2 Oxygen-Rich Staged Combustion Preburner and Thrust Chamber Assembly with Gas-Centered Swirl Coaxial Injector Elements

Science.gov (United States)

Casiano, M. J.; Kenny, R. J.; Protz, C. S.; Garcia, C. P.; Simpson, S. P.; Elmore, J. L.; Fischbach, S. R.; Giacomoni, C. B.; Hulka, J. R.

2016-01-01

The Combustion Stability Tool Development (CSTD) project, funded by the Air Force Space and Missile Systems Center, began in March 2015 supporting a renewed interest in the development of a liquid oxygen/hydrocarbon, oxygen-rich combustion engine. The project encompasses the design, assembly, and hot-fire testing of the NASA Marshall Space Flight Center 40-klbf Integrated Test Rig (MITR). The test rig models a staged-combustion configuration by combining an oxygen-rich preburner (ORPB), to generate hot gas, with a thrust chamber assembly (TCA) using gas-centered swirl coaxial injector elements. There are five separately designed interchangeable injectors in the TCA that each contain 19- or 27- injector elements. A companion paper in this JANNAF conference describes the design characteristics, rationale, and fabrication issues for all the injectors. The data acquired from a heavily instrumented rig encompasses several injectors, several operating points, and stability bomb tests. Another companion paper in this JANNAF conference describes this test program in detail. In this paper, dynamic data from the hot-fire testing is characterized and used to identify the responses in the ORPB and TCA. A brief review of damping metrics are discussed and applied as a measure of stability margin for damped acoustic modes. Chug and longitudinal combustion stability models and predictions are described which includes new dynamic models for compressible flow through an orifice and a modification to incorporate a third feed line for inclusion of the fuel-film coolant. Flow-acoustics finite element modeling is used to investigate the anticipated TCA acoustics, the effects of injector element length on stability margin, and the potential use of an ORPB orifice trip ring for improving longitudinal stability margin.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

Directory of Open Access Journals (Sweden)

Qiang WEI

2017-08-01

Full Text Available To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions. The overall model is benchmarked under various impingement angles, jet momentum and off-center ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

Institute of Scientific and Technical Information of China (English)

Qiang WEI; Guozhu LIANG

2017-01-01

To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray,the conventional uncoupled spray model for impinging injectors is extended by considering the couplingof the jet impingement process and the ambient gas field.The new coupled model consists of the plain-orifice sub-model,the jet-jet impingement sub-model and the droplet collision sub-model.The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles,jet momentum and offcenter ratios.Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics,such as the mass flux and mixture ratio distributions in quiescent air.Besides,impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions.First,a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile.The minimum average droplet diameter is achieved when the orifices work in cavitation state,and is about 30％ smaller than the steady single phase state.Second,the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°.The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

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

Science.gov (United States)

Zibit, Benjamin Seth

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

Major accomplishments of America's nuclear rocket program (ROVER)

International Nuclear Information System (INIS)

Finseth, J.L.

1991-01-01

The United States embarked on a program to develop nuclear rocket engines in 1955. This program was known as project Rover. Initially nuclear rockets were considered as a potential backup for intercontinental ballistic missile propulsion but later proposed applications included both a lunar second stage as well as use in manned-Mars flights. Under the Rover program, 19 different reactors were built and tested during the period of 1959-1969. Additionally, several cold flow (non-fuelled) reactors were tested as well as a nuclear fuels test cell. The Rover program was terminated in 1973, due to budget constraints and an evolving political climate. The Rover program would have led to the development of a flight engine had the program continued through a logical continuation. The Rover program was responsible for a number of technological achievements. The successful operation of nuclear rocket engines on a system level represents the pinnacle of accomplishment. This paper will discuss the engine test program as well as several subsystems

Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

Directory of Open Access Journals (Sweden)

Bruce Chehroudi

2012-01-01

Full Text Available Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME or Vulcain (Ariane 5, the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL, spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

Application of the FIRST Combustion model to Spray Combustion

NARCIS (Netherlands)

de Jager, B.; Kok, Jacobus B.W.

2004-01-01

Liquid fuel is of interest to apply to gas turbines. The large advantage is that liquids are easily storable as compared to gaseous fuels. Disadvantage is that liquid fuel has to be sprayed, vaporized and mixed with air. Combustion occurs at some stage of mixing and ignition. Depending on the

History of Solid Rockets

Science.gov (United States)

Green, Rebecca

2017-01-01

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

Optical measurements in rocket engine liquid sprays

Science.gov (United States)

Feikema, Douglas A.

1994-01-01

The performance of liquid propellant rocket engines is dependent upon many elements of the entire system. One of the most fundamental and most critical is the performance of the injector elements. Their characterization is an important part of the development of combustion devices. Optical measurements within these environments have proven to be invaluable tools in quantifying the physical environment of two phase flows. The effort reported herein involves the measurement of drop velocity, drop size, and most importantly mass flux using Phase-Doppler Particle Anemometry within a spray generated by a single swirl injector element operating in atmospheric pressure conditions. The mass flux has been determined and validated by mechanical patternation methods and by profile integration of the mass flux.

Second stage of Saturn V being assembled with the first stage.

Science.gov (United States)

1965-01-01

The hydrogen-powered second stage is being lowered into place during the final phase of fabrication of the Saturn V moon rocket at North American's Seal Beach, California facility. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

SSTO rockets. A practical possibility

Science.gov (United States)

Bekey, Ivan

1994-07-01

Most experts agree that single-stage-to-orbit (SSTO) rockets would become feasible if more advanced technologies were available to reduce the vehicle dry weight, increase propulsion system performance, or both. However, these technologies are usually judged to be very ambitious and very far off. This notion persists despite major advances in technology and vehicle design in the past decade. There appears to be four major misperceptions about SSTOs, regarding their mass fraction, their presumed inadequate performance margin, their supposedly small payloads, and their extreme sensitivity to unanticipated vehicle weight growth. These misperceptions can be dispelled for SSTO rockets using advanced technologies that could be matured and demonstrated in the near term. These include a graphite-composite primary structure, graphite-composite and Al-Li propellant tanks with integral reusable thermal protection, long-life tripropellant or LOX-hydrogen engines, and several technologies related to operational effectiveness, including vehicle health monitoring, autonomous avionics/flight control, and operable launch and ground handling systems.

Prediction of Non-Equilibrium Kinetics of Fuel-Rich Kerosene/LOX Combustion in Gas Generator

International Nuclear Information System (INIS)

Yu, Jung Min; Lee, Chang Jin

2007-01-01

Gas generator is the device to produce high enthalpy gases needed to drive turbo-pump system in liquid rocket engine. And, the combustion temperature in gas generator should be controlled below around 1,000K to avoid any possible thermal damages to turbine blade by using either fuel rich combustion or oxidizer rich combustion. Thus, nonequilibrium chemical reaction dominates in fuel-rich combustion of gas generator. Meanwhile, kerosene is a compounded fuel with various types of hydrocarbon elements and difficult to model the chemical kinetics. This study focuses on the prediction of the non-equilibrium reaction of fuel rich kerosene/LOX combustion with detailed kinetics developed by Dagaut using PSR (Perfectly Stirred Reactor) assumption. In Dagaut's surrogate model for kerosene, chemical kinetics of kerosene consists of 1,592 reaction steps with 207 chemical species. Also, droplet evaporation time is taken into account in the PSR calculation by changing the residence time of droplet in the gas generator. Frenklach's soot model was implemented along with detailed kinetics to calculate the gas properties of fuel rich combustion efflux. The results could provide very reliable and accurate numbers in the prediction of combustion gas temperature,species fraction and material properties

Design, Development and Hotfire Testing of Monolithic Copper and Bimetallic Additively Manufactured Combustion Chambers

Science.gov (United States)

Gradl, Paul; Barnett, Greg; Brandsmeier, Will; Greene, Sandy Elam; Protz, Chris

2016-01-01

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM) otherwise commonly referred to as additive manufacturing. The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for the GRCop-84 copper-alloy commensurate with powder bed additive manufacturing, evaluate bimetallic deposition and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. As a direct spin off of this program, NASA is working with industry partners to further develop the printing process for the GRCop-84 material in addition to the C-18150 (CuCrZr) material. To advance the process further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic additively manufactured chambers. A 1.2k sized thrust-chamber was designed and developed to compare the printing process of the GRCop-84 and C-18150 SLM materials. A series of similar MCC liners also completed development with an Inconel 625 jacket bonded to the GRcop-84 liner evaluating direct metal deposition (DMD) laser and arc-based techniques. This paper describes the design, development, manufacturing and testing of these combustion chambers and associated lessons learned throughout the design and development process.

2017 Laser Diagnostics in Combustion GRC/GRS

Science.gov (United States)

2017-10-06

scientists and engineers . The goal of GRC was to bring together scientists and engineers with a strong interest in laser-based combustion diagnostics to...processes. The GRS, held the weekend before the GRC, was organized to bring together emerging leaders in the field at an early stage in their careers ...coupled multi-scale combustion processes. The GRC on Laser Diagnostics in Combustion was designed to bring together scientists and engineers working at

ANALYSIS OF INTERNAL COMBUSTION ENGINE WITH A NEW CONCEPT OF POROUS MEDIUM COMBUSTION FOR THE FUTURE CLEAN ENGINE

Directory of Open Access Journals (Sweden)

Ashok A Dhale

2010-01-01

Full Text Available At present, the emissions of internal combustion engine can only be improved by catalytic treatments of the exhaust gases. Such treatments, however, result in high costs and relatively low conversion efficiency. This suggests that a new combustion technique should be developed to yield improved primary combustion processes inside the engine with drastically reduced exhaust gas emissions. To fulfill all requirements, Dr. Franz Drust has proposed a new combustion concept to perform homogenous combustion in internal combustion engines. This concept used the porous medium combustion technique and is called "PM-engine". It is shown that the PM combustion technique can be applied to internal combustion engines. Theoretical considerations are presented for internal combustion engines, indicating that an overall improvement in thermal efficiency can be achieved for the PM-engine. This is explained and general performance of the new PM-engines is demonstrated for a single cylinder, water cooled, direct injection diesel engine. Verification of experiments at primary stage is described that were carried out as a part of the present study.

The flight of uncontrolled rockets

CERN Document Server

Gantmakher, F R; Dryden, H L

1964-01-01

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

Modeling of Combined Heat and Power Plant Based on a Multi-Stage Gasifier and Internal Combustion Engines of Various Power Outputs

Science.gov (United States)

Khudyakova, G. I.; Kozlov, A. N.; Svishchev, D. A.

2017-11-01

The paper is concerned with an integrated system of internal combustion engine and mini combined heat and power plant (ICE-CHP). The system is based on multi-stage wood biomass gasification. The use of producer gas in the system affects negatively the internal combustion engine performance and, therefore, reduces the efficiency of the ICE-CHP plant. A mathematical model of an internal combustion engine running on low-calorie producer gas was developed using an overview of Russian and foreign manufacturers of reciprocating units, that was made in the research. A thermal calculation was done for four-stroke gas engines of different rated power outputs (30, 100 and 250 kW), running on producer gas (CO2 - 10.2, CO - 45.8, N2 - 38.8%). Thermal calculation demonstrates that the engine exhaust gas temperature reaches 500 - 600°C at the rated power level and with the lower engine power, the temperature gets higher. For example, for an internal combustion engine power of 1000 kW the temperature of exhaust gases equals 400°C. A comparison of the efficiency of engine operation on natural gas and producer gas shows that with the use of producer gas the power output declines from 300 to 250 kWe. The reduction in the effective efficiency in this case makes up 2%. The measures are proposed to upgrade the internal combustion engine to enable it to run on low-calorie producer gas.

Feasibility and Performance of the Microwave Thermal Rocket Launcher

Science.gov (United States)

Parkin, Kevin L. G.; Culick, Fred E. C.

2004-03-01

Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit (SSTO) microwave thermal rocket. We present an SSTO concept employing a scaled X-33 aeroshell. The flat aeroshell underside is covered by a thin-layer microwave absorbent heat-exchanger that forms part of the thruster. During ascent, the heat-exchanger faces the microwave beam. A simple ascent trajectory analysis incorporating X-33 aerodynamic data predicts a 10% payload fraction for a 1 ton craft of this type. In contrast, the Saturn V had 3 non-reusable stages and achieved a payload fraction of 4%.

Nuclear rockets

Energy Technology Data Exchange (ETDEWEB)

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

1972-02-01

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

Research Technology (ASTP) Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

Combustion of poultry litter in a fluidised bed combustor

Energy Technology Data Exchange (ETDEWEB)

P. Abelha; I. Gulyurtlu; D. Boavida; J. Seabra Barros; I. Cabrita; J. Leahy; B. Kelleher; M. Leahy [DEECA-INETI, Lisbon (Portugal)

2003-04-01

Combustion studies of poultry litter alone or mixed with peat by 50% on weight basis were undertaken in an atmospheric bubbling fluidised bed. Because of high moisture content of poultry litter, there was some uncertainty whether the combustion could be sustained on 100% poultry litter and as peat is very available in Ireland, its presence was considered to help to improve the combustion. However, the results showed that, as long as the moisture content of poultry litter was kept below 25%, the combustion did not need the addition of peat. The main parameters that were investigated are (i) moisture content, (ii) air staging, and (iii) variations in excess air levels along the freeboard. The main conclusions of the results are (i) combustion was influenced very much by the conditions of the fuel supply, (ii) the steady fuel supply was strongly dependent on the moisture content of the poultry litter, (iii) temperature appeared to be still very influential in reducing the levels of unburned carbon and hydrocarbons released from residues, (iv) the air staging in the freeboard improved combustion efficiency by enhancing the combustion of volatiles released from residues in the riser and (vi) NOx emissions were influenced by air staging in the freeboard. Particles collected from the bed and the two cyclones were analysed to determine the levels of heavy metals and the leachability tests were carried out with ashes collected to verify whether or not they could safely be used in agricultural lands. 8 refs., 1 fig., 8 tabs.

Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion

Science.gov (United States)

Rahman, Shamim A.

2010-01-01

Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.

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

Science.gov (United States)

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

2002-01-01

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

Combustion visualization and experimental study on spark induced compression ignition (SICI) in gasoline HCCI engines

International Nuclear Information System (INIS)

Wang Zhi; He Xu; Wang Jianxin; Shuai Shijin; Xu Fan; Yang Dongbo

2010-01-01

Spark induced compression ignition (SICI) is a relatively new combustion control technology and a promising combustion mode in gasoline engines with high efficiency. SICI can be divided into two categories, SACI and SI-CI. This paper investigated the SICI combustion process using combustion visualization and engine experiment respectively. Ignition process of SICI was captured by high speed photography in an optical engine with different compression ratios. The results show that SICI is a combustion mode combined with partly flame propagation and main auto-ignition. The spark ignites the local mixture near spark electrodes and the flame propagation occurs before the homogeneous mixture is auto-ignited. The heat release from central burned zone due to the flame propagation increases the in-cylinder pressure and temperature, resulting in the unburned mixture auto-ignition. The SICI combustion process can be divided into three stages of the spark induced stage, the flame propagation stage and the compression ignition stage. The SICI combustion mode is different from the spark ignition (SI) knocking in terms of the combustion and emission characteristics. Furthermore, three typical combustion modes including HCCI, SICI, SI, were compared on a gasoline direct injection engine with higher compression ratio and switchable cam-profiles. The results show that SICI has an obvious combustion characteristic with two-stage heat release and lower pressure rise rate. The SICI combustion mode can be controlled by spark timings and EGR rates and utilized as an effective method for high load extension on the gasoline HCCI engine. The maximum IMEP of 0.82 MPa can be achieved with relatively low NO x emission and high thermal efficiency. The SICI combustion mode can be applied in medium-high load region for high efficiency gasoline engines.

Combustion visualization and experimental study on spark induced compression ignition (SICI) in gasoline HCCI engines

Energy Technology Data Exchange (ETDEWEB)

Wang Zhi, E-mail: wangzhi@tsinghua.edu.c [State Key Laboratory of Automotive Safety and Energy, Tsinghua University, Beijing 100084 (China); He Xu; Wang Jianxin; Shuai Shijin; Xu Fan; Yang Dongbo [State Key Laboratory of Automotive Safety and Energy, Tsinghua University, Beijing 100084 (China)

2010-05-15

Spark induced compression ignition (SICI) is a relatively new combustion control technology and a promising combustion mode in gasoline engines with high efficiency. SICI can be divided into two categories, SACI and SI-CI. This paper investigated the SICI combustion process using combustion visualization and engine experiment respectively. Ignition process of SICI was captured by high speed photography in an optical engine with different compression ratios. The results show that SICI is a combustion mode combined with partly flame propagation and main auto-ignition. The spark ignites the local mixture near spark electrodes and the flame propagation occurs before the homogeneous mixture is auto-ignited. The heat release from central burned zone due to the flame propagation increases the in-cylinder pressure and temperature, resulting in the unburned mixture auto-ignition. The SICI combustion process can be divided into three stages of the spark induced stage, the flame propagation stage and the compression ignition stage. The SICI combustion mode is different from the spark ignition (SI) knocking in terms of the combustion and emission characteristics. Furthermore, three typical combustion modes including HCCI, SICI, SI, were compared on a gasoline direct injection engine with higher compression ratio and switchable cam-profiles. The results show that SICI has an obvious combustion characteristic with two-stage heat release and lower pressure rise rate. The SICI combustion mode can be controlled by spark timings and EGR rates and utilized as an effective method for high load extension on the gasoline HCCI engine. The maximum IMEP of 0.82 MPa can be achieved with relatively low NO{sub x} emission and high thermal efficiency. The SICI combustion mode can be applied in medium-high load region for high efficiency gasoline engines.

Effect of Metal Additives on the Combustion Characteristics of High-Energy Materials

Directory of Open Access Journals (Sweden)

Korotkikh Alexander

2016-01-01

Full Text Available Thermodynamic calculation of combustion parameters and equilibrium composition of HEMs combustion products showed, that at the increase of aluminum powder dispersity the specific impulse and combustion temperature of solid propellants are reduced due to the decrease of the mass fraction of active aluminum in particles. Partial or complete replacement of aluminum by metal powder (B, Mg, AlB2, Al\\Mg alloy, Fe, Ti and Zr in HEMs composition leads to the reduce of the specific impulse and combustion temperature. Replacement of aluminum powder by boron and magnesium in HEM reduces the mass fraction of condensed products in the combustion chamber of solid rocket motor. So, for compositions HEMs with boron and aluminum boride the mass fraction in chamber is reduced by 24 and 36 %, respectively, with respect to the composition HEMs with Al powder. But the mass fraction of CCPs in the nozzle exit increases by 13 % for HEMs with aluminum boride due to the formation of boron oxide in the condensed combustion products. Partial replacement of 2 wt. % aluminum powder by iron and copper additives in HEM leads to the reduce of CCPs mass fraction in chamber by 4–10 % depending on the aluminum powder dispersity duo to these metals are not formed condensed products at the HEMs combustion in chamber.

Parametric study and performance analysis of hybrid rocket motors with double-tube configuration

Science.gov (United States)

Yu, Nanjia; Zhao, Bo; Lorente, Arnau Pons; Wang, Jue

2017-03-01

The practical implementation of hybrid rocket motors has historically been hampered by the slow regression rate of the solid fuel. In recent years, the research on advanced injector designs has achieved notable results in the enhancement of the regression rate and combustion efficiency of hybrid rockets. Following this path, this work studies a new configuration called double-tube characterized by injecting the gaseous oxidizer through a head end injector and an inner tube with injector holes distributed along the motor longitudinal axis. This design has demonstrated a significant potential for improving the performance of hybrid rockets by means of a better mixing of the species achieved through a customized injection of the oxidizer. Indeed, the CFD analysis of the double-tube configuration has revealed that this design may increase the regression rate over 50% with respect to the same motor with a conventional axial showerhead injector. However, in order to fully exploit the advantages of the double-tube concept, it is necessary to acquire a deeper understanding of the influence of the different design parameters in the overall performance. In this way, a parametric study is carried out taking into account the variation of the oxidizer mass flux rate, the ratio of oxidizer mass flow rate injected through the inner tube to the total oxidizer mass flow rate, and injection angle. The data for the analysis have been gathered from a large series of three-dimensional numerical simulations that considered the changes in the design parameters. The propellant combination adopted consists of gaseous oxygen as oxidizer and high-density polyethylene as solid fuel. Furthermore, the numerical model comprises Navier-Stokes equations, k-ε turbulence model, eddy-dissipation combustion model and solid-fuel pyrolysis, which is computed through user-defined functions. This numerical model was previously validated by analyzing the computational and experimental results obtained for

Fuel staging tests at the Kymijaervi power plant

International Nuclear Information System (INIS)

Kivelae, M.; Rotter, H.; Virkki, J.

1990-01-01

The aim of this study was to measure nitrogen oxide (NO x ) emissions and find the methods to reduce them in plants using coal and natural gas as fuel. The tests involved were made at the Kymijaervi Power Plant, Lahti, Finland. Coal and natural gas was used alone or mixed. With natural gas when using flue gas recirculation, the NO x emission level dropped from 330 mg/m 3 down to 60 mg/m 3 . A negative side effect was that the flue gas temperature increased. At coal combustion and staged combustion, the flue gas recirculation had no significant effect on the NO x emission level. At coal combustion, the staging of combustion air halved the NO x emission but the combustibles increased strongly. With fuel staging, using coal as main fuel and gas as staging fuel, the NO x emission level was decreased from 340 mg/m 3 to 170 mg/m 3 . At the same time the combustibles increased 2 %- units. Also the flue gas temperature increased a little. At the tests, the proportion of natural gas was rather high, one third of the fuel energy input, but it could not be decreased, because the gas flow ratio was already too low to ensure good mixing

Nuclear rockets

International Nuclear Information System (INIS)

Sarram, M.

1972-01-01

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

Nuclear Propulsion through Direct Conversion of Fusion Energy: The Fusion Driven Rocket

Science.gov (United States)

Slough, John; Pancotti, Anthony; Kirtley, David; Pihl, Christopher; Pfaff, Michael

2012-01-01

The future of manned space exploration and development of space depends critically on the creation of a dramatically more proficient propulsion architecture for in-space transportation. A very persuasive reason for investigating the applicability of nuclear power in rockets is the vast energy density gain of nuclear fuel when compared to chemical combustion energy. Current nuclear fusion efforts have focused on the generation of electric grid power and are wholly inappropriate for space transportation as the application of a reactor based fusion-electric system creates a colossal mass and heat rejection problem for space application.

Study of combustion properties of a solid propellant by highly time-resolved passive FTIR

Energy Technology Data Exchange (ETDEWEB)

Zhang, Liming; Zhang, Lin; Li, Yan; Liu, Bingping; Wang, Junde [Laboratory of Advanced Spectroscopy, Nanjing University of Science and Technology, Nanjing 210014 (China)

2006-10-15

With a time resolution of 0.125 s and a spectral resolution of 4 cm{sup -1}, emission spectra of the combustion process of a solid propellant were recorded by highly time-resolved passive FTIR. Some gaseous combustion products, such as H{sub 2}O, CO, CO{sub 2}, NO and HCl, were distinguished by the characteristic emission band of each molecule. The equation for flame temperature calculation based on the diatomic molecule emission fine structure theory was improved through judicious utilization of the spectral running number 'm' which makes the temperature measurement simpler and faster. Some combustion information of the solid propellant had been given including the characteristic spectral profile, the distribution of the absolute spectral energy, the distribution of the combustion flame temperature, and the concentration distributions of HCl and NO versus burning time. The results will provide theoretical and experimental bases for improving the formula and raising combustion efficiency of solid propellant, and developing the design of rocket motor, infrared guidance and antiguidance systems. (Abstract Copyright [2006], Wiley Periodicals, Inc.)

Experimental studies on combustion of composite biomass pellets in fluidized bed.

Science.gov (United States)

Guo, Feihong; Zhong, Zhaoping

2017-12-01

This work presents studies on the combustion of Composite Biomass Pellets (CBP S ) in fluidized bed using bauxite particles as the bed material. Prior to the combustion experiment, cold-flow characterization and thermogravimetric analysis are performed to investigate the effect of air velocity and combustion mechanism of CBP S . The cold-state test shows that CBPs and bauxite particles fluidize well in the fluidized bed. However, because of the presence of large CBPs, optimization of the fluidization velocity is rather challenging. CBPs can gather at the bottom of the fluidized bed at lower gas velocities. On the contrary, when the velocity is too high, they accumulate in the upper section of the fluidized bed. The suitable fluidization velocity for the system in this study was found to be between 1.5-2.0m/s. At the same time, it is found that the critical fluidization velocity and the pressure fluctuation of the two-component system increase with the increase of CBPs mass concentration. The thermogravimetric experiment verifies that the combustion of CBPs is a first-order reaction, and it is divided into three stages: (i) dehydration, (ii) release and combustion of the volatile and (iii) the coke combustion. The combustion of CBPs is mainly based on the stage of volatile combustion, and its activation energy is greater than that of char combustion. During the combustion test, CBP S are burned at a 10kg/h feed rate, while the excess air is varied from 25% to 100%. Temperatures of the bed and flue gas concentrations (O 2 , CO, SO 2 and NO) are recorded. CBPs can be burnt stably, and the temperature of dense phase is maintained at 765-780°C. With the increase of the air velocity, the main combustion region has a tendency to move up. While the combustion is stable, O 2 and CO 2 concentrations are maintained at about 7%, and 12%, respectively. The concentration of SO 2 in the flue gas after the initial stage of combustion is nearly zero. Furthermore, NO concentration

Catalytic combustion in small wood burning appliances

Energy Technology Data Exchange (ETDEWEB)

Oravainen, H [VTT Energy, Jyvaeskylae (Finland)

1997-12-31

There is over a million hand fired small heating appliances in Finland where about 5,4 million cubic meters of wood fuel is used. Combustion in such heating appliances is a batch-type process. In early stages of combustion when volatiles are burned, the formation of carbon monoxide (CO) and other combustible gases are difficult to avoid when using fuels that have high volatile matter content. Harmful emissions are formed mostly after each fuel adding but also during char burnout period. When the CO-content in flue gases is, say over 0.5 %, also other harmful emissions will be formed. Methane (CH{sub 4}) and other hydrocarbons are released and the amount of polycyclic aromatic hydrocarbons (PAH)-compounds can be remarkable. Some PAH-compounds are very carcinogenic. It has been estimated that in Finland even more than 90 % of hydrocarbon and PAH emissions are due to small scale wood combustion. Emissions from transportation is excluded from these figures. That is why wood combustion has a net effect on greenhouse gas phenomena. For example carbon monoxide emissions from small scale wood combustion are two fold compared to that of energy production in power plants. Methane emission is of the same order as emission from transportation and seven fold compared with those of energy production. Emissions from small heating appliances can be reduced by developing the combustion techniques, but also by using other means, for example catalytic converters. In certain stages of the batch combustion, temperature is not high enough, gas mixing is not good enough and residence time is too short for complete combustion. When placed to a suitable place inside a heating appliance, a catalytic converter can oxidize unburned gases in the flue gas into compounds that are not harmful to the environment. (3 refs.)

Catalytic combustion in small wood burning appliances

Energy Technology Data Exchange (ETDEWEB)

Oravainen, H. [VTT Energy, Jyvaeskylae (Finland)

1996-12-31

There is over a million hand fired small heating appliances in Finland where about 5,4 million cubic meters of wood fuel is used. Combustion in such heating appliances is a batch-type process. In early stages of combustion when volatiles are burned, the formation of carbon monoxide (CO) and other combustible gases are difficult to avoid when using fuels that have high volatile matter content. Harmful emissions are formed mostly after each fuel adding but also during char burnout period. When the CO-content in flue gases is, say over 0.5 %, also other harmful emissions will be formed. Methane (CH{sub 4}) and other hydrocarbons are released and the amount of polycyclic aromatic hydrocarbons (PAH)-compounds can be remarkable. Some PAH-compounds are very carcinogenic. It has been estimated that in Finland even more than 90 % of hydrocarbon and PAH emissions are due to small scale wood combustion. Emissions from transportation is excluded from these figures. That is why wood combustion has a net effect on greenhouse gas phenomena. For example carbon monoxide emissions from small scale wood combustion are two fold compared to that of energy production in power plants. Methane emission is of the same order as emission from transportation and seven fold compared with those of energy production. Emissions from small heating appliances can be reduced by developing the combustion techniques, but also by using other means, for example catalytic converters. In certain stages of the batch combustion, temperature is not high enough, gas mixing is not good enough and residence time is too short for complete combustion. When placed to a suitable place inside a heating appliance, a catalytic converter can oxidize unburned gases in the flue gas into compounds that are not harmful to the environment. (3 refs.)

Rocket Flight.

Science.gov (United States)

Van Evera, Bill; Sterling, Donna R.

2002-01-01

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

Rocket observations

Science.gov (United States)

1984-05-01

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

Co-combustion of anthracite coal and wood pellets: Thermodynamic analysis, combustion efficiency, pollutant emissions and ash slagging.

Science.gov (United States)

Guo, Feihong; Zhong, Zhaoping

2018-08-01

This work presents studies on the co-combustion of anthracite coal and wood pellets in fluidized bed. Prior to the fluidized bed combustion, thermogravimetric analysis are performed to investigate the thermodynamic behavior of coal and wood pellets. The results show that the thermal decomposition of blends is divided into four stages. The co-firing of coal and wood pellets can promote the combustion reaction and reduce the emission of gaseous pollutants, such as SO 2 and NO. It is important to choose the proportion of wood pellets during co-combustion due to the low combustion efficiency caused by large pellets with poor fluidization. Wood pellets can inhibit the volatilization of trace elements, especially for Cr, Ni and V. In addition, the slagging ratio of wood pellets ash is reduced by co-firing with coal. The research on combustion of coal and wood pellets is of great significance in engineering. Copyright © 2018 Elsevier Ltd. All rights reserved.

Ramjet Application Possibilities for Increasing Fire Range of the Multiple Launch Rocket Systems Ammunition

Directory of Open Access Journals (Sweden)

V. N. Zubov

2015-01-01

Full Text Available The article considers a possibility to increase a flying range of the perspective rockets equipped with the control unit with aerodynamic controllers for the multiple launch rocket systems “Smerch”.To increase a flying range and reduce a starting mass of the rocket, the paper studies a possibility to replace the single-mode rocket engine used in the solid-fuel rocket motor for the direct-flow propulsion jet engine (DFPJE with not head sector air intakes. The DFPJE is implemented according to the classical scheme with a fuel charged in the combustion chamber. A separated solid propellant starting accelerator provides the rocket acceleration to reach a speed necessary for the DFPJE to run.When designing the DFPJE a proper choice of not head air intake parameters is one of the most difficult points. For this purpose a COSMOS Flow Simulation software package and analytical dependences were used to define the following: a boundary layer thickness where an air intake is set, maximum permissible and appropriate angles of attack and deviation angles of controllers at the section where the DFPJE works, and some other parameters as well.Calculation of DFPJE characteristics consisted in determining parameters of an air-gas path of the propulsion system, geometrical sizes of the pipeline flow area, sizes of a fuel charge, and dependence of the propulsion system impulse on the flight height and speed. Calculations were performed both in thermodynamic statement of problem and in using software package of COSMOS Flow Simulation.As a result of calculations and design engineering activities the air intake profile is created and mass-dimensional characteristics of DFPJE are defined. Besides, calculations of the starting solid fuel accelerator were carried out. Further design allowed us to create the rocket shape, estimate its mass-dimensional characteristics, and perform ballistic calculations, which proved that achieving a range of 120 km for the rocket is

THE ADIABATIC DEMAGNETIZATION REFRIGERATOR FOR THE MICRO-X SOUNDING ROCKET TELESCOPE

International Nuclear Information System (INIS)

Wikus, P.; Bagdasarova, Y.; Figueroa-Feliciano, E.; Leman, S. W.; Rutherford, J. M.; Trowbridge, S. N.; Adams, J. S.; Bandler, S. R.; Eckart, M. E.; Kelley, R. L.; Kilbourne, C. A.; Porter, F. S.; Doriese, W. B.; McCammon, D.

2010-01-01

The Micro-X Imaging X-ray Spectrometer is a sounding rocket payload slated for launch in 2011. An array of Transition Edge Sensors, which is operated at a bath temperature of 50 mK, will be used to obtain a high resolution spectrum of the Puppis-A supernova remnant. An Adiabatic Demagnetization Refrigerator (ADR) with a 75 gram Ferric Ammonium Alum (FAA) salt pill in the bore of a 4 T superconducting magnet provides a stable heat sink for the detector array only a few seconds after burnout of the rocket motors. This requires a cold stage design with very short thermal time constants. A suspension made from Kevlar strings holds the 255 gram cold stage in place. It is capable of withstanding loads in excess of 200 g. Stable operation of the TES array in proximity to the ADR magnet is ensured by a three-stage magnetic shielding system which consists of a superconducting can, a high-permeability shield and a bucking coil. The development and testing of the Micro-X payload is well underway.

Rockets two classic papers

CERN Document Server

Goddard, Robert

2002-01-01

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

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

National Research Council Canada - National Science Library

Shelley, J

2001-01-01

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

Rocket Engine Innovations Advance Clean Energy

Science.gov (United States)

2012-01-01

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

Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

Science.gov (United States)

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

2001-01-01

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

Multisized Inert Particle Loading for Solid Rocket Axial Combustion Instability Suppression

Directory of Open Access Journals (Sweden)

David R. Greatrix

2012-01-01

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

Two-Dimensional Motions of Rockets

Science.gov (United States)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

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

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

Science.gov (United States)

Sutton, George P.

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

CSP-based chemical kinetics mechanisms simplification strategy for non-premixed combustion: An application to hybrid rocket propulsion

KAUST Repository

Ciottoli, Pietro P.; Malpica Galassi, Riccardo; Lapenna, Pasquale E.; Leccese, G.; Bianchi, D.; Nasuti, F.; Creta, F.; Valorani, M.

2017-01-01

A set of simplified chemical kinetics mechanisms for hybrid rocket applications using gaseous oxygen (GOX) and hydroxyl-terminated polybutadiene (HTPB) is proposed. The starting point is a 561-species, 2538-reactions, detailed chemical kinetics

Laboratory weathering of combusted oil shale

International Nuclear Information System (INIS)

Essington, M.E.

1991-01-01

The objective of this study was to examine the mineralogy and leachate chemistry of three combusted oil shales (two Green River Formation and one New Albany) in a laboratory weathering environment using the humidity cell technique. The mineralogy of the combusted western oil shales (Green River Formation) is process dependent. In general, processing resulted in the formation of anhydrite, lime, periclase, and hematite. During the initial stages of weathering, lime, periclase, and hematite. During the initial stages of weathering, lime, periclase, and anhydrite dissolve and ettringite precipitates. The initial leachates are highly alkaline, saline, and dominated by Na, hydroxide, and SO 4 . As weathering continues, ettringite precipitates. The initial leachates are highly alkaline, saline, and dominated by Na, hydroxide, and SO 4 . As weathering continues, ettringite dissolves, gypsum and calcite precipitate, and the leachates are dominated by Mg, SO 4 , and CO 3 . Leachate pH is rapidly reduced to between 8.5 and 9 with leaching. The combusted eastern oil shale (New Albany) is composed of quartz, illite, hematite, and orthoclase. Weathering results in the precipitation of gypsum. The combusted eastern oil shale did not display a potential to produce acid drainage. Leachate chemistry was dominated by Ca and SO 4 . Element concentrations continually decreased with weathering. IN a western disposal environment receiving minimal atmospheric precipitation, spent oil shale will remain in the initial stages of weathering, and highly alkaline and saline conditions will dominate leachate chemistry. In an eastern disposal environment, soluble salts will be rapidly removed from the spent oil shale to potentially affect the surrounding environment

Final report: Prototyping a combustion corridor; FINAL

International Nuclear Information System (INIS)

Rutland, Christopher J.; Leach, Joshua

2001-01-01

The Combustion Corridor is a concept in which researchers in combustion and thermal sciences have unimpeded access to large volumes of remote computational results. This will enable remote, collaborative analysis and visualization of state-of-the-art combustion science results. The Engine Research Center (ERC) at the University of Wisconsin - Madison partnered with Lawrence Berkeley National Laboratory, Argonne National Laboratory, Sandia National Laboratory, and several other universities to build and test the first stages of a combustion corridor. The ERC served two important functions in this partnership. First, we work extensively with combustion simulations so we were able to provide real world research data sets for testing the Corridor concepts. Second, the ERC was part of an extension of the high bandwidth based DOE National Laboratory connections to universities

A method for determining the completeness of fuel combustion

Energy Technology Data Exchange (ETDEWEB)

Tavger, M.D.; Chepkin, V.M.; Gruzdev, V.N.; Talantov, A.V.

1982-01-01

The current of conductivity (ionization) of gaseous combustion products, which forms with feeding of electric voltage to a special probe, is proposed for determining the completeness of fuel combustion. Here, the charged particles are formed from substances which form in the intermediate stages of the combustion reaction. The volume of charged particles is proportional to the volume of the intermediate substances, whose presence attests to the incompleteness of the combustion reaction. The fullness of fuel combustion is determined from a formula which includes the stoichiometric coefficient, a gas constant, the energy of activation, the characteristics of the chemical activity of the intermediate substances, the coefficient of air excess, the temperature of the combustion products and the conductivity current.

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

Energy Technology Data Exchange (ETDEWEB)

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

2016-11-15

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

Modeling and simulation of combustion chamber and propellant dynamics and issues in active control of combustion instabilities

Science.gov (United States)

Isella, Giorgio Carlo

A method for a comprehensive approach to analysis of the dynamics of an actively controlled combustion chamber, with detailed analysis of the combustion models for the case of a solid rocket propellant, is presented here. The objective is to model the system as interconnected blocks describing the dynamics of the chamber, combustion and control. The analytical framework for the analysis of the dynamics of a combustion chamber is based on spatial averaging, as introduced by Culick. Combustion dynamics are analyzed for the case of a solid propellant. Quasi-steady theory is extended to include the dynamics of the gas-phase and also of a surface layer. The models are constructed so that they produce a combustion response function for the solid propellant that can be immediately introduced in the our analytical framework. The principal objective mechanisms responsible for the large sensitivity, observed experimentally, of propellant response to small variations. We show that velocity coupling, and not pressure coupling, has the potential to be the mechanism responsible for that high sensitivity. We also discuss the effect of particulate modeling on the global dynamics of the chamber and revisit the interpretation of the intrinsic stability limit for burning of solid propellants. Active control is also considered. Particular attention is devoted to the effect of time delay (between sensing and actuation); several methods to compensate for it are discussed, with numerical examples based on the approximate analysis produced by our framework. Experimental results are presented for the case of a Dump Combustor. The combustor exhibits an unstable burning mode, defined through the measurement of the pressure trace and shadowgraph imaging. The transition between stable and unstable modes of operation is characterized by the presence of hysteresis, also observed in other experimental works, and hence not a special characteristic of this combustor. Control is introduced in the

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

Science.gov (United States)

Thorpe, Douglas G.

1991-01-01

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

Description and Flight Performance Results of the WASP Sounding Rocket

Science.gov (United States)

De Pauw, J. F.; Steffens, L. E.; Yuska, J. A.

1968-01-01

A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

Combustion Chamber Fluid Dynamics and Hypergolic Gel Propellant Chemistry Simulations for Selectable Thrust Rocket Engines

National Research Council Canada - National Science Library

Nusca, Michael J; Chen, Chiung-Chu; McQuaid, Michael J

2007-01-01

.... Computational fluid dynamics is employed to model the chemically reacting flow within a system's combustion chamber, and computational chemistry is employed to characterize propellant physical and reactive properties...

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

Science.gov (United States)

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

2002-01-01

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

Eddie Rocket's Franchise

OpenAIRE

Vahter, Jenni

2008-01-01

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

Effects of catalysts on combustion characteristics and kinetics of coal-char blends

Science.gov (United States)

Hu, Yingjie; Wang, Zhiqiang; Cheng, Xingxing; Liu, Ming; Ma, Chunyuan

2018-04-01

The effects of Fe2O3, CaO, and MnO2 on the combustion characteristics and kinetics of coal-char blends were investigated using thermogravimetric analysis (TGA) and differential scanning calorimetry (DSC). The results indicated that catalysts exhibited positive effects on the combustion characteristics of coal-char blends, especially in the initial period of coal-char blends combustion. With catalysts addition (mass 1.5%), it could improves volatile matter release, and reduces ignition point, promotes char to begin burning under lower temperature. The ignition index (C) was increased, respectively, by 27% for Fe2O3, 6% for CaO, 11.3% for MnO2, and the combustion characteristic index ( S ) was increased respectively, by 29% for Fe2O3, 5% for CaO, 8.3% for MnO2. In addition, two kinetic models (R2 and F1) were adopted to calculate the kinetic parameters in different stage of combustion processes. The results showed that with Fe2O3 or CaO addition, the activation energy at second stage decreases from 86.0 KJ/mol to 76.92 KJ/mol and 75.12 KJ/mol, respectively. There are no obvious decreases at the third stage of samples combustion process.

Concentric traveling ionospheric disturbances triggered by the launch of a SpaceX Falcon 9 rocket

Science.gov (United States)

Lin, Charles C. H.; Shen, Ming-Hsueh; Chou, Min-Yang; Chen, Chia-Hung; Yue, Jia; Chen, Po-Cheng; Matsumura, Mitsuru

2017-08-01

We report the first observation of concentric traveling ionospheric disturbances (CTIDs) triggered by the launch of a SpaceX Falcon 9 rocket on 17 January 2016. The rocket-triggered ionospheric disturbances show shock acoustic wave signature in the time rate change (time derivative) of total electron content (TEC), followed by CTIDs in the 8-15 min band-pass filtering of TEC. The CTIDs propagated northward with phase velocity of 241-617 m/s and reached distances more than 1000 km away from the source on the rocket trajectory. The wave characteristics of CTIDs with periods of 10.5-12.7 min and wavelength 200-400 km agree well with the gravity wave dispersion relation. The optimal wave source searching and gravity wave ray tracing technique suggested that the CTIDs have multiple sources which are originated from 38-120 km altitude before and after the ignition of the second-stage rocket, 200 s after the rocket was launched.

Multi-Point Combustion System: Final Report

Science.gov (United States)

Goeke, Jerry; Pack, Spencer; Zink, Gregory; Ryon, Jason

2014-01-01

A low-NOx emission combustor concept has been developed for NASA's Environmentally Responsible Aircraft (ERA) program to meet N+2 emissions goals for a 70,000 lb thrust engine application. These goals include 75 percent reduction of LTO NOx from CAEP6 standards without increasing CO, UHC, or smoke from that of current state of the art. An additional key factor in this work is to improve lean combustion stability over that of previous work performed on similar technology in the early 2000s. The purpose of this paper is to present the final report for the NASA contract. This work included the design, analysis, and test of a multi-point combustion system. All design work was based on the results of Computational Fluid Dynamics modeling with the end results tested on a medium pressure combustion rig at the UC and a medium pressure combustion rig at GRC. The theories behind the designs, results of analysis, and experimental test data will be discussed in this report. The combustion system consists of five radially staged rows of injectors, where ten small scale injectors are used in place of a single traditional nozzle. Major accomplishments of the current work include the design of a Multipoint Lean Direct Injection (MLDI) array and associated air blast and pilot fuel injectors, which is expected to meet or exceed the goal of a 75 percent reduction in LTO NOx from CAEP6 standards. This design incorporates a reduced number of injectors over previous multipoint designs, simplified and lightweight components, and a very compact combustor section. Additional outcomes of the program are validation that the design of these combustion systems can be aided by the use of Computational Fluid Dynamics to predict and reduce emissions. Furthermore, the staging of fuel through the individually controlled radially staged injector rows successfully demonstrated improved low power operability as well as improvements in emissions over previous multipoint designs. Additional comparison

Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

Science.gov (United States)

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

2004-01-01

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

Two-Rockets Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

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

Combustion characteristics of crude jatropha oil droplets using rhodium liquid as a homogeneous combustion catalyst

Science.gov (United States)

Nanlohy, Hendry Y.; Wardana, I. N. G.; Hamidi, N.; Yuliati, L.

2018-01-01

Combustion characteristics of crude jatropha oil droplet at room temperature with and without catalyst have been studied experimentally. Its combustion characteristics have been observed by igniting the oil droplet on a junction of a thermocouple, and the combustion characteristics of oil droplets are observed using a high-speed camera. The results show that the uniqueness of crude jatropha oil as alternative fuel is evidenced by the different stages of combustion caused by thermal cracking in burning droplets. The results also show that the role of the catalyst is not only an accelerator agent, but there are other unique functions and roles as a stabilizer. Moreover, the results also found that the catalyst was able to shorten the ignition timing and burnout time. This phenomenon proves that the presence of catalysts alters and weakens the structure of the triglyceride geometry so that the viscosity and flash point is reduced, the fuel absorbs heat well and flammable.

Bibliography of Books and Published Reports on Gas Turbines, Jet Propulsion, and Rocket Power Plants

Science.gov (United States)

1951-06-01

Ink , New York, 1945. W. Ley, Rockets. Viking Press. New York. 1945. LI. S. Zim, Rockets and jets. Harcourt Brace, New York, 1945. Jet propulsion...Hausenblas, Design nomograms for turbine stages. Motortechnische Zeit. 11, 96 (Aug. 1950). S. L. Koutz et al., Effect of beat and power extraction on...Edelman, The pulsating engine-its evolution and future prospects. SAE Quart. Trans. 1, 204 (1947). R. McLarren, Project Squid probes pulsejet. Aviation

Rocketing into the future the history and technology of rocket planes

CERN Document Server

van Pelt, Michel

2012-01-01

Rocketing into the Future journeys into the exciting world of rocket planes, examining the exotic concepts and actual flying vehicles that have been devised over the last one hundred years. Lavishly illustrated with over 150 photographs, it recounts the history of rocket planes from the early pioneers who attached simple rockets on to their wooden glider airplanes to the modern world of high-tech research vehicles. The book then looks at the possibilities for the future. The technological and economic challenges of the Space Shuttle proved insurmountable, and thus the program was unable to fulfill its promise of low-cost access to space. However, the burgeoning market of suborbital space tourism may yet give the necessary boost to the development of a truly reusable spaceplane.

Combustion of Sewage Sludge as Alternative Fuel for Cement Industry

Institute of Scientific and Technical Information of China (English)

LI Fuzhou; ZHANG Wei

2011-01-01

The combustion of sewage sludge and coal was studied by thermogravimetric analysis.Both differential scanning calorimetric analysis and derivative thermogravimetric profiles showed differences between combustion of sewage sludge and coal, and non-isothermal kinetics analysis method was applied to evaluate the combustion process. Based on Coats-Redfem integral method, some reaction models were tested,the mechanism and kinetics of the combustion reaction were discussed. The results show that the combustion of sewage sludge is mainly in the Iow temperature stage, meanwhile the ignition temperature and Arrhenius activation energy are lower than that of coal. The combustion of sewage sludge has the advantage over coal in some aspects, thus sewage sludge can partly replace coal used as cement industry fuel.

High-Temperature Polymer Composites Tested for Hypersonic Rocket Combustor Backup Structure

Science.gov (United States)

Sutter, James K.; Shin, E. Eugene; Thesken, John C.; Fink, Jeffrey E.

2005-01-01

Significant component weight reductions are required to achieve the aggressive thrust-toweight goals for the Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-toorbit vehicle. A collaboration between the NASA Glenn Research Center and Boeing Rocketdyne was formed under the Higher Operating Temperature Propulsion Components (HOTPC) program and, currently, the Ultra-Efficient Engine Technology (UEET) Project to develop carbon-fiber-reinforced high-temperature polymer matrix composites (HTPMCs). This program focused primarily on the combustor backup structure to replace all metallic support components with a much lighter polymer-matrixcomposite- (PMC-) titanium honeycomb sandwich structure.

Metallic hydrogen: The most powerful rocket fuel yet to exist

Energy Technology Data Exchange (ETDEWEB)

Silvera, Isaac F [Lyman Laboratory of Physics, Harvard University, Cambridge MA 02138 (United States); Cole, John W, E-mail: silvera@physics.harvard.ed [NASA MSFC, Huntsville, AL 35801 (United States)

2010-03-01

Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, I{sub sp}. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of {approx}460s; metallic hydrogen has a theoretical I{sub sp} of 1700s. Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

Two-dimensional motions of rockets

International Nuclear Information System (INIS)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the descending parts of the trajectories tend to be gentler and straighter slopes than the ascending parts for relatively large launching angles due to the non-vanishing thrusts. We discuss the ranges, the maximum altitudes and the engine performances of the rockets. It seems that the exponential fuel exhaustion can be the most potent engine for the longest and highest flights

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

Science.gov (United States)

Roche, Joseph M.

2002-01-01

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

Advanced CFD modelling of air and recycled flue gas staging in a waste wood-fired grate boiler for higher combustion efficiency and greater environmental benefits.

Science.gov (United States)

Rajh, Boštjan; Yin, Chungen; Samec, Niko; Hriberšek, Matjaž; Kokalj, Filip; Zadravec, Matej

2018-07-15

Grate-fired boilers are commonly used to burn biomass/wastes for heat and power production. In spite of the recent breakthrough in integration of advanced secondary air systems in grate boilers, grate-firing technology needs to be advanced for higher efficiency and lower emissions. In this paper, innovative staging of combustion air and recycled flue gas in a 13 MW th waste wood-fired grate boiler is comprehensively studied based on a numerical model that has been previously validated. In particular, the effects of the jet momentum, position and orientation of the combustion air and recycled flue gas streams on in-furnace mixing, combustion and pollutant emissions from the boiler are examined. It is found that the optimized air and recycled flue gas jets remarkably enhance mixing and heat transfer, result in a more uniform temperature and velocity distribution, extend the residence time of the combustibles in the hot zone and improve burnout in the boiler. Optimizing the air and recycled flue gas jet configuration can reduce carbon monoxide emission from the boiler by up to 86%, from the current 41.0 ppm to 5.7 ppm. The findings of this study can serve as useful guidelines for novel design and optimization of the combustion air supply and flue gas recycling for grate boilers of this type. Copyright © 2018 Elsevier Ltd. All rights reserved.

Heat Transfer by Thermo-Capillary Convection. Sounding Rocket COMPERE Experiment SOURCE

Science.gov (United States)

Fuhrmann, Eckart; Dreyer, Michael

2009-08-01

This paper describes the results of a sounding rocket experiment which was partly dedicated to study the heat transfer from a hot wall to a cold liquid with a free surface. Natural or buoyancy-driven convection does not occur in the compensated gravity environment of a ballistic phase. Thermo-capillary convection driven by a temperature gradient along the free surface always occurs if a non-condensable gas is present. This convection increases the heat transfer compared to a pure conductive case. Heat transfer correlations are needed to predict temperature distributions in the tanks of cryogenic upper stages. Future upper stages of the European Ariane V rocket have mission scenarios with multiple ballistic phases. The aims of this paper and of the COMPERE group (French-German research group on propellant behavior in rocket tanks) in general are to provide basic knowledge, correlations and computer models to predict the thermo-fluid behavior of cryogenic propellants for future mission scenarios. Temperature and surface location data from the flight have been compared with numerical calculations to get the heat flux from the wall to the liquid. Since the heat flux measurements along the walls of the transparent test cell were not possible, the analysis of the heat transfer coefficient relies therefore on the numerical modeling which was validated with the flight data. The coincidence between experiment and simulation is fairly good and allows presenting the data in form of a Nusselt number which depends on a characteristic Reynolds number and the Prandtl number. The results are useful for further benchmarking of Computational Fluid Dynamics (CFD) codes such as FLOW-3D and FLUENT, and for the design of future upper stage propellant tanks.

Regulating low-NOx and high-burnout deep-air-staging combustion under real-furnace conditions in a 600 MWe down-fired supercritical boiler by strengthening the staged-air effect.

Science.gov (United States)

Kuang, Min; Wang, Zhihua; Zhu, Yanqun; Ling, Zhongqian; Li, Zhengqi

2014-10-21

A 600 MW(e) down-fired pulverized-coal supercritical boiler, which was equipped with a deep-air-staging combustion system for reducing the particularly high NOx emissions, suffered from the well-accepted contradiction between low NOx emissions and high carbon in fly ash, in addition to excessively high gas temperatures in the hopper that jeopardized the boiler's safe operations. Previous results uncovered that under low-NOx conditions, strengthening the staged-air effect by decreasing the staged-air angle and simultaneously increasing the staged-air damper opening alleviated the aforementioned problems to some extent. To establish low-NOx and high-burnout circumstances and control the aforementioned hopper temperatures, a further staged-air retrofit with horizontally redirecting staged air through an enlarged staged-air slot area was performed to greatly strengthen the staged-air effect. Full-load industrial-size measurements were performed to confirm the availability of this retrofit. The present data were compared with those published results before the retrofit. High NOx emissions, low carbon in fly ah, and high hopper temperatures (i.e., levels of 1036 mg/m(3) at 6% O2, 3.72%, and about 1300 °C, respectively) appeared under the original conditions with the staged-air angle of 45° and without overfire air (OFA) application. Applying OFA and reducing the angle to 20° achieved an apparent NOx reduction and a moderate hopper temperature decrease while a sharp increase in carbon in fly ash (i.e., levels of 878 mg/m(3) at 6% O2, about 1200 °C, and 9.81%, respectively). Fortunately, the present staged-air retrofit was confirmed to be applicable in regulating low-NOx, high-burnout, and low hopper temperature circumstances (i.e., levels of 867 mg/m(3) at 6% O2, 5.40%, and about 1100 °C, respectively).

An experimental investigation of the combustion process of a heavy-duty diesel engine enriched with H{sub 2}

Energy Technology Data Exchange (ETDEWEB)

Liew, C.; Li, H.; Nuszkowski, J.; Liu, S.; Gatts, T.; Atkinson, R.; Clark, N. [Department of Mechanical and Aerospace Engineering, West Virginia University, Morgantown, WV 26506-6106 (United States)

2010-10-15

This paper investigated the effect of hydrogen (H{sub 2}) addition on the combustion process of a heavy-duty diesel engine. The addition of a small amount of H{sub 2} was shown to have a mild effect on the cylinder pressure and combustion process. When operated at high load, the addition of a relatively large amount of H{sub 2} substantially increased the peak cylinder pressure and the peak heat release rate. Compared to the two-stage combustion process of diesel engines, a featured three-stage combustion process of the H{sub 2}-diesel dual fuel engine was observed. The extremely high peak heat release rate represented a combination of diesel diffusion combustion and the premixed combustion of H{sub 2} consumed by multiple turbulent flames, which substantially enhanced the combustion process of H{sub 2}-diesel dual fuel engine. However, the addition of a relatively large amount of H{sub 2} at low load did not change the two-stage heat release process pattern. The premixed combustion was dramatically inhibited while the diffusion combustion was slightly enhanced and elongated. The substantially reduced peak cylinder pressure at low load was due to the deteriorated premixed combustion. (author)

Another Look at Rocket Thrust

Science.gov (United States)

Hester, Brooke; Burris, Jennifer

2012-01-01

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

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

Science.gov (United States)

Matogawa, Yasunori

2007-12-01

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

South Pole rockets, (1)

International Nuclear Information System (INIS)

Kimura, Iwane

1977-01-01

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

Combustion in microgravity: The French contribution

Science.gov (United States)

Prud'homme, Roger; Legros, Guillaume; Torero, José L.

2017-01-01

Microgravity (drop towers, parabolic flights, sounding rockets and space stations) are particularly relevant to combustion problems given that they show high-density gradients and in many cases weak forced convection. For some configurations where buoyancy forces result in complex flow fields, microgravity leads to ideal conditions that correspond closely to canonical problems, e.g., combustion of a spherical droplet in a far-field still atmosphere, Emmons' problem for flame spreading over a solid flat plate, deflagration waves, etc. A comprehensive chronological review on the many combustion studies in microgravity was written first by Law and Faeth (1994) and then by F.A. Williams (1995). Later on, new recommendations for research directions have been delivered. In France, research has been managed and supported by CNES and CNRS since the creation of the microgravity research group in 1992. At this time, microgravity research and future activities contemplated the following: Droplets: the "D2 law" has been well verified and high-pressure behavior of droplet combustion has been assessed. The studies must be extended in two main directions: vaporization in mixtures near the critical line and collective effects in dense sprays. Flame spread: experiments observed blue flames governed by diffusion that are in accordance with Emmons' theory. Convection-dominated flames showed significant departures from the theory. Some theoretical assumptions appeared controversial and it was noted that radiation effects must be considered, especially when regarding the role of soot production in quenching. Heterogeneous flames: two studies are in progress, one in Poitiers and the other in Marseilles, about flame/suspension interactions. Premixed and triple flames: the knowledge still needs to be complemented. Triple flames must continue to be studied and understanding of "flame balls" still needs to be addressed.

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

Science.gov (United States)

Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

2016-09-01

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

Micro-Rockets for the Classroom.

Science.gov (United States)

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

1999-01-01

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

Heavy fuel oil pyrolysis and combustion: kinetics and evolved gases investigated by TGA-FTIR

KAUST Repository

Abdul Jameel, Abdul Gani

2017-08-24

Heavy fuel oil (HFO) obtained from crude oil distillation is a widely used fuel in marine engines and power generation technologies. In the present study, the pyrolysis and combustion of a Saudi Arabian HFO in nitrogen and in air, respectively, were investigated using non-isothermal thermo-gravimetric analysis (TGA) coupled with a Fourier-transform infrared (FTIR) spectrometer. TG and DTG (differential thermo-gravimetry) were used for the kinetic analysis and to study the mass loss characteristics due to the thermal degradation of HFO at temperatures up to 1000°C and at various heating rates of 5, 10 and 20°C/min, in air and N2 atmospheres. FTIR analysis was then performed to study the composition of the evolved gases. The TG/DTG curves during HFO combustion show the presence of three distinct stages: the low temperature oxidation (LTO); fuel decomposition (FD); and high temperature oxidation (HTO) stages. The TG/DTG curves obtained during HFO pyrolysis show the presence of two devolatilization stages similar to that seen in the LTO stage of HFO combustion. Apart from this, the TG/DTG curves obtained during HFO combustion and pyrolysis differ significantly. Kinetic analysis was also performed using the distributed activation energy model, and the kinetic parameter (E) was determined for the different stages of HFO combustion and pyrolysis processes, yielding a good agreement with the measured TG profiles. FTIR analysis showed the signal of CO2 as approximately 50 times more compared to the other pollutant gases under combustion conditions. Under pyrolytic conditions, the signal intensity of alkane functional groups was the highest followed by alkenes. The TGA-FTIR results provide new insights into the overall HFO combustion processes, which can be used to improve combustor designs and control emissions.

Analysis of exergy loss of gasoline surrogate combustion process based on detailed chemical kinetics

International Nuclear Information System (INIS)

Sun, Hongjie; Yan, Feng; Yu, Hao; Su, W.H.

2015-01-01

Highlights: • We explored the exergy loss sources of gasoline engine like combustion process. • The model combined non-equilibrium thermodynamics with detailed chemical kinetics. • We explored effects of initial conditions on exergy loss of combustion process. • Exergy loss decreases 15% of fuel chemical exergy by design of initial conditions. • Correspondingly, the second law efficiency increases from 38.9% to 68.9%. - Abstract: Chemical reaction is the most important source of combustion irreversibility in premixed conditions, but details of the exergy loss mechanisms have not been explored yet. In this study numerical analysis based on non-equilibrium thermodynamics combined with detailed chemical kinetics is conducted to explore the exergy loss mechanism of gasoline engine like combustion process which is simplified as constant volume combustion. The fuel is represented by the common accepted gasoline surrogates which consist of four components: iso-octane (57%), n-heptane (16%), toluene (23%), and 2-pentene (4%). We find that overall exergy loss is mainly composed of three peaks along combustion generated from chemical reactions in three stages, the conversion from large fuel molecules into small molecules (as Stage 1), the H 2 O 2 loop-related reactions (as Stage 2), and the violent oxidation reactions of CO, H, and O (as Stage 3). The effects of individual combustion boundaries, including temperature, pressure, equivalence ratio, oxygen concentration, on combustion exergy loss have been widely investigated. The combined effects of combustion boundaries on the total loss of gasoline surrogates are also investigated. We find that in a gasoline engine with a compression ratio of 10, the total loss can be reduced from 31.3% to 24.3% using lean combustion. The total loss can be further reduced to 22.4% by introducing exhaust gas recirculation and boosting the inlet charge. If the compression ratio is increased to 17, the total loss can be decreased to

NOx reduction and NO2 emission characteristics in rich-lean combustion of hydrogen

OpenAIRE

Shudo, Toshio; Omori, Kento; Hiyama, Osamu

2008-01-01

Hydrogen is a clean alternative to conventional hydrocarbon fuels, but it is very important to reduce the nitrogen oxides (NOx) emissions generated by hydrogen combustion. The rich-lean combustion or staged combustion is known to reduce NOx emissions from continuous combustion burners such as gas turbines and boilers, and NOx reduction effects have been demonstrated for hydrocarbon fuels. The authors applied rich-lean combustion to a hydrogen gas turbine and showed its NOx reduction effect in...

A spectroscopy study of gasoline partially premixed compression ignition spark assisted combustion

International Nuclear Information System (INIS)

Pastor, J.V.; García-Oliver, J.M.; García, A.; Micó, C.; Durrett, R.

2013-01-01

Highlights: ► PPC combustion combined with spark assistance and gasoline fuel on a CI engine. ► Chemiluminescence of different chemical species describes the progress of combustion reaction. ► Spectra of a novel combustion mode under SACI conditions is described. ► UV–Visible spectrometry, high speed imaging and pressure diagnostic were employed for analysis. - Abstract: Nowadays many research efforts are focused on the study and development of new combustion modes, mainly based on the use of locally lean air–fuel mixtures. This characteristic, combined with exhaust gas recirculation, provides low combustion temperatures that reduces pollutant formation and increases efficiency. However these combustion concepts have some drawbacks, related to combustion phasing control, which must be overcome. In this way, the use of a spark plug has shown to be a good solution to improve phasing control in combination with lean low temperature combustion. Its performance is well reported on bibliography, however phenomena involving the combustion process are not completely described. The aim of the present work is to develop a detailed description of the spark assisted compression ignition mode by means of application of UV–Visible spectrometry, in order to improve insight on the combustion process. Tests have been performed in an optical engine by means of broadband radiation imaging and emission spectrometry. The engine hardware is typical of a compression ignition passenger car application. Gasoline was used as the fuel due to its low reactivity. Combining broadband luminosity images with pressure-derived heat-release rate and UV–Visible spectra, it was possible to identify different stages of the combustion reaction. After the spark discharge, a first flame kernel appears and starts growing as a premixed flame front, characterized by a low and constant heat-release rate in combination with the presence of remarkable OH radical radiation. Heat release increases

Advanced air staging techniques to improve fuel flexibility, reliability and emissions in fluidized bed co-combustion

Energy Technology Data Exchange (ETDEWEB)

Aamand, Lars-Erik; Leckner, Bo [Chalmers Technical Univ., Goeteborg (Sweden); Luecke, Karsten; Werther, Joachim [Technical Univ. of Hamburg-Harburg (Germany)

2001-12-01

A joint research project between the Technical University of Hamburg-Harburg and Chalmers Technical University. For operation under co-combustion the following results should be considered: The high ash content of the sewage sludge results in significantly increased ash flows. Although high alkali metal concentrations are found in the sewage sludge ash, no critical concentrations were reached and tendencies to fouling were not observed. The trace metal input rises with increased sludge fraction. However, emissions of metal compounds were well below legal limits. The trace metals tend to accumulate on the fly ash. In general, very low fuel nitrogen conversions to NO and N{sub 2}O of 2 - 4 % are achievable. With coal as a base fuel alternative air staging with secondary air supply after solids separation attains even lower NO emissions than normal staging without strongly affecting CO and SO{sub 2} emissions. Alternative staging also reduces N{sub 2}O emissions. An optimum for the excess air ratio in the riser of 1.05 was found for a total excess air ratio of 1.2. The higher the volatile content of the fuel is, the less effective the NO reduction due to air staging becomes. The measurements suggest that the optimum gas residence time regarding the emissions in CFB combustors is around 6 to 7 s. These times are achieved in commercial scale plants due to their large cyclones that perhaps partly can replace a large afterburner chamber. The circulating fluidized bed boiler can be operated in a very flexible way with various fuel mixtures up to an energy fraction of sludge of 25% without exceeding legal emission limits.

Effects of air jet duration and timing on the combustion characteristics of high-pressure air jet controlled compression ignition combustion mode in a hybrid pneumatic engine

International Nuclear Information System (INIS)

Long, Wuqiang; Meng, Xiangyu; Tian, Jiangping; Tian, Hua; Cui, Jingchen; Feng, Liyan

2016-01-01

Highlights: • A 3-D CFD model of the power cylinder in HPE was developed. • High-pressure air JCCI combustion mode includes two-stage high-temperature reaction. • The combustion phasing of the pre-mixture is controllable via the SOJ timing. • There exists an optimum SOJ timing for obtaining the highest combustion efficiency and shortest burning duration. - Abstract: The high-pressure air jet controlled compression ignition (JCCI) combustion mode was employed to control the premixed diesel compression ignition combustion phasing by using the compound thermodynamic cycle under all operating conditions, which is accomplished in a hybrid pneumatic engine (HPE). A three-dimensional computational fluid dynamics (CFD) numerical simulation coupled with reduced n-heptane chemical kinetics mechanism has been applied to investigate the effects of high-pressure air jet duration and the start of jet (SOJ) timing on the combustion characteristics in the power cylinder of HPE. By sweeping the high-pressure air jet durations from 6 to 14 °CA and SOJ timings from −12 °CA ATDC to the top dead center (TDC) under the air jet temperatures of 400 and 500 K, respectively, the low- and high-temperature reactions, combustion efficiency, as well as the combustion phasing and burning duration have been analyzed in detail. The results illustrated that a longer air jet duration results in a higher peak in the first-stage high-temperature reaction, and the short air jet duration of 6 °CA can lead to a higher combustion efficiency. The SOJ timing sweep results showed that there exists an optimum timing for obtaining the highest combustion efficiency and shortest burning duration.

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

Science.gov (United States)

2010-01-01

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

Improving combustion characteristics and NO(x) emissions of a down-fired 350 MW(e) utility boiler with multiple injection and multiple staging.

Science.gov (United States)

Kuang, Min; Li, Zhengqi; Xu, Shantian; Zhu, Qunyi

2011-04-15

Within a Mitsui Babcock Energy Limited down-fired pulverized-coal 350 MW(e) utility boiler, in situ experiments were performed, with measurements taken of gas temperatures in the burner and near the right-wall regions, and of gas concentrations (O(2) and NO) from the near-wall region. Large combustion differences between zones near the front and rear walls and particularly high NO(x) emissions were found in the boiler. With focus on minimizing these problems, a new technology based on multiple-injection and multiple-staging has been developed. Combustion improvements and NO(x) reductions were validated by investigating three aspects. First, numerical simulations of the pulverized-coal combustion process and NO(x) emissions were compared in both the original and new technologies. Good agreement was found between simulations and in situ measurements with the original technology. Second, with the new technology, gas temperature and concentration distributions were found to be symmetric near the front and rear walls. A relatively low-temperature and high-oxygen-concentration zone formed in the near-wall region that helps mitigate slagging in the lower furnace. Third, NO(x) emissions were found to have decreased by as much as 50%, yielding a slight decrease in the levels of unburnt carbon in the fly ash.

Feasibility and Performance of the Microwave Thermal Rocket Launcher

OpenAIRE

Parkin, Kevin L. G.; Culick, Fred E. C.

2004-01-01

Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit...

A Concept of Two-Stage-To-Orbit Reusable Launch Vehicle

Science.gov (United States)

Yang, Yong; Wang, Xiaojun; Tang, Yihua

2002-01-01

Reusable Launch Vehicle (RLV) has a capability of delivering a wide rang of payload to earth orbit with greater reliability, lower cost, more flexibility and operability than any of today's launch vehicles. It is the goal of future space transportation systems. Past experience on single stage to orbit (SSTO) RLVs, such as NASA's NASP project, which aims at developing an rocket-based combined-cycle (RBCC) airplane and X-33, which aims at developing a rocket RLV, indicates that SSTO RLV can not be realized in the next few years based on the state-of-the-art technologies. This paper presents a concept of all rocket two-stage-to-orbit (TSTO) reusable launch vehicle. The TSTO RLV comprises an orbiter and a booster stage. The orbiter is mounted on the top of the booster stage. The TSTO RLV takes off vertically. At the altitude about 50km the booster stage is separated from the orbiter, returns and lands by parachutes and airbags, or lands horizontally by means of its own propulsion system. The orbiter continues its ascent flight and delivers the payload into LEO orbit. After completing orbit mission, the orbiter will reenter into the atmosphere, automatically fly to the ground base and finally horizontally land on the runway. TSTO RLV has less technology difficulties and risk than SSTO, and maybe the practical approach to the RLV in the near future.

Fluid-bed combustion

Energy Technology Data Exchange (ETDEWEB)

Hunt, G.; Schoebotham, N.

1981-02-01

In Energy Equipment Company's two-stage fluidized bed system, partial combustion in a fluidized bed is followed by burn-off of the generated gases above the bed. The system can be retrofitted to existing boilers, and can burn small, high ash coal efficiently. It has advantages when used as a hot gas generator for process drying. Tests on a boiler at a Cadbury Schweppes plant are reported.

Combustion, cofiring and emissions characteristics of torrefied biomass in a drop tube reactor

International Nuclear Information System (INIS)

Ndibe, Collins; Maier, Jörg; Scheffknecht, Günter

2015-01-01

The study investigates cofiring characteristics of torrefied biomass fuels at 50% thermal shares with coals and 100% combustion cases. Experiments were carried out in a 20 kW, electrically heated, drop-tube reactor. Fuels used include a range of torrefied biomass fuels, non-thermally treated white wood pellets, a high volatile bituminous coal and a lignite coal. The reactor was maintained at 1200 °C while the overall stoichiometric ratio was kept constant at 1.15 for all combustion cases. Measurements were performed to evaluate combustion reactivity, emissions and burn-out. Torrefied biomass fuels in comparison to non-thermally treated wood contain a lower amount of volatiles. For the tests performed at a similar particle size distribution, the reduced volatile content did not impact combustion reactivity significantly. Delay in combustion was only observed for test fuel with a lower amount of fine particles. The particle size distribution of the pulverised grinds therefore impacts combustion reactivity more. Sulphur and nitrogen contents of woody biomass fuels are low. Blending woody biomass with coal lowers the emissions of SO 2 mainly as a result of dilution. NO X emissions have a more complex dependency on the nitrogen content. Factors such as volatile content of the fuels, fuel type, furnace and burner configurations also impact the final NO X emissions. In comparison to unstaged combustion, the nitrogen conversion to NO X declined from 34% to 9% for air-staged co-combustion of torrefied biomass and hard coal. For the air-staged mono-combustion cases, nitrogen conversion to NO X declined from between 42% and 48% to about 10%–14%. - Highlights: • Impact of torrefaction on cofiring was studied at high heating rates in a drop tube. • Cofiring of torrefied biomasses at high thermal shares (50% and higher) is feasible. • Particle size impacts biomass combustion reactivity more than torrefaction. • In a drop tube reactor, torrefaction has no negative

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

Science.gov (United States)

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

2003-01-01

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

Liquid Rocket Engine Testing

Science.gov (United States)

2016-10-21

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

Additive Manufacturing of Low Cost Upper Stage Propulsion Components

Science.gov (United States)

Protz, Christopher; Bowman, Randy; Cooper, Ken; Fikes, John; Taminger, Karen; Wright, Belinda

2014-01-01

NASA is currently developing Additive Manufacturing (AM) technologies and design tools aimed at reducing the costs and manufacturing time of regeneratively cooled rocket engine components. These Low Cost Upper Stage Propulsion (LCUSP) tasks are funded through NASA's Game Changing Development Program in the Space Technology Mission Directorate. The LCUSP project will develop a copper alloy additive manufacturing design process and develop and optimize the Electron Beam Freeform Fabrication (EBF3) manufacturing process to direct deposit a nickel alloy structural jacket and manifolds onto an SLM manufactured GRCop chamber and Ni-alloy nozzle. In order to develop these processes, the project will characterize both the microstructural and mechanical properties of the SLMproduced GRCop-84, and will explore and document novel design techniques specific to AM combustion devices components. These manufacturing technologies will be used to build a 25K-class regenerative chamber and nozzle (to be used with tested DMLS injectors) that will be tested individually and as a system in hot fire tests to demonstrate the applicability of the technologies. These tasks are expected to bring costs and manufacturing time down as spacecraft propulsion systems typically comprise more than 70% of the total vehicle cost and account for a significant portion of the development schedule. Additionally, high pressure/high temperature combustion chambers and nozzles must be regeneratively cooled to survive their operating environment, causing their design to be time consuming and costly to build. LCUSP presents an opportunity to develop and demonstrate a process that can infuse these technologies into industry, build competition, and drive down costs of future engines.

Baking Soda and Vinegar Rockets

Science.gov (United States)

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

2009-01-01

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

Understanding Kelvin-Helmholtz instability in paraffin-based hybrid rocket fuels

Science.gov (United States)

Petrarolo, Anna; Kobald, Mario; Schlechtriem, Stefan

2018-04-01

Liquefying fuels show higher regression rates than the classical polymeric ones. They are able to form, along their burning surface, a low viscosity and surface tension liquid layer, which can become unstable (Kelvin-Helmholtz instability) due to the high velocity gas flow in the fuel port. This causes entrainment of liquid droplets from the fuel surface into the oxidizer gas flow. To better understand the droplets entrainment mechanism, optical investigations on the combustion behaviour of paraffin-based hybrid rocket fuels in combination with gaseous oxygen have been conducted in the framework of this research. Combustion tests were performed in a 2D single-slab burner at atmospheric conditions. High speed videos were recorded and analysed with two decomposition techniques. Proper orthogonal decomposition (POD) and independent component analysis (ICA) were applied to the scalar field of the flame luminosity. The most excited frequencies and wavelengths of the wave-like structures characterizing the liquid melt layer were computed. The fuel slab viscosity and the oxidizer mass flow were varied to study their influence on the liquid layer instability process. The combustion is dominated by periodic, wave-like structures for all the analysed fuels. Frequencies and wavelengths characterizing the liquid melt layer depend on the fuel viscosity and oxidizer mass flow. Moreover, for very low mass flows, no wavelength peaks are detected for the higher viscosity fuels. This is important to better understand and predict the onset and development of the entrainment process, which is connected to the amplification of the longitudinal waves.

Multi-Rocket Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

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

Rocket Science 101 Interactive Educational Program

Science.gov (United States)

Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

2007-01-01

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

Design and Experimental Study on Spinning Solid Rocket Motor

Science.gov (United States)

Xue, Heng; Jiang, Chunlan; Wang, Zaicheng

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

Impact of higher n-butanol addition on combustion and performance of GDI engine in stoichiometric combustion

International Nuclear Information System (INIS)

Chen, Zheng; Yang, Feng; Xue, Shuo; Wu, Zhenkuo; Liu, Jingping

2015-01-01

Highlights: • Effects of 0–50% n-butanol addition on GDI engine are experimentally studied. • Higher n-butanol fractions increase combustion pressure and fasten burning rate. • Higher n-butanol fractions increase BSFC but improve BTE. • Higher n-butanol fractions enhance combustion stability but increase knock intensity. • Higher n-butanol fractions reduce exhaust temperature and NOx emissions. - Abstract: An experimental study was carried out on a turbocharged gasoline direct injection (GDI) engine fueled by n-butanol/gasoline blends. Effects of n-butanol percents (15%, 30%, and 50%) on combustion and performance of the engine operating on stoichiometric combustion condition were discussed and also compared with pure gasoline in this paper. The results indicate that n-butanol/gasoline blends increase combustion pressure and pressure rise rate, fasten burning rate, and shorten ignition delay and combustion duration, as compared to pure gasoline. Moreover, these trends are impacted more evidently with increased n-butanol fraction in the blends. In addition, higher n-butanol percent of gasoline blends increase combustion temperature but decrease the temperature in the later stage of expansion stroke, which contributes to the control of exhaust temperature at high-load. With regards to engine performance, higher n-butanol percent in the blends results in increased brake specific fuel consumption (BSFC) and higher brake thermal efficiency (BTE). However, higher n-butanol addition helps to improve combustion stability but shows slightly higher knock possibility in high-load. In that case, the knock trend could be weakened by retarding ignition timing. Moreover, higher n-butanol addition significantly decreases NOx emissions, but it increases CO emissions obviously.

Computer Program for Analysis, Design and Optimization of Propulsion, Dynamics, and Kinematics of Multistage Rockets

Science.gov (United States)

Lali, Mehdi

2009-03-01

A comprehensive computer program is designed in MATLAB to analyze, design and optimize the propulsion, dynamics, thermodynamics, and kinematics of any serial multi-staging rocket for a set of given data. The program is quite user-friendly. It comprises two main sections: "analysis and design" and "optimization." Each section has a GUI (Graphical User Interface) in which the rocket's data are entered by the user and by which the program is run. The first section analyzes the performance of the rocket that is previously devised by the user. Numerous plots and subplots are provided to display the performance of the rocket. The second section of the program finds the "optimum trajectory" via billions of iterations and computations which are done through sophisticated algorithms using numerical methods and incremental integrations. Innovative techniques are applied to calculate the optimal parameters for the engine and designing the "optimal pitch program." This computer program is stand-alone in such a way that it calculates almost every design parameter in regards to rocket propulsion and dynamics. It is meant to be used for actual launch operations as well as educational and research purposes.

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

Science.gov (United States)

Arrighi, Robert S.

2016-01-01

Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

Improving of technical characteristics of launch vehicles with liquid rocket engines using active onboard de-orbiting systems

Science.gov (United States)

Trushlyakov, V.; Shatrov, Ya.

2017-09-01

In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing. Factual technical characteristics on the example of a promising type of rocket ;Soyuz-2.1.v.; are analyzed. Meeting the TR in relation to anthropogenic impact decrease based on the conventional design approach and the content of the onboard system does not prove to be efficient and leads to depreciation of the initial technical characteristics obtained at the first design stage if these requirements are not included. In this concern, it is shown that the implementation of additional active onboard de-orbiting system (AODS) of worked-off stages (WS) into the onboard LV stages systems allows to meet the TR related to the LV environmental characteristics, including fire-explosion safety. In some cases, the orbital payload mass increases.

High Thermal Conductivity NARloy-Z-Diamond Composite Liner for Advanced Rocket Engines

Science.gov (United States)

Bhat, Biliyar; Greene, Sandra

2015-01-01

NARloy-Z (Cu-3Ag-0.5Zr) alloy is state-of-the-art combustion chamber liner material used in liquid propulsion engines such as the RS-68 and RS-25. The performance of future liquid propulsion systems can be improved significantly by increasing the heat transfer through the combustion chamber liner. Prior work1 done at NASA Marshall Space Flight Center (MSFC) has shown that the thermal conductivity of NARloy-Z alloy can be improved significantly by embedding high thermal conductivity diamond particles in the alloy matrix to form NARloy-Z-diamond composite (fig. 1). NARloy-Z-diamond composite containing 40vol% diamond showed 69% higher thermal conductivity than NARloy-Z. It is 24% lighter than NARloy-Z and hence the density normalized thermal conductivity is 120% better. These attributes will improve the performance and life of the advanced rocket engines significantly. The research work consists of (a) developing design properties (thermal and mechanical) of NARloy-Z-D composite, (b) fabrication of net shape subscale combustion chamber liner, and (c) hot-fire testing of the liner to test performance. Initially, NARloy-Z-D composite slabs were made using the Field Assisted Sintering Technology (FAST) for the purpose of determining design properties. In the next step, a cylindrical shape was fabricated to demonstrate feasibility (fig. 3). The liner consists of six cylinders which are sintered separately and then stacked and diffusion bonded to make the liner (fig. 4). The liner will be heat treated, finish-machined, and assembled into a combustion chamber and hot-fire tested in the MSFC test facility (TF 115) to determine perform.

Effect of Chamber Pressurization Rate on Combustion and Propagation of Solid Propellant Cracks

Science.gov (United States)

Yuan, Wei-Lan; Wei, Shen; Yuan, Shu-Shen

2002-01-01

area of the propellant grain satisfies the designed value. But cracks in propellant grain can be generated during manufacture, storage, handing and so on. The cracks can provide additional surface area for combustion. The additional combustion may significantly deviate the performance of the rocket motor from the designed conditions, even lead to explosive catastrophe. Therefore a thorough study on the combustion, propagation and fracture of solid propellant cracks must be conducted. This paper takes an isolated propellant crack as the object and studies the effect of chamber pressurization rate on the combustion, propagation and fracture of the crack by experiment and theoretical calculation. deformable, the burning inside a solid propellant crack is a coupling of solid mechanics and combustion dynamics. In this paper, a theoretical model describing the combustion, propagation and fracture of the crack was formulated and solved numerically. The interaction of structural deformation and combustion process was included in the theoretical model. The conservation equations for compressible fluid flow, the equation of state for perfect gas, the heat conducting equation for the solid-phase, constitutive equation for propellant, J-integral fracture criterion and so on are used in the model. The convective burning inside the crack and the propagation and fracture of the crack were numerically studied by solving the set of nonlinear, inhomogeneous gas-phase governing equations and solid-phase equations. On the other hand, the combustion experiments for propellant specimens with a precut crack were conducted by RTR system. Predicted results are in good agreement with experimental data, which validates the reasonableness of the theoretical model. Both theoretical and experimental results indicate that the chamber pressurization rate has strong effects on the convective burning in the crack, crack fracture initiation and fracture pattern.

Combustion of drops of Mexican fuel oils with high asphaltenes content; Combustion de gotas de combustoleos mexicanos con alto contenido de asfaltenos

Energy Technology Data Exchange (ETDEWEB)

Garcia Rodriguez, Jose Francisco [Instituto de Investigaciones Electricas, Cuernavaca (Mexico)

1998-09-01

In this work the combustion of fuel drops with a content of 18% of asphaltenes has been studied . The results obtained for this fuel were compared with the ones obtained for another with a content of 12% asphaltenes. The drops were suspended in a platinum filament and burned in an spherical radiant furnace. The drop size varied between 600 and 800 microns. The fuel drops with 12% asphaltenes showed shorter combustion times, a smaller diameter increment of the smaller diameter during the combustion stages and also a shorter burning time of the carbonaceous residue than the fuel drops with a content of 18% asphaltenes. [Espanol] En el presente trabajo se ha estudiado la combustion de gotas de combustible con 18% de contenido de asfaltenos. Los resultados obtenidos para este combustible se compararon con los obtenidos para otro con 12% de contenido de asfaltenos. Las gotas fueron suspendidas en un filamento de platino y quemadas en un horno radiante esferico. El tamano de las gotas vario entre 600 y 800 micras. Las gotas de combustible con 12% de asfaltenos mostraron tiempos de combustion mas cortos, un incremento del diametro menor durante las etapas de combustion y un tiempo de quemado del residuo carbonoso tambien mas corto que las gotas del combustible con 18% de contenido de asfaltenos.

Design and experimental investigation of an oxy-fuel combustion system for magnetohydrodynamic power extraction

Science.gov (United States)

Hernandez, Manuel Johannes

design concepts. Therefore, numerical computational fluid dynamics (CFD) models were developed to design and optimize the combustion flow fields of oxy-fuel combustion systems. These models were analyzed to understand the boundary layer and heat transfer profile and qualitative behaviors in the product designs. Advanced materials for high-temperature applications were assessed for their possible implementation in the product design. A trade-off analysis indicated that this scheme may incur elevated product cost and a difficulty in manufacturing. Active cooling strategies were considered for product development. A rocket-based cooling scheme, regenerative cooling, was implemented to provide active cooling. In the hot gas path (HGP) cooling design, CFD models were developed to predict the variation of heat removal along the oxy-combustion wall for various operating conditions. The oxy-combustion technology was manufactured using electrical discharge machining (EDM). The product development lifecycle in this dissertation encompassed preliminary design, detailed design, and demonstration and validation of the product. Towards the final stages of the product development, Fuel-rich oxy combustion experiments were carried out to demonstrate and observe flame characteristics from the designed technology and to predict heat transfer loads. The demonstration findings of oxy-combustion flames are presented in this work to contribute the developing field of MHD direct power extraction, which lacks oxy-combustion design data and qualitative combustion datasets. The findings show that this oxy-combustion concept is capable of providing a high-enthalpy MHD environment for seeding, in order to render the flow to be conductive. Based on previous findings, temperatures in the range of 2800-3000 K may enable magnetohydrodynamic power extraction. The combustor hardware design was developed to contribute to engineered systems rated less than 100 kW for demonstration. The product hardware was

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

Science.gov (United States)

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

2009-09-01

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

Design study of laser fusion rocket

International Nuclear Information System (INIS)

Nakashima, Hideki; Shoyama, Hidetoshi; Kanda, Yukinori

1991-01-01

A design study was made on a rocket powered by laser fusion. Dependence of its flight performance on target gain, driver repetition rate and fuel composition was analyzed to obtain optimal design parameters of the laser fusion rocket. The results indicate that the laser fusion rocket fueled with DT or D 3 He has the potential advantages over other propulsion systems such as fission rocket for interplanetary travel. (author)

Thermogravimetric investigation of the co-combustion between the pyrolysis oil distillation residue and lignite.

Science.gov (United States)

Li, Hao; Xia, Shuqian; Ma, Peisheng

2016-10-01

Co-combustion of lignite with distillation residue derived from rice straw pyrolysis oil was investigated by non-isothermal thermogravimetric analysis (TGA). The addition of distillation residue improved the reactivity and combustion efficiency of lignite, such as increasing the weight loss rate at peak temperature and decreasing the burnout temperature and the total burnout. With increasing distillation residue content in the blended fuels, the synergistic interactions between distillation residue and lignite firstly increased and then decreased during co-combustion stage. Results of XRF, FTIR, (13)C NMR and SEM analysis indicated that chemical structure, mineral components and morphology of samples have great influence on the synergistic interactions. The combustion mechanisms and kinetic parameters were calculated by the Coats Redfern model, suggesting that the lowest apparent activation energy (120.19kJ/mol) for the blended fuels was obtained by blending 60wt.% distillation residue during main co-combustion stage. Copyright © 2016 Elsevier Ltd. All rights reserved.

Ionospheric shock waves triggered by rockets

Directory of Open Access Journals (Sweden)

C. H. Lin

2014-09-01

Full Text Available This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK Taepodong-2 and 2013 South Korea (SK Korea Space Launch Vehicle-II (KSLV-II rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100–600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800–1200 m s−1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800–1200 m s−1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.

Influence of Coal Quality on Combustion Performance

DEFF Research Database (Denmark)

Lans, Robert Pieter Van Der; Glarborg, Peter; Dam-Johansen, Kim

1998-01-01

mixing pattern on NO formation under these conditions. Emissions from the opposed fired plant with all combustion air introduced through the burners could only be qualitatively reproduced by the pilot furnace. Under single stage conditions the test rig provided higher NO levels. Carbon in ash levels did...... not show any correlation between the coals and the furnaces. An engineering, mathematical model has been developed describing radiation heat transfer and coal combustion in full scale furnaces. The model has been validated against measured temperatures and the amount of carbon in fly ash. The model...

Nonlinear rocket motor stability prediction: Limit amplitude, triggering, and mean pressure shifta)

Science.gov (United States)

Flandro, Gary A.; Fischbach, Sean R.; Majdalani, Joseph

2007-09-01

High-amplitude pressure oscillations in solid propellant rocket motor combustion chambers display nonlinear effects including: (1) limit cycle behavior in which the fluctuations may dwell for a considerable period of time near their peak amplitude, (2) elevated mean chamber pressure (DC shift), and (3) a triggering amplitude above which pulsing will cause an apparently stable system to transition to violent oscillations. Along with the obvious undesirable vibrations, these features constitute the most damaging impact of combustion instability on system reliability and structural integrity. The physical mechanisms behind these phenomena and their relationship to motor geometry and physical parameters must, therefore, be fully understood if instability is to be avoided in the design process, or if effective corrective measures must be devised during system development. Predictive algorithms now in use have limited ability to characterize the actual time evolution of the oscillations, and they do not supply the motor designer with information regarding peak amplitudes or the associated critical triggering amplitudes. A pivotal missing element is the ability to predict the mean pressure shift; clearly, the designer requires information regarding the maximum chamber pressure that might be experienced during motor operation. In this paper, a comprehensive nonlinear combustion instability model is described that supplies vital information. The central role played by steep-fronted waves is emphasized. The resulting algorithm provides both detailed physical models of nonlinear instability phenomena and the critically needed predictive capability. In particular, the origin of the DC shift is revealed.

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

Science.gov (United States)

Gordon, Sanford; Kastner, Michael E

1958-01-01

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

Experimental Study of the Swirling Oxidizer Flow in HTPB/N2O Hybrid Rocket Motor

Directory of Open Access Journals (Sweden)

Mohammad Mahdi Heydari

2017-01-01

Full Text Available Effects of swirling oxidizer flow on the performance of a HTPB/N2O Hybrid rocket motor were studied. A hybrid propulsion laboratory has been developed, to characterize internal ballistics characteristics of swirl flow hybrid motors and to define the operating parameters, like fuel regression rate, specific impulse, and characteristics velocity and combustion efficiency. Primitive variables, like pressure, thrust, temperature, and the oxidizer mass flow rate, were logged. A modular motor with 70 mm outer diameter and variable chamber length is designed for experimental analysis. The injector module has four tangential injectors and one axial injector. Liquid nitrous oxide (N2O as an oxidizer is injected at the head of combustion chamber into the motor. The feed system uses pressurized air as the pressurant. Two sets of tests have been performed. Some tests with axial and tangential oxidizer injection and a test with axial oxidizer injection were done. The test results show that the fuel grain regression rate has been improved by applying tangential oxidizer injection at the head of the motor. Besides, it was seen that combustion efficiency of motors with the swirl flow was about 10 percent more than motors with axial flow.

Oxy-fuel combustion characteristics and kinetics of microalgae Chlorella vulgaris by thermogravimetric analysis.

Science.gov (United States)

Chen, Chunxiang; Lu, Ziguang; Ma, Xiaoqian; Long, Jun; Peng, Yuning; Hu, Likun; Lu, Quan

2013-09-01

Oxy-fuel or O2/CO2 combustion technology was used to investigate the combustion of Chlorella vulgaris by thermogravimetric analysis (TGA). Oxy-fuel combustion occurs in an O2/CO2 atmosphere instead of an O2/N2 atmosphere and offers an alternative method of C. vulgaris preparation for biofuels processing. Our results show that three stages were observed during C. vulgaris combustion and the main combustion process occurred at the second stage. Compared with a 20%O2/80%N2 atmosphere, the mass loss rate at the DTG peaks (Rp) and the average reaction rate (Rv) in a 20%O2/80%CO2 atmosphere was lower, while the ignition temperature (TI) was higher. As oxygen concentration increases in an O2/CO2 atmosphere, Rp, Rv and the apparent activation energy (E) increases, while TI, the final temperature detected as mass stabilization (Tf) and the residue mass (Mr) decreases; As the heating rate (β) increases, TI, Tf and Rp increase, while Mr decreases. Crown Copyright © 2013. Published by Elsevier Ltd. All rights reserved.

Burnout behaviour of bituminous coals in air-staged combustion conditions

Energy Technology Data Exchange (ETDEWEB)

Kluger, F.; Spliethoff, H.; Hein, K.R.G. [University of Stuttgart, Stuttgart (Germany). Inst. of Process Engineering and Power Plant (IVD)

2001-07-01

In order to determine the influence on burnout by the combustion conditions and the coal preparation, three bituminous coals sold on the world market, from three different locations in Poland, South Africa, and Australia, were studied more closely. For this purpose, the coals were ground in two different particle size ranges, which, besides the influence of the combustion conditions, such a temperature, residence time, and stoichiometry, made it possible to also investigate the impact on burnout by the coal preparation. The experiments were carried out in an electrically heated entrained-flow reactor with a thermal input of 8.5 kW. The parameters for the experiments are wall temperature (1000-1350{degree}C), air ratio (0.6-1.15) and two particle sizes (70% {lt} 75 {mu}m, 90% {lt} 75 {mu}m). The results show that in general, for increasing temperatures, the burnout quality will improve. For the Australian Illawara coal, another outcome is increased NOx emissions. Lowering the air ratio in the reduction zone leads to less NOx emission but to increased unburnt matter in ash. For the smaller particle size fraction, the analysis of the different particle sizes shows an improvement of the burnout without a change in NOx emissions. 10 refs., 10 figs., 2 tabs.

The influence of charge stratification on the spectral signature of partially premixed combustion in a light-duty optical engine

KAUST Repository

Najafabadi, M. Izadi; Egelmeers, Luc; Somers, Bart; Deen, Niels; Johansson, Bengt; Dam, Nico

2017-01-01

* chemiluminescence, we see two different chemical mechanisms for formation of C2* in the PPC regime: first during the early stage of combustion by the breakup of bigger molecules and the second during the late stage of combustion when soot particles are forming.

High-speed schlieren imaging of rocket exhaust plumes

Science.gov (United States)

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

2016-11-01

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

A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

Science.gov (United States)

Stanley, Thomas Troy; Alexander, Reginald

1999-01-01

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

Safety and Environment- Masterplan 2020 of DLR's Rocket Test Center Lampoldhausen

Science.gov (United States)

Haberzettl, Andreas; Dommers, Michael

2013-09-01

The German Aerospace Center DLR is the German research institute with approximately 7000 employees in 16 domestic locations. Among the research priorities of the German Aerospace Center DLR includes aerospace, energy and transport. DLR is institutionally supported by federal and state governments.Next funding sources arise in the context of third-party funds business (contract research and public contracts and subsidiaries). Main activities of the test center Lampoldshausen are testing of ARIANE's main and upper stage engines in the frame of ESA contracts.In the last years the test center of the DLR in Lampoldshausen has grown strongly, so that the number of employees is actually of about 230. The testing department is mainly responsible for rocket combustion testing according to customer requirements.Two kinds of test facilities are operated, sea level test benches and the altitude simulation test facilities.In addition to the DLR's growth also the activities of the industrial partner ASTRIUM has been elevated so that actually nearly 600 employees are present on site Lampoldshausen.The management of the site in relation to safety and security requires special measures with special respect to the presence of more people inside the testing area in order to guarantee trouble-free and safe experimental operation onsite the DLR's test plants. In order to meet with the future needs of continuing growth, the security and safety requirements have to be adopted.This report gives comprehensive outlook information about future possible scenarios of our coming tasks.Main driving force for future requests is the evolution of the rocket ARIANE. The testing of the new upper stage test facility for ARIANE 5 midlife evolution has been started. A new test position P5.2 is foreseen to perform the qualification of the new upper stage with the VINCI engine. This project will be very complex, in parallel running operation processes will require special procedures related to the overall

Combustion of Pure, Hydrolyzed and Methyl Ester Formed of Jatropha Curcas Lin oil

Directory of Open Access Journals (Sweden)

Muhaji Muhaji

2015-10-01

Full Text Available The density and viscosity of vegetable oil are higher than that of diesel oil. Thus its direct combustion in the diesel engine results many problems. This research was conducted to investigate the flame characteristics of combustion of jatropha curcas lin in pure, hydrolyzed and methyl ester form. The results indicated that the combustion of pure jatropha curcas lin occurs in three stages, hydrolyzed in two stages    and methyl ester in one stage. For pure jatropha curcas lin, in the first stage, unsaturated fatty acid burned for  0.265 s.  It is followed by saturated fatty acid, burned for 0.389 s in the second stage. And, in the last stage is the burned of glycerol for 0.560 s. Meanwhile for hydrolyzed one, in the first stage, unsaturated fatty acid burned for 0.736 s, followed by saturated fatty acid, burned  for 0.326 s in the second stage. And the last, for methyl ester is the burned for 0.712 s. The highest burning rate was for methyl ester which was 0.003931cc/s. The energy releasing rate of methyl ester, which was for 13,628.67 kcal/(kg.s resembled that of diesel oil the most, while the lowest rate was for pure jatropha curcas lin which was 8,200.94 kcal/(kg.s. In addition, massive explosion occurred in the fuel containing unsaturated fatty acid and glycerol

Nitrogen Chemistry in Fluidized Bed Combustion of Coal

DEFF Research Database (Denmark)

Jensen, Anker Degn

and reduction by homogeneous and heterogeneous reactions. The data for the estimation of kinetics of the heterogeneous reactions were measured by one of the partners in the project for char and bed material sampled from a pressurized FBC pilot plant burning Kiveton Park coal. Experimental data from the pilot...... plant were used for model verification. The simulations of the NO emission during staged combustion and NH3 injection for NO reduction were in qualitative agreement with the experimental data. A parametric study of the influence of operating conditions on the conversion of fuel-N to NO showed......, the gas interchange coefficient, the bubble size and the bubble rise velocity. The most important combustion parameters were the rate of CO and CH4 combustion and the fraction of CO produced from char combustion. By using a rate of production analysis, the important reactions in the NO model were...

16 CFR 1507.10 - Rockets with sticks.

Science.gov (United States)

2010-01-01

... 16 Commercial Practices 2 2010-01-01 2010-01-01 false Rockets with sticks. 1507.10 Section 1507.10... FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall utilize a straight and rigid stick to provide a direct and stable flight. Such sticks shall...

Results from flamelet and non-flamelet models for supersonic combustion

Science.gov (United States)

Ladeinde, Foluso; Li, Wenhai

2017-11-01

Air-breathing propulsion systems (scramjets) have been identified as a viable alternative to rocket engines for improved efficiency. A scramjet engine, which operates at flight Mach numbers around 7 or above, is characterized by the existence of supersonic flow conditions in the combustor. In a dual-mode scramjet, this phenomenon is possible because of the relatively low value of the equivalence ratio and high stagnation temperature, which, together, inhibits thermal choking downstream of transverse injectors. The flamelet method has been our choice for turbulence-combustion interaction modeling and we have extended the basic approach in several dimensions, with a focus on the way the pressure and progress variable are modeled. Improved results have been obtained. We have also examined non-flamelet models, including laminar chemistry (QL), eddy dissipation concept (EDC), and partially-stirred reactor (PaSR). The pressure/progress variable-corrected simulations give better results compared with the original model, with reaction rates that are lower than those from EDC and PaSR. In general, QL tends to over-predict the reaction rate for the supersonic combustion problems investigated in our work.

Five-Segment Solid Rocket Motor Development Status

Science.gov (United States)

Priskos, Alex S.

2012-01-01

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

The environmental assessment of the wood combustion

International Nuclear Information System (INIS)

Dinca, Cristian; Badea, Adrian; Apostol, Tiberiu

2007-01-01

In this paper, the authors analysed the emissions from residential boilers burning wood logs, bark pellets, wood briquettes and wood pellets. Three boilers, selected with respect to age, design, connection to heat storage tank, and type of biofuel, were included in the study. The emissions captured comprised carbon monoxide (CO), carbon dioxide (CO 2 ), oxygen (O 2 ), total organic carbons (TOC), nitrogen oxides (NO x ), polycyclic aromatic hydrocarbons (PAC) and 33 volatile organic compounds (VOC). We have used the Life Cycle Inventory method in order to identify the main stressors generated by the wood combustion stage. In this purpose, we have analysed one type of old boiler, one type of modern boiler and a multi-fuel boiler, which can burn wood logs, bark pellets, wood briquettes and wood pellets. In this article, we selected only the wood combustion stage because it is the most important according to the emissions produced. (authors)

Direct Numerical Simulation of Turbulent Multi-Stage Autoignition Relevant to Engine Conditions

Science.gov (United States)

Chen, Jacqueline

2017-11-01

Due to the unrivaled energy density of liquid hydrocarbon fuels combustion will continue to provide over 80% of the world's energy for at least the next fifty years. Hence, combustion needs to be understood and controlled to optimize combustion systems for efficiency to prevent further climate change, to reduce emissions and to ensure U.S. energy security. In this talk I will discuss recent progress in direct numerical simulations of turbulent combustion focused on providing fundamental insights into key `turbulence-chemistry' interactions that underpin the development of next generation fuel efficient, fuel flexible engines for transportation and power generation. Petascale direct numerical simulation (DNS) of multi-stage mixed-mode turbulent combustion in canonical configurations have elucidated key physics that govern autoignition and flame stabilization in engines and provide benchmark data for combustion model development under the conditions of advanced engines which operate near combustion limits to maximize efficiency and minimize emissions. Mixed-mode combustion refers to premixed or partially-premixed flames propagating into stratified autoignitive mixtures. Multi-stage ignition refers to hydrocarbon fuels with negative temperature coefficient behavior that undergo sequential low- and high-temperature autoignition. Key issues that will be discussed include: 1) the role of mixing in shear driven turbulence on the dynamics of multi-stage autoignition and cool flame propagation in diesel environments, 2) the role of thermal and composition stratification on the evolution of the balance of mixed combustion modes - flame propagation versus spontaneous ignition - which determines the overall combustion rate in autoignition processes, and 3) the role of cool flames on lifted flame stabilization. Finally prospects for DNS of turbulent combustion at the exascale will be discussed in the context of anticipated heterogeneous machine architectures. sponsored by DOE

Two technicians apply insulation to S-II second stage

Science.gov (United States)

1964-01-01

Two technicians apply insulation to the outer surface of the S-II second stage booster for the Saturn V moon rocket. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

JANNAF "Test and Evaluation Guidelines for Liquid Rocket Engines": Status and Application

Science.gov (United States)

Parkinson, Douglas; VanLerberghe, Wayne M.; Rahman, Shamim A.

2017-01-01

For many decades, the U.S. rocket propulsion industrial base has performed remarkably in developing complex liquid rocket engines that can propel critical payloads into service for the nation, as well as transport people and hardware for missions that open the frontiers of space exploration for humanity. This has been possible only at considerable expense given the lack of detailed guidance that captures the essence of successful practices and knowledge accumulated over five decades of liquid rocket engine development. In an effort to provide benchmarks and guidance for the next generation of rocket engineers, the Joint Army Navy NASA Air Force (JANNAF) Interagency Propulsion Committee published a liquid rocket engine (LRE) test and evaluation (T&E) guideline document in 2012 focusing on the development challenges and test verification considerations for liquid rocket engine systems. This document has been well received and applied by many current LRE developers as a benchmark and guidance tool, both for government-driven applications as well as for fully commercial ventures. The USAF Space and Missile Systems Center (SMC) has taken an additional near-term step and is directing activity to adapt and augment the content from the JANNAF LRE T&E guideline into a standard for potential application to future USAF requests for proposals for LRE development initiatives and launch vehicles for national security missions. A draft of this standard was already sent out for review and comment, and is intended to be formally approved and released towards the end of 2017. The acceptance and use of the LRE T&E guideline is possible through broad government and industry participation in the JANNAF liquid propulsion committee and associated panels. The sponsoring JANNAF community is expanding upon this initial baseline version and delving into further critical development aspects of liquid rocket propulsion testing at the integrated stage level as well as engine component level, in

MODELING OF FUEL SPRAY CHARACTERISTICS AND DIESEL COMBUSTION CHAMBER PARAMETERS

Directory of Open Access Journals (Sweden)

G. M. Kukharonak

2011-01-01

Full Text Available The computer model for coordination of fuel spray characteristics with diesel combustion chamber parameters has been created in the paper.  The model allows to observe fuel sprays  develоpment in diesel cylinder at any moment of injection, to calculate characteristics of fuel sprays with due account of a shape and dimensions of a combustion chamber, timely to change fuel injection characteristics and supercharging parameters, shape and dimensions of a combustion chamber. Moreover the computer model permits to determine parameters of holes in an injector nozzle that provides the required fuel sprays characteristics at the stage of designing a diesel engine. Combustion chamber parameters for 4ЧН11/12.5 diesel engine have been determined in the paper.

Pressure-Equalizing Cradle for Booster Rocket Mounting

Science.gov (United States)

Rutan, Elbert L. (Inventor)

2015-01-01

A launch system and method improve the launch efficiency of a booster rocket and payload. A launch aircraft atop which the booster rocket is mounted in a cradle, is flown or towed to an elevation at which the booster rocket is released. The cradle provides for reduced structural requirements for the booster rocket by including a compressible layer, that may be provided by a plurality of gas or liquid-filled flexible chambers. The compressible layer contacts the booster rocket along most of the length of the booster rocket to distribute applied pressure, nearly eliminating bending loads. Distributing the pressure eliminates point loading conditions and bending moments that would otherwise be generated in the booster rocket structure during carrying. The chambers may be balloons distributed in rows and columns within the cradle or cylindrical chambers extending along a length of the cradle. The cradle may include a manifold communicating gas between chambers.

Subscale Winged Rocket Development and Application to Future Reusable Space Transportation

Directory of Open Access Journals (Sweden)

Koichi YONEMOTO

2018-03-01

Full Text Available Kyushu Institute of Technology has been studying unmanned suborbital winged rocket called WIRES (WInged REusable Sounding rocket and its research subjects concerning aerodynamics, NGC (Navigation, Guidance and Control, cryogenic composite tanks etc., and conducting flight demonstration of small winged rocket since 2005. WIRES employs the original aerodynamic shape of HIMES (HIghly Maneuverable Experimental Sounding rocket studied by ISAS (Institute of Space and Astronautical Science of JAXA (Japan Aerospace Exploration Agency in 1980s. This paper presents the preliminary design of subscale non-winged and winged rockets called WIRES#013 and WIRES#015, respectively, that are developed in collaboration with JAXA, USC (University of Southern California, UTEP (University of Texas at El Paso and Japanese industries. WIRES#013 is a conventional pre-test rocket propelled by two IPA-LOX (Isopropyl Alcohol and Liquid Oxygen engines under development by USC. It has the total length of 4.6m, and the weight of 1000kg to reach the altitude of about 6km. The flight objective is validation of the telemetry and ground communication system, recovery parachute system, and launch operation of liquid engine. WIRES#015, which has the same length of WIRES#013 and the weight of 1000kg, is a NGC technology demonstrator propelled by a fully expander-cycle LOX-Methane engine designed and developed by JAXA to reach the altitude more than 6km. The flight tests of both WIRES#013 and WIRES#015 will be conducted at the launch facility of FAR (Friends of Amateur Rocketry, Inc., which is located at Mojave Desert of California in United States of America, in May 2018 and March 2019 respectively. After completion of WIRES#015 flight tests, the suborbital demonstrator called WIRES-X will be developed and its first flight test well be performed in 2020. Its application to future fully reusable space transportation systems, such as suborbital space tour vehicles and two-stage

Environmental control implications of generating electric power from coal. 1977 technology status report. Appendix D. Assessment of NO/sub x/ control technology for coal fired utility boilers. [Low-excess-air, staged combustion, flu gas recirculation and burner design

Energy Technology Data Exchange (ETDEWEB)

1977-12-01

An NOx control technology assessment study was conducted to examine the effectiveness of low-excess-air firing, staged combustion, flue gas recirculation, and current burner/boiler designs as applied to coal-fired utility boilers. Significant variations in NOx emissions exist with boiler type, firing method, and coal type, but a relative comparison of emissions control performance, cost, and operational considerations is presented for each method. The study emphasized the numerous operational factors that are of major importance to the user in selecting and implementing a combustion modification technique. Staged combustion and low-excess-air operation were identified as the most cost-effective methods for existing units. Close control of local air/fuel ratios and rigorous combustion equipment maintenance are essential to the success of both methods. Flue gas recirculation is relatively ineffective and has the added concern of tube erosion. More research is needed to resolve potential corrosion concerns with low-NOx operating modes. Low-NOx burners in conjunction with a compartmentalized windbox are capable of meeting a 0.6-lb/million Btu emission level on new units. Advanced burner designs are being developed to meet research emission goals of approximately 0.25 lb/MBtu.

EXPERIMENTAL COMBUSTION ANALYSIS OF A HSDI DIESEL ENGINE FUELLED WITH PALM OIL BIODIESEL-DIESEL FUEL BLENDS

Directory of Open Access Journals (Sweden)

JOHN AGUDELO

2009-01-01

Full Text Available Differences in the chemical nature between petroleum diesel fuels and vegetable oils-based fuels lead to differences in their physical properties affecting the combustion process inside the engine. In this work a detailed combustion diagnosis was applied to a turbocharged automotive diesel engine operating with neat palm oil biodiesel (POB, No. 2 diesel fuel and their blends at 20 and 50% POB by volume (B20 and B50 respectively. To isolate the fuel effect, tests were executed at constant power output without carrying out any modification of the engine or its fuel injection system. As the POB content in the blend increased, there was a slight reduction in the fuel/air equivalence ratio from 0.39 (B0 to 0.37 (B100, an advance of injection timing and of start of combustion. Additionally, brake thermal efficiency, combustion duration, maximum mean temperature, temperature at exhaust valve opening and exhaust gas efficiency decreased; while the peak pressure, exergy destruction rate and specific fuel consumption increased. With diesel fuel and the blends B20 and B50 the same combustion stages were noticed. However, as a consequence of the differences pointed out, the thermal history of the process was affected. The diffusion combustion stage became larger with POB content. For B100 no premixed stage was observed.

Development and Performance of the 10 kN Hybrid Rocket Motor for the Stratos II Sounding Rocket

NARCIS (Netherlands)

Werner, R.M.; Knop, T.R.; Wink, J; Ehlen, J; Huijsman, R; Powell, S; Florea, R.; Wieling, W; Cervone, A.; Zandbergen, B.T.C.

2016-01-01

This paper presents the development work of the 10 kN hybrid rocket motor DHX-200 Aurora. The DHX-200 Aurora was developed by Delft Aerospace Rocket Engineering (DARE) to power the Stratos II and Stratos II+ sounding rocket, with the later one being launched in October 2015. Stratos II and Stratos

Failure characteristics analysis and fault diagnosis for liquid rocket engines

CERN Document Server

Zhang, Wei

2016-01-01

This book concentrates on the subject of health monitoring technology of Liquid Rocket Engine (LRE), including its failure analysis, fault diagnosis and fault prediction. Since no similar issue has been published, the failure pattern and mechanism analysis of the LRE from the system stage are of particular interest to the readers. Furthermore, application cases used to validate the efficacy of the fault diagnosis and prediction methods of the LRE are different from the others. The readers can learn the system stage modeling, analyzing and testing methods of the LRE system as well as corresponding fault diagnosis and prediction methods. This book will benefit researchers and students who are pursuing aerospace technology, fault detection, diagnostics and corresponding applications.

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

Science.gov (United States)

1972-01-01

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

Numerical Simulation of Combustion and Rotor-Stator Interaction in a Turbine Combustor

Directory of Open Access Journals (Sweden)

Dragos D. Isvoranu

2003-01-01

Full Text Available This article presents the development of a numerical algorithm for the computation of flow and combustion in a turbine combustor. The flow and combustion are modeled by the Reynolds-averaged Navier-Stokes equations coupled with the species-conservation equations. The chemistry model used herein is a two-step, global, finite-rate combustion model for methane and combustion gases. The governing equations are written in the strong conservation form and solved using a fully implicit, finite-difference approximation. The gas dynamics and chemistry equations are fully decoupled. A correction technique has been developed to enforce the conservation of mass fractions. The numerical algorithm developed herein has been used to investigate the flow and combustion in a one-stage turbine combustor.

The Swedish sounding rocket programme

International Nuclear Information System (INIS)

Bostroem, R.

1980-01-01

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

Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors

Science.gov (United States)

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

2014-01-01

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

Combustion synthesis of TiB2-based cermets: modeling and experimental results

International Nuclear Information System (INIS)

Martinez Pacheco, M.; Bouma, R.H.B.; Katgerman, L.

2008-01-01

TiB 2 -based cermets are prepared by combustion synthesis followed by a pressing stage in a granulate medium. Products obtained by combustion synthesis are characterized by a large remaining porosity (typically 50%). To produce dense cermets, a subsequent densification step is performed after the combustion process and when the reacted material is still hot. To design the process, numerical simulations are carried out and compared to experimental results. In addition, physical and electrical properties of the products related to electrical contact applications are evaluated. (orig.)

Combustion

CERN Document Server

Glassman, Irvin

2008-01-01

Combustion Engineering, a topic generally taught at the upper undergraduate and graduate level in most mechanical engineering programs, and many chemical engineering programs, is the study of rapid energy and mass transfer usually through the common physical phenomena of flame oxidation. It covers the physics and chemistry of this process and the engineering applications-from the generation of power such as the internal combustion automobile engine to the gas turbine engine. Renewed concerns about energy efficiency and fuel costs, along with continued concerns over toxic and particulate emissions have kept the interest in this vital area of engineering high and brought about new developments in both fundamental knowledge of flame and combustion physics as well as new technologies for flame and fuel control. *New chapter on new combustion concepts and technologies, including discussion on nanotechnology as related to combustion, as well as microgravity combustion, microcombustion, and catalytic combustion-all ...

Measuring Model Rocket Engine Thrust Curves

Science.gov (United States)

Penn, Kim; Slaton, William V.

2010-01-01

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

SAFE testing nuclear rockets economically

International Nuclear Information System (INIS)

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

2003-01-01

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

Sounding rockets explore the ionosphere

International Nuclear Information System (INIS)

Mendillo, M.

1990-01-01

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

A simplified computational fluid-dynamic approach to the oxidizer injector design in hybrid rockets

Science.gov (United States)

Di Martino, Giuseppe D.; Malgieri, Paolo; Carmicino, Carmine; Savino, Raffaele

2016-12-01

Fuel regression rate in hybrid rockets is non-negligibly affected by the oxidizer injection pattern. In this paper a simplified computational approach developed in an attempt to optimize the oxidizer injector design is discussed. Numerical simulations of the thermo-fluid-dynamic field in a hybrid rocket are carried out, with a commercial solver, to investigate into several injection configurations with the aim of increasing the fuel regression rate and minimizing the consumption unevenness, but still favoring the establishment of flow recirculation at the motor head end, which is generated with an axial nozzle injector and has been demonstrated to promote combustion stability, and both larger efficiency and regression rate. All the computations have been performed on the configuration of a lab-scale hybrid rocket motor available at the propulsion laboratory of the University of Naples with typical operating conditions. After a preliminary comparison between the two baseline limiting cases of an axial subsonic nozzle injector and a uniform injection through the prechamber, a parametric analysis has been carried out by varying the oxidizer jet flow divergence angle, as well as the grain port diameter and the oxidizer mass flux to study the effect of the flow divergence on heat transfer distribution over the fuel surface. Some experimental firing test data are presented, and, under the hypothesis that fuel regression rate and surface heat flux are proportional, the measured fuel consumption axial profiles are compared with the predicted surface heat flux showing fairly good agreement, which allowed validating the employed design approach. Finally an optimized injector design is proposed.

Combustion

CERN Document Server

Glassman, Irvin

1997-01-01

This Third Edition of Glassman's classic text clearly defines the role of chemistry, physics, and fluid mechanics as applied to the complex topic of combustion. Glassman's insightful introductory text emphasizes underlying physical and chemical principles, and encompasses engine technology, fire safety, materials synthesis, detonation phenomena, hydrocarbon fuel oxidation mechanisms, and environmental considerations. Combustion has been rewritten to integrate the text, figures, and appendixes, detailing available combustion codes, making it not only an excellent introductory text but also an important reference source for professionals in the field. Key Features * Explains complex combustion phenomena with physical insight rather than extensive mathematics * Clarifies postulates in the text using extensive computational results in figures * Lists modern combustion programs indicating usage and availability * Relates combustion concepts to practical applications.

On use of hybrid rocket propulsion for suborbital vehicles

Science.gov (United States)

Okninski, Adam

2018-04-01

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

Rocket center Peenemünde — Personal memories

Science.gov (United States)

Dannenberg, Konrad; Stuhlinger, Ernst

Von Braun built his first rockets as a young teenager. At 14, he started making plans for rockets for human travel to the Moon and Mars. The German Army began a rocket program in 1929. Two years later, Colonel (later General) Becker contacted von Braun who experimented with rockets in Berlin, gave him a contract in 1932, and, jointly with the Air Force, in 1936 built the rocket center Peenemünde where von Braun and his team developed the A-4 (V-2) rocket under Army auspices, while the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes. Albert Speer, impressed by the work of the rocketeers, allowed a modest growth of the Peenemünde project; this brought Dannenberg to the von Braun team in 1940. Hitler did not believe in rockets; he ignored the A-4 project until 1942 when he began to support it, expecting that it could turn the fortunes of war for him. He drastically increased the Peenemünde work force and allowed the transfer of soldiers from the front to Peenemünde; that was when Stuhlinger, in 1943, came to Peenemünde as a Pfc.-Ph.D. Later that year, Himmler wrenched the authority over A-4 production out of the Army's hands, put it under his command, and forced production of the immature rocket at Mittelwerk, and its military deployment against targets in France, Belgium, and England. Throughout the development of the A-4 rocket, von Braun was the undisputed leader of the project. Although still immature by the end of the war, the A-4 had proceeded to a status which made it the first successful long-range precision rocket, the prototype for a large number of military rockets built by numerous nations after the war, and for space rockets that launched satellites and traveled to the Moon and the planets.

Current status of rocket developments in universities -development of a small hybrid rocket with a swirling oxidizer flow type engine

OpenAIRE

Yuasa, Saburo; Kitagawa, Koki

2005-01-01

To develop an experimental small hybrid rocket with a swirling gaseous oxygen flow type engine, we made a flight model engine. Burning tests of the engine showed that a maximum thrust of 692 N and a specific impulse of 263 s (at sea level) were achieved. We designed a small hybrid rocket with this engine. The rocket measured 1.8 m in length and 15.4 kg in mass. To confirm the flight stability of the rocket, wind tunnel tests using a 112-scale model of the rocket and simulations of the flight ...

Photometric observations of local rocket-atmosphere interactions

Science.gov (United States)

Greer, R. G. H.; Murtagh, D. P.; Witt, G.; Stegman, J.

1983-06-01

Photometric measurements from rocket flights which recorded a strong foreign luminance in the altitude region between 90 and 130 km are reported. From one Nike-Orion rocket the luminance appeared on both up-leg and down-leg; from a series of Petrel rockets the luminance was apparent only on the down-leg. The data suggest that the luminance may be distributed mainly in the wake region along the rocket trajectory. The luminance is believed to be due to a local interaction between the rocket and the atmosphere although the precise nature of the interaction is unknown. It was measured at wavelengths ranging from 275 nm to 1.61 microns and may be caused by a combination of reactions.

Rocket Flight Path

Directory of Open Access Journals (Sweden)

Jamie Waters

2014-09-01

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

Combustion engineering

CERN Document Server

Ragland, Kenneth W

2011-01-01

Introduction to Combustion Engineering The Nature of Combustion Combustion Emissions Global Climate Change Sustainability World Energy Production Structure of the Book   Section I: Basic Concepts Fuels Gaseous Fuels Liquid Fuels Solid Fuels Problems Thermodynamics of Combustion Review of First Law Concepts Properties of Mixtures Combustion StoichiometryChemical EnergyChemical EquilibriumAdiabatic Flame TemperatureChemical Kinetics of CombustionElementary ReactionsChain ReactionsGlobal ReactionsNitric Oxide KineticsReactions at a Solid SurfaceProblemsReferences  Section II: Combustion of Gaseous and Vaporized FuelsFlamesLaminar Premixed FlamesLaminar Flame TheoryTurbulent Premixed FlamesExplosion LimitsDiffusion FlamesGas-Fired Furnaces and BoilersEnergy Balance and EfficiencyFuel SubstitutionResidential Gas BurnersIndustrial Gas BurnersUtility Gas BurnersLow Swirl Gas BurnersPremixed-Charge Engine CombustionIntroduction to the Spark Ignition EngineEngine EfficiencyOne-Zone Model of Combustion in a Piston-...

EMISSION AND COMBUSTION CHARACTERISTICS OF DIFFERENT FUELS IN A HCCI ENGINE

Directory of Open Access Journals (Sweden)

S. Sendilvelan

2011-06-01

Full Text Available Different intake valve timings and fuel injection amounts were tested in order to identify their effects on exhaust emissions and combustion characteristics using variable valve actuation (VVA in a Homogeneous Charge Compression Ignition (HCCI engine. The HCCI engine is a promising concept for future automobile engines and stationary power plants. The two-stage ignition process in a HCCI engine creates advanced ignition and stratified combustion, which makes the ignition timing and combustion rate controllable. Meanwhile, the periphery of the fuel-rich zone leads to fierce burning, which results in slightly high NOx emissions. The experiments were conducted in a modified single cylinder water-cooled diesel engine. In this experiment we use diesel, bio-diesel (Jatropha and gasoline as the fuel at different mixing ratios. HCCI has advantages in high thermal efficiency and low emissions and could possibly become a promising combustion method in internal combustion engines.

Wake effect in rocket observation

International Nuclear Information System (INIS)

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

1975-01-01

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

Hot Firing of a Full Scale Copper Tubular Combustion Chamber

National Research Council Canada - National Science Library

Cooley, C

2002-01-01

This paper describes the chamber design and hot firing test results for a full-scale copper tubular combustion chamber that has future application in a high-thrust, upper-stage expander cycle engine...

Nitrous Oxide/Paraffin Hybrid Rocket Engines

Science.gov (United States)

Zubrin, Robert; Snyder, Gary

2010-01-01

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

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

Science.gov (United States)

Marshall, Joel H.

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

A Flight Demonstration of Plasma Rocket Propulsion

Science.gov (United States)

Petro, Andrew

1999-01-01

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

Species- and stage-specific differences in trace element tissue concentrations in amphibians: implications for the disposal of coal-combustion wastes

Energy Technology Data Exchange (ETDEWEB)

Roe, J.H.; Hopkins, W.A.; Jackson, B.P. [University of Georgia, Aiken, SC (US)

2005-07-01

Information on species- and stage-specific patterns of contaminant accumulation is generally lacking for amphibians, yet such information could provide valuable knowledge on how amphibians interact with contaminants. We assessed concentrations of As, Cd, Cu, Ni, Pb, Se, Sr, and Zn in whole bodies of larval, recently metamorphosed, and adult life stages in Bufo terrestris and Rana sphenocephala from a site that currently receives coal combustion waste (CCW) discharge, a site where CCW was formerly discharged that has undergone natural attenuation for 30 years, and a nearby reference site. For the majority of elements (As, Cd, Cu, Ni, Pb, Zn), concentrations were highest in larvae, but Se and Sr concentrations remained elevated in later life stages, likely because these elements are S and Ca analogs, respectively, and are thus retained throughout structural changes during metamorphosis. Element concentrations were generally higher in B. terrestris than in R. sphenocephala. Concentrations of As, Se, and Sr were up to 11-35 times higher in metamorphs emigrating from CCW-polluted wetlands compared to unpolluted wetlands, suggesting metamorphosed amphibians can transport trace elements from aquatic disposal basins to nearby uncontaminated terrestrial habitats. In addition, anurans utilizing naturally revegetated sites up to 30 years after CCW disposal ceases are exposed to trace elements, although to a lesser degree than sites where CCW is currently discharged.

Numerical study on NO formation in a pulverized coal-fired furnace using oxy-fuel combustion

Energy Technology Data Exchange (ETDEWEB)

Zhang, Aiyue; Chen, Yuan; Sheng, Changdong [Southeast Univ., Nanjing (China). School of Energy and Environment

2013-07-01

Computational fluid dynamics (CFD) approach was employed to numerically investigate NO formation in a 600 MW wall-fired pulverized coal-fired furnace retrofitted for oxy-coal combustion, aimed at the impacts of flue gas recycle ratio, O{sub 2} staging and recycled NO with the recycled flue gas (RFG) on NO formation and emission. An in-house CFD research code for conventional air combustion was developed and extended to simulate O{sub 2}/RFG combustion with specific considerations of the change of gas properties and its impact on coal particle combustion processes. The extended De Soete mechanisms including NO reburning mechanism were applied to describe transformations of fuel nitrogen. It was shown that CFD simulation represented the significant reduction of NO formation during O{sub 2}/RFG combustion compared to that during air combustion. The in-burner and particularly the in-furnace O{sub 2} staging were confirmed still to play very important roles in NO formation control. Changing the recycle ratio had significant impact on the combustion performance and consequently on NO formation and emission. With the combustion performance ensured, decreasing the flue gas recycle ratio or increasing the inlet O{sub 2} concentration of combustion gas led to reduction of NO formation and emission. Although NO formation and emission was found to increase with increasing the inlet NO concentration of combustion gas, CFD simulation indicated that {proportional_to}74% of the inlet NO was reduced in the furnace, consistent with the experimental data reported in the literature. This demonstrated the significant contribution of reburning mechanism to the reduction of the recycled NO in the furnace.

Nitrogen Chemistry During Burnout in Fuel-Staged Combustion

DEFF Research Database (Denmark)

Kristensen, Per Gravers; Glarborg, Peter; Dam-Johansen, Kim

1996-01-01

A parametric study involving flow reactor experiments and chemical kinetic modeling is presented for the burnout zone in fuel-staging (reburning). The results provide guidelines for optimizing the reburn process and provide a test basis for verifying kinetic models for nitrogen chemistry at tempe......A parametric study involving flow reactor experiments and chemical kinetic modeling is presented for the burnout zone in fuel-staging (reburning). The results provide guidelines for optimizing the reburn process and provide a test basis for verifying kinetic models for nitrogen chemistry...

Modeling of Supersonic Combustion Systems for Sustained Hypersonic Flight

Directory of Open Access Journals (Sweden)

Stephen M. Neill

2017-11-01

Full Text Available Through Computational Fluid Dynamics and validation, an optimal scramjet combustor has been designed based on twin-strut Hydrogen injection to sustain flight at a desired speed of Mach 8. An investigation undertaken into the efficacy of supersonic combustion through various means of injection saw promising results for Hydrogen-based systems, whereby strut-style injectors were selected over transverse injectors based on their pressure recovery performance and combustive efficiency. The final configuration of twin-strut injectors provided robust combustion and a stable region of net thrust (1873 kN in the nozzle. Using fixed combustor inlet parameters and injection equivalence ratio, the finalized injection method advanced to the early stages of two-dimensional (2-D and three-dimensional (3-D scramjet engine integration. The overall investigation provided a feasible supersonic combustion system, such that Mach 8 sustained cruise could be achieved by the aircraft concept in a computational design domain.

The combustion behavior of large scale lithium titanate battery

Science.gov (United States)

Huang, Peifeng; Wang, Qingsong; Li, Ke; Ping, Ping; Sun, Jinhua

2015-01-01

Safety problem is always a big obstacle for lithium battery marching to large scale application. However, the knowledge on the battery combustion behavior is limited. To investigate the combustion behavior of large scale lithium battery, three 50 Ah Li(NixCoyMnz)O2/Li4Ti5O12 batteries under different state of charge (SOC) were heated to fire. The flame size variation is depicted to analyze the combustion behavior directly. The mass loss rate, temperature and heat release rate are used to analyze the combustion behavior in reaction way deeply. Based on the phenomenon, the combustion process is divided into three basic stages, even more complicated at higher SOC with sudden smoke flow ejected. The reason is that a phase change occurs in Li(NixCoyMnz)O2 material from layer structure to spinel structure. The critical temperatures of ignition are at 112–121°C on anode tab and 139 to 147°C on upper surface for all cells. But the heating time and combustion time become shorter with the ascending of SOC. The results indicate that the battery fire hazard increases with the SOC. It is analyzed that the internal short and the Li+ distribution are the main causes that lead to the difference. PMID:25586064

Cycle-to-cycle fluctuation of combustion in a spark-ignition engine; Hibana tenka engine no nensho hendo

Energy Technology Data Exchange (ETDEWEB)

Hamamoto, Y; Yoshiyama, S; Tomita, E; Hamagami, T [Okayama University, Okayama (Japan); Otsubo, H [Yammer Diesel Engine Co. Ltd. Tokyo (Japan)

1997-10-01

In a homogeneous charge spark-ignition engine, the duration of early stage of combustion is a dominant factor for determining the fluctuation of mean effective pressure. And the early stage of combustion varies with the equivalence ratio and turbulence characteristics of the mixture. In this study, the fluctuations of 1% combustion duration and indicated mean effective pressure Pmi were computed as the function of fluctuations both in the equivalence ratio {phi} of the mixture and in the turbulence characteristics of the cylinder charge. And effects of the spark timing {theta}ig and {phi} on the cycle-to-cycle fluctuation in Pmi were investigated. 16 refs., 6 figs.

Propellant injection strategy for suppressing acoustic combustion instability

Science.gov (United States)

Diao, Qina

Shear-coaxial injector elements are often used in liquid-propellant-rocket thrust chambers, where combustion instabilities remain a significant problem. A conventional solution to the combustion instability problem relies on passive control techniques that use empirically-developed hardware such as acoustic baffles and tuned cavities. In addition to adding weight and decreasing engine performance, these devices are designed using trial-and-error methods, which do not provide the capability to predict the overall system stability characteristics in advance. In this thesis, two novel control strategies that are based on propellant fluid dynamics were investigated for mitigating acoustic instability involving shear-coaxial injector elements. The new control strategies would use a set of controlled injectors allowing local adjustment of propellant flow patterns for each operating condition, particularly when instability could become a problem. One strategy relies on reducing the oxidizer-fuel density gradient by blending heavier methane with the main fuel, hydrogen. Another strategy utilizes modifying the equivalence ratio to affect the acoustic impedance through mixing and reaction rate changes. The potential effectiveness of these strategies was assessed by conducting unit-physics experiments. Two different model combustors, one simulating a single-element injector test and the other a double-element injector test, were designed and tested for flame-acoustic interaction. For these experiments, the Reynolds number of the central oxygen jet was kept between 4700 and 5500 making the injector flames sufficiently turbulent. A compression driver, mounted on one side of the combustor wall, provided controlled acoustic excitation to the injector flames, simulating the initial phase of flame-acoustic interaction. Acoustic excitation was applied either as band-limited white noise forcing between 100 Hz and 5000 Hz or as single-frequency, fixed-amplitude forcing at 1150 Hz

Large eddy simulation of the low temperature ignition and combustion processes on spray flame with the linear eddy model

Science.gov (United States)

Wei, Haiqiao; Zhao, Wanhui; Zhou, Lei; Chen, Ceyuan; Shu, Gequn

2018-03-01

Large eddy simulation coupled with the linear eddy model (LEM) is employed for the simulation of n-heptane spray flames to investigate the low temperature ignition and combustion process in a constant-volume combustion vessel under diesel-engine relevant conditions. Parametric studies are performed to give a comprehensive understanding of the ignition processes. The non-reacting case is firstly carried out to validate the present model by comparing the predicted results with the experimental data from the Engine Combustion Network (ECN). Good agreements are observed in terms of liquid and vapour penetration length, as well as the mixture fraction distributions at different times and different axial locations. For the reacting cases, the flame index was introduced to distinguish between the premixed and non-premixed combustion. A reaction region (RR) parameter is used to investigate the ignition and combustion characteristics, and to distinguish the different combustion stages. Results show that the two-stage combustion process can be identified in spray flames, and different ignition positions in the mixture fraction versus RR space are well described at low and high initial ambient temperatures. At an initial condition of 850 K, the first-stage ignition is initiated at the fuel-lean region, followed by the reactions in fuel-rich regions. Then high-temperature reaction occurs mainly at the places with mixture concentration around stoichiometric mixture fraction. While at an initial temperature of 1000 K, the first-stage ignition occurs at the fuel-rich region first, then it moves towards fuel-richer region. Afterwards, the high-temperature reactions move back to the stoichiometric mixture fraction region. For all of the initial temperatures considered, high-temperature ignition kernels are initiated at the regions richer than stoichiometric mixture fraction. By increasing the initial ambient temperature, the high-temperature ignition kernels move towards richer

Gas-phase reactions at combustion and gasification

International Nuclear Information System (INIS)

Hupa, M.; Kilpinen, P.; Chowdhury, K.; Brink, A.; Mueller, C.

1995-01-01

Formation and destruction of gaseous nitrogen pollutants at combustion (NO x , N 2 O) and gasification (NH 3 , HCN) are studied based on detailed chemical kinetic modelling and experiments in laboratory reactors. During 1994 the following topics have been studied: (a) nitrogen reactions in pressurized combustion processes (in co-operation with the LIEKKI projects 202 and 204), (b) NO x reduction by staging techniques in CO 2 , rich combustion processes, (c) HCN reactions at pyrolysis, (d) formation of soot precursors in a blast furnace (in co-operation with the SULA project 103) (e) incorporation of better NO x description into furnace models, (in co-operation with the LIEKKI project Y01). NH 3 conversion to N 2 in gasification product gases, (in co-operation with the LIEKKI project 203). In this report, some results of the items (a-c) will be presented. The results of items (d-f) are described in the reports by the co-operation projects. (author)

EMISSION AND COMBUSTION CHARACTERISTICS OF DIFFERENT FUELS IN A HCCI ENGINE

OpenAIRE

S. Sendilvelan; S.Mohanamurugan

2011-01-01

Different intake valve timings and fuel injection amounts were tested in order to identify their effects on exhaust emissions and combustion characteristics using variable valve actuation (VVA) in a Homogeneous Charge Compression Ignition (HCCI) engine. The HCCI engine is a promising concept for future automobile engines and stationary power plants. The two-stage ignition process in a HCCI engine creates advanced ignition and stratified combustion, which makes the ignition timing and combus...

Combustion

CERN Document Server

Glassman, Irvin

1987-01-01

Combustion, Second Edition focuses on the underlying principles of combustion and covers topics ranging from chemical thermodynamics and flame temperatures to chemical kinetics, detonation, ignition, and oxidation characteristics of fuels. Diffusion flames, flame phenomena in premixed combustible gases, and combustion of nonvolatile fuels are also discussed. This book consists of nine chapters and begins by introducing the reader to heats of reaction and formation, free energy and the equilibrium constants, and flame temperature calculations. The next chapter explores the rates of reactio

Ignition and combustion characteristics of metallized propellants, phase 2

Science.gov (United States)

Mueller, D. C.; Turns, S. R.

1994-01-01

Experimental and analytical investigations focusing on aluminum/hydrocarbon gel droplet secondary atomization and its effects on gel-fueled rocket engine performance are being conducted. A single laser sheet sizing/velocimetry diagnostic technique, which should eliminate sizing bias in the data collection process, has been designed and constructed to overcome limitations of the two-color forward-scatter technique used in previous work. Calibration of this system is in progress and the data acquisition/validation code is being written. Narrow-band measurements of radiant emission, discussed in previous reports, will be used to determine if aluminum ignition has occurred in a gel droplet. A one-dimensional model of a gel-fueled rocket combustion chamber, described in earlier reports, has been exercised in conjunction with a two-dimensional, two-phase nozzle code to predict the performance of an aluminum/hydrocarbon fueled engine. Estimated secondary atomization effects on propellant burnout distance, condensed particle radiation losses to the chamber walls, and nozzle two phase flow losses are also investigated. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size, and radiation heat losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated, depending on secondary atomization intensity. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two phase flow effects on overall engine performance. Radiation losses yielded a one percent decrease in engine Isp. Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine Isp was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass

Optical Signature Analysis of Tumbling Rocket Bodies via Laboratory Measurements

Science.gov (United States)

Cowardin, H.; Lederer, S.; Liou, J.-C.

2012-01-01

The NASA Orbital Debris Program Office has acquired telescopic lightcurve data on massive intact objects, specifically spent rocket bodies, in order to ascertain tumble rates in support of the Active Debris Removal (ADR) task to help remediate the LEO environment. Rotation rates are needed to plan and develop proximity operations for potential future ADR operations. To better characterize and model optical data acquired from ground-based telescopes, the Optical Measurements Center (OMC) at NASA/JSC emulates illumination conditions in space using equipment and techniques that parallel telescopic observations and source-target-sensor orientations. The OMC employs a 75-watt Xenon arc lamp as a solar simulator, an SBIG CCD camera with standard Johnson/Bessel filters, and a robotic arm to simulate an object's position and rotation. The light source is mounted on a rotary arm, allowing access any phase angle between 0 -- 360 degrees. The OMC does not attempt to replicate the rotation rates, but focuses on how an object is rotating as seen from multiple phase angles. The two targets studied are scaled (1:48), SL-8 Cosmos 3M second stages. The first target is painted in the standard government "gray" scheme and the second target is primary white, as used for commercial missions. This paper summarizes results of the two scaled rocket bodies, each rotated about two primary axes: (a) a spin-stabilized rotation and (b) an end-over-end rotation. The two rotation states are being investigated as a basis for possible spin states of rocket bodies, beginning with simple spin states about the two primary axes. The data will be used to create a database of potential spin states for future works to convolve with more complex spin states. The optical signatures will be presented for specific phase angles for each rocket body and shown in conjunction with acquired optical data from multiple telescope sources.

Rocket Ozone Data Recovery for Digital Archival

Science.gov (United States)

Hwang, S. H.; Krueger, A. J.; Hilsenrath, E.; Haffner, D. P.; Bhartia, P. K.

2014-12-01

Ozone distributions in the photochemically-controlled upper stratosphere and mesosphere were first measured using spectrometers on V-2 rockets after WWII. The IGY(1957-1958) spurred development of new optical and chemical instruments for flight on meteorological and sounding rockets. In the early 1960's, the US Navy developed an Arcas rocket-borne optical ozonesonde and NASA GSFC developed chemiluminescent ozonesonde onboard Nike_Cajun and Arcas rocket. The Navy optical ozone program was moved in 1969 to GSFC where rocket ozone research was expanded and continued until 1994 using Super Loki-Dart rocket at 11 sites in the range of 0-65N and 35W-160W. Over 300 optical ozone soundings and 40 chemiluminescent soundings were made. The data have been used to produce the US Standard Ozone Atmosphere, determine seasonal and diurnal variations, and validate early photochemical models. The current effort includes soundings conducted by Australia, Japan, and Korea using optical techniques. New satellite ozone sounding techniques were initially calibrated and later validated using the rocket ozone data. As satellite techniques superseded the rocket methods, the sponsoring agencies lost interest in the data and many of those records have been discarded. The current task intends to recover as much of the data as possible from the private records of the experimenters and their publications, and to archive those records in the WOUDC (World Ozone and Ultraviolet Data Centre). The original data records are handwritten tabulations, computer printouts that are scanned with OCR techniques, and plots digitized from publications. This newly recovered digital rocket ozone profile data from 1965 to 2002 could make significant contributions to the Earth science community in atmospheric research including long-term trend analysis.

Scientific Experiences Using Argentinean Sounding Rockets in Antarctica

Science.gov (United States)

Sánchez-Peña, Miguel

2000-07-01

Argentina in the sixties and seventies, had experience for developing and for using sounding rockets and payloads to perform scientific space experiments. Besides they have several bases in Antarctica with adequate premises and installations, also duly equipped aircrafts and trained crews to flight to the white continent. In February 1965, scientists and technical people from the "Instituto de Investigacion Aeronáutica y Espacial" (I.I.A.E.) with the cooperation of the Air Force and the Tucuman University, conducted the "Matienzo Operation" to measure X radiation and temperature in the upper atmosphere, using the Gamma Centauro rocket and also using big balloons. The people involved in the experience, the launcher, other material and equipment flew from the south tip of Argentina to the Matienzo base in Antarctica, in a C-47 aircraft equipped with skies an additional jet engine Marbore 2-C. Other experience was performed in 1975 in the "Marambio" Antartic Base, using the two stages solid propellent sounding rocket Castor, developed in Argentina. The payload was developed in cooperation with the Max Planck Institute of Germany. It consist of a special mixture including a shape charge to form a ionized cloud producing a jet of electrons travelling from Marambio base to the conjugate point in the Northern hemisphere. The cloud was observed by several ground stations in Argentina and also by a NASA aircraft with TV cameras, flying at East of New York. The objective of this experience was to study the electric and magnetic fields in altitude, the neutral points, the temperature and electrons profile. The objectives of both experiments were accomplished satisfactorily.

Study of the combustion of aluminium and magnesium particulates: influence of the composition of the gaseous mixture and of pressure; Etude de la combustion de particules d'aluminium et de magnesium: influence de la composition du melange gazeux et de la pression

Energy Technology Data Exchange (ETDEWEB)

Legrand, B.

2000-07-01

The combustion of metal particulates has a major interest in the domain of space propulsion. Aluminium is today used as doping material in the solid propellant of Ariane 5 rocket engines. Magnesium represents a possible fuel for propellers allowing a come back from Mars. An electrostatic levitation device has been used to study the combustion in controlled environment of particulates having a size representative of those encountered in propellers. The particulates are ignited with a laser and observed by fast cinematography. The inhibitive property of hydrogen chloride, an important constituent of the propellant atmosphere, on the combustion of aluminium particulates has been evidenced. These results have been compared with those obtained with a kinetic model in gaseous phase. The combustion of magnesium particulates in carbon dioxide has been studied for 53-63 {mu}m and 1-2 mm particulates. It is shown that the ignition of small particulates is controlled by the chemical kinetics and that the limit ignition pressure is reversely proportional to the particulates size. A study on big samples, performed in normal gravity but also in reduced gravity to get rid of the natural convection phenomena, has permitted to show a pulsed combustion regime linked with the presence of heterogenous reactions. The measurement of the combustion durations for the different sizes of particulates has permitted to propose a correlation between these two parameters for the particulate diameters comprised between 50 {mu}m and 2 mm. (J.S.)

Rocket in situ observation of equatorial plasma irregularities in the region between E and F layers over Brazil

Directory of Open Access Journals (Sweden)

S. Savio Odriozola

2017-03-01

Full Text Available A two-stage VS-30 Orion rocket was launched from the equatorial rocket launching station in Alcântara, Brazil, on 8 December 2012 soon after sunset (19:00 LT, carrying a Langmuir probe operating alternately in swept and constant bias modes. At the time of launch, ground equipment operated at equatorial stations showed rapid rise in the base of the F layer, indicating the pre-reversal enhancement of the F region vertical drift and creating ionospheric conditions favorable for the generation of plasma bubbles. Vertical profiles of electron density estimated from Langmuir probe data showed wave patterns and small- and medium-scale plasma irregularities in the valley region (100–300 km during the rocket upleg and downleg. These irregularities resemble those detected by the very high frequency (VHF radar installed at Jicamarca and so-called equatorial quasi-periodic echoes. We present evidence suggesting that these observations could be the first detection of this type of irregularity made by instruments onboard a rocket.

Direct electrical arc ignition of hybrid rocket motors

Science.gov (United States)

Judson, Michael I., Jr.

Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

Numerical simulations of a sounding rocket in ionospheric plasma: Effects of magnetic field on the wake formation and rocket potential

Science.gov (United States)

Darian, D.; Marholm, S.; Paulsson, J. J. P.; Miyake, Y.; Usui, H.; Mortensen, M.; Miloch, W. J.

2017-09-01

The charging of a sounding rocket in subsonic and supersonic plasma flows with external magnetic field is studied with numerical particle-in-cell (PIC) simulations. A weakly magnetized plasma regime is considered that corresponds to the ionospheric F2 layer, with electrons being strongly magnetized, while the magnetization of ions is weak. It is demonstrated that the magnetic field orientation influences the floating potential of the rocket and that with increasing angle between the rocket axis and the magnetic field direction the rocket potential becomes less negative. External magnetic field gives rise to asymmetric wake downstream of the rocket. The simulated wake in the potential and density may extend as far as 30 electron Debye lengths; thus, it is important to account for these plasma perturbations when analyzing in situ measurements. A qualitative agreement between simulation results and the actual measurements with a sounding rocket is also shown.

Pulsating combustion - Combustion characteristics and reduction of emissions

Energy Technology Data Exchange (ETDEWEB)

Lindholm, Annika

1999-11-01

In the search for high efficiency combustion systems pulsating combustion has been identified as one of the technologies that potentially can meet the objectives of clean combustion and good fuel economy. Pulsating combustion offers low emissions of pollutants, high heat transfer and efficient combustion. Although it is an old technology, the interest in pulsating combustion has been renewed in recent years, due to its unique features. Various applications of pulsating combustion can be found, mainly as drying and heating devices, of which the latter also have had commercial success. It is, however, in the design process of a pulse combustor, difficult to predict the operating frequency, the heat release etc., due to the lack of a well founded theory of the phenomenon. Research concerning control over the combustion process is essential for developing high efficiency pulse combustors with low emissions. Natural gas fired Helmholtz type pulse combustors have been the experimental objects of this study. In order to investigate the interaction between the fluid dynamics and the chemistry in pulse combustors, laser based measuring techniques as well as other conventional measuring techniques have been used. The experimental results shows the possibilities to control the combustion characteristics of pulsating combustion. It is shown that the time scales in the large vortices created at the inlet to the combustion chamber are very important for the operation of the pulse combustor. By increasing/decreasing the time scale for the large scale mixing the timing of the heat release is changed and the operating characteristics of the pulse combustor changes. Three different means for NO{sub x} reduction in Helmholtz type pulse combustors have been investigated. These include exhaust gas recirculation, alteration of air/fuel ratio and changed inlet geometry in the combustion chamber. All used methods achieved less than 10 ppm NO{sub x} emitted (referred to stoichiometric

Experimental studies of thermal preparation of internal combustion engine

Science.gov (United States)

Karnaukhov, N. N.; Merdanov, Sh M.; V, Konev V.; Borodin, D. M.

2018-05-01

In conditions of autonomous functioning of road construction machines, it becomes necessary to use its internal sources. This can be done by using a heat recovery system of an internal combustion engine (ICE). For this purpose, it is proposed to use heat accumulators that accumulate heat of the internal combustion engine during the operation of the machine. Experimental studies have been carried out to evaluate the efficiency of using the proposed pre-start thermal preparation system, which combines a regular system based on liquid diesel fuel heaters and an ICE heat recovery system. As a result, the stages of operation of the preheating thermal preparation system, mathematical models and the dependence of the temperature change of the antifreeze at the exit from the internal combustion engine on the warm-up time are determined.

Chinese modify CZ-2/3 rocket boosters, focus on commercial launch market

Science.gov (United States)

Covault, C.

1985-07-01

A program underway in the People's Republic of China to modify the Titan-class CZ-2/3 satellite-launch and ICBM boosters is described on the basis of a recent visit to the manufacturing plant in Shanghai. The present two-stage CZ-2 and three-stage CZ-3 can place 5000 lbs in LEO or 3080 lbs in GEO, respectively, and are produced on a custom basis with a delivery time of about 2 yrs. Modifications introduced include 4 x 6-ft fins and a pogo-suppression system for the four-engine first stage and a steel support band for the combustion chamber of the 80-ton-thrust second-stage main engine.

Smog chamber study on the evolution of fume from residential coal combustion.

Science.gov (United States)

Geng, Chunmei; Wang, Kun; Wang, Wei; Chen, Jianhua; Liu, Xiaoyu; Liu, Hongjie

2012-01-01

Domestic coal stoves are widely used in countryside and greenbelt residents in China for heating and cooking, and emit considerable pollutants to the atmosphere because of no treatment of their exhaust, which can result in deteriorating local air quality. In this study, a dynamic smog chamber was used to investigate the real-time emissions of gaseous and particulate pollutants during the combustion process and a static smog chamber was used to investigate the fume evolution under simulate light irradiation. The real-time emissions revealed that the total hydrocarbon (THC) and CO increased sharply after ignition, and then quickly decreased, indicating volatilization of hydrocarbons with low molecular weight and incomplete combustion at the beginning stage of combustion made great contribution to these pollutants. There was evident shoulder peak around 10 min combustion for both THC and CO, revealing the emissions from vitrinite combustion. Additionally, another broad emission peak of CO after 30 min was also observed, which was ascribed to the incomplete combustion of the inertinite. Compared with THC and CO, there was only one emission peak for NOx, SO2 and particular matters at the beginning stage of combustion. The fume evolution with static chamber simulation indicated that evident consumption of SO2 and NOx as well as new particle formation were observed. The consumption rates for SO2 and NOx were about 3.44% hr(-1) and 3.68% hr(-1), the new particle formation of nuclei particles grew at a rate of 16.03 nm/hr during the first reaction hour, and the increase of the diameter of accumulation mode particles was evident. The addition of isoprene to the diluted mixture of the fume could promote 03 and secondary particle formation.

Investigation on the transient enthalpy of coal combustion

Energy Technology Data Exchange (ETDEWEB)

Fu, Pei-fang; Wang, Na; Yu, Bo; Zhang, Bin; Liu, Yang; Zhou, Huai-chun [Huazhong Univ. of Science and Technology, Wuhan (China). State Key Lab. of Coal Combustion

2013-07-01

The transient enthalpy ({Delta}h) of coal/char combustion of the three different coals (including anthracite, bituminous, and lignite) during the process of combustion is determined as a function of burn-off degree by using thermo-gravimetric-differential scanning calorimeter (TG-DSC) simultaneous thermal analyzer, and The error of determining calorific values of coals/chars is less 5% compared the results of TG-DSC with that of an automatic isoperibol calorimeter. It is found that In the initial stage, all the {Delta}h of coals are greater than that of the char pyrolysized from parent coal for many of volatiles contained more a great deal of heat per unit mass oxidized at low temperature, it also imply that coal is more easily ignited than char corresponded; And in the middle stage, all the {Delta}h of coals is lower than that of the char pyrolysized, so the pyrolysized char oxidation can supply much more of thermo-energy per unit mass. {Delta}h are almost a constant when the burn-off degree is equal to between 0.35/0.15 and 0.95/0.85 for ZCY bituminous coal/char and JWY anthracite/char, between 0.35/0.35 and 0.75/0.9 for SLH lignite/char; In the later stage, the {Delta}h of the coal/char decreased with the burn-off degree, it imply that the activity of the coal/char decreases. Therefore, coal pyrolysis changes not only the structure of char, but also the property of release heat; the transient enthalpy of coal/char combustion has been in change with the burn-out degree.

Nuclear thermal rocket propulsion application to Mars missions

International Nuclear Information System (INIS)

Emrich, W.J. Jr.; Young, A.C.; Mulqueen, J.A.

1991-01-01

Options for vehicle configurations are reviewed in which nuclear thermal rocket (NTR) propulsion is used for a reference mission to Mars. The scenario assumes an opposition-class Mars transfer trajectory, a 435-day mission, and the use of a single nuclear engine with 75,000 lbs of thrust. Engine parameters are examined by calculating mission variables for a range of specific impulses and thrust/weight ratios. The reference mission is found to have optimal values of 925 s for the specific impulse and thrust/weight ratios of 4.0 and 0.06 for the engine and total stage ratios respectively. When the engine thrust/weight ratio is at least 4/1 the most critical engine parameter is engine specific impulse for reducing overall stage weight. In the context of this trans-Mars three-burn maneuver the NTR engine with an expander engine cycle is considered a more effective alternative than chemical/aerobrake and other propulsion options

Soot temperature and KL factor for biodiesel and diesel spray combustion in a constant volume combustion chamber

KAUST Repository

Zhang, Ji

2013-07-01

This paper presents measurements of the soot temperature and KL factor for biodiesel and diesel combustion in a constant volume chamber using a two-color technique. This technique uses a high-speed camera coupled with two narrowband filters (550. nm and 650. nm, 10. nm FWHM). After calibration, statistical analysis shows that the uncertainty of the two-color temperature is less than 5%, while it is about 50% for the KL factor. This technique is then applied to the spray combustion of biodiesel and diesel fuels under an ambient oxygen concentration of 21% and ambient temperatures of 800, 1000 and 1200. K. The heat release result shows higher energy utilization efficiency for biodiesel compared to diesel under all conditions; meanwhile, diesel shows a higher pressure increase due to its higher heating value. Biodiesel yields a lower temperature inside the flame area, a longer soot lift-off length, and a smaller soot area compared to diesel. Both the KL factor and the total soot with biodiesel are lower than with diesel throughout the entire combustion process, and this difference becomes larger as the ambient temperature decreases. Biodiesel shows approximately 50-100. K lower temperatures than diesel at the quasi-steady stage for 1000 and 1200. K ambient temperature, while diesel shows a lower temperature than biodiesel at 800. K ambient. This result may raise the question of how important the flame temperature is in explaining the higher NO. x emissions often observed during biodiesel combustion. Other factors may also play an important role in controlling NO. x emissions. Both biodiesel and diesel temperature measurements show a monotonic dependence on the ambient temperature. However, the ambient temperature appears to have a more significant effect on the soot formation and oxidation in diesel combustion, while biodiesel combustion soot characteristics shows relative insensitivity to the ambient temperature. © 2013 Elsevier Ltd.

Furnace devices aerodynamics optimization for fuel combustion efficiency improvement and nitrogen oxide emission reduction

Science.gov (United States)

Volkov, E. P.; Prokhorov, V. B.; Arkhipov, A. M.; Chernov, S. L.; Kirichkov, V. S.; Kaverin, A. A.

2017-11-01

MPEI conducts researches on physical and mathematical models of furnace chambers for improvement of power-generation equipment fuel combustion efficiency and ecological safety. Results of these researches are general principles of furnace aerodynamics arrangement for straight-flow burners and various fuels. It has been shown, that staged combustion arrangement with early heating and igniting with torch distribution in all furnace volume allows to obtain low carbon in fly ash and nitrogen oxide emission and also to improve boiler operation reliability with expand load adjustment range. For solid fuel combustion efficiency improvement it is practical to use high-placed and strongly down-tilted straight-flow burners, which increases high-temperature zone residence time for fuel particles. In some cases, for this combustion scheme it is possible to avoid slag-tap removal (STR) combustion and to use Dry-bottom ash removal (DBAR) combustion with tolerable carbon in fly ash level. It is worth noting that boilers with STR have very high nitrogen oxide emission levels (1200-1800 mg/m3) and narrow load adjustment range, which is determined by liquid slag output stability, so most industrially-developed countries don’t use this technology. Final decision about overhaul of boiler unit is made with regard to physical and mathematical modeling results for furnace and zonal thermal calculations for furnace and boiler as a whole. Overhaul of boilers to provide staged combustion and straight-flow burners and nozzles allows ensuring regulatory nitrogen oxide emission levels and corresponding best available technology criteria, which is especially relevant due to changes in Russian environmental regulation.

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

Science.gov (United States)

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

1987-01-01

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

Plasma waves observed by sounding rockets

International Nuclear Information System (INIS)

Kimura, I.

1977-01-01

Observations of plasma wave phenomena have been conducted with several rockets launched at Kagoshima Space Center, Kyushu, Japan, and at Showa Base, Antarctica. This report presents some results of the observations in anticipation of having valuable comments from other plasma physicists, especially from those who are concerned with laboratory plasma. In the K-9M-41 rocket experiment, VLF plasma waves were observed. In this experiment, the electron beam of several tens of uA was emitted from a hot cathode when a positive dc bias changing from 0 to 10V at 1V interval each second was applied to a receiving dipole antenna. The discrete emissions with 'U' shaped frequency spectrum were observed for the dc bias over 3 volts. The U emissions appeared twice per spin period of the rocket. Similar rocket experiment was performed at Showa Base using a loop and dipole antenna and without hot cathode. Emissions were observed with varying conditions. At present, the authors postulate that such emissions may be produced just in the vicinity of a rocket due to a kind of wake effect. (Aoki, K.)

Energetic Combustion Devices for Aerospace Propulsion and Power

Science.gov (United States)

Litchford, Ron J.

2000-01-01

Chemical reactions have long been the mainstay thermal energy source for aerospace propulsion and power. Although it is widely recognized that the intrinsic energy density limitations of chemical bonds place severe constraints on maximum realizable performance, it will likely be several years before systems based on high energy density nuclear fuels can be placed into routine service. In the mean time, efforts to develop high energy density chemicals and advanced combustion devices which can utilize such energetic fuels may yield worthwhile returns in overall system performance and cost. Current efforts in this vein are being carried out at NASA MSFC under the direction of the author in the areas of pulse detonation engine technology development and light metals combustion devices. Pulse detonation engines are touted as a low cost alternative to gas turbine engines and to conventional rocket engines, but actual performance and cost benefits have yet to be convincingly demonstrated. Light metal fueled engines also offer potential benefits in certain niche applications such as aluminum/CO2 fueled engines for endo-atmospheric Martian propulsion. Light metal fueled MHD generators also present promising opportunities with respect to electric power generation for electromagnetic launch assist. This presentation will discuss the applications potential of these concepts with respect to aero ace propulsion and power and will review the current status of the development efforts.

A review of active control approaches in stabilizing combustion systems in aerospace industry

Science.gov (United States)

Zhao, Dan; Lu, Zhengli; Zhao, He; Li, X. Y.; Wang, Bing; Liu, Peijin

2018-02-01

Self-sustained combustion instabilities are one of the most plaguing challenges and problems in lean-conditioned propulsion and land-based engine systems, such as rocket motors, gas turbines, industrial furnace and boilers, and turbo-jet thrust augmenters. Either passive or active control in open- or closed-loop configurations can be implemented to mitigate such instabilities. One of the classical disadvantages of passive control is that it is only implementable to a designed combustor over a limited frequency range and can not respond to the changes in operating conditions. Compared with passive control approaches, active control, especially in closed-loop configuration is more adaptive and has inherent capacity to be implemented in practice. The key components in closed-loop active control are 1) sensor, 2) controller (optimization algorithm) and 3) dynamic actuator. The present work is to outline the current status, technical challenges and development progress of the active control approaches (in open- or closed-loop configurations). A brief description of feedback control, adaptive control, model-based control and sliding mode control are provided first by introducing a simplified Rijke-type combustion system. The modelled combustion system provides an invaluable platform to evaluate the performance of these feedback controllers and a transient growth controller. The performance of these controllers are compared and discussed. An outline of theoretical, numerical and experimental investigations are then provided to overview the research and development progress made during the last 4 decades. Finally, potential, challenges and issues involved with the design, application and implementation of active combustion control strategies on a practical engine system are highlighted.

Species-and stage-specific differences in trace element tissue concentrations in amphibians: implications for the disposal of coal-combustion wastes

International Nuclear Information System (INIS)

Roe, John H.; Hopkins, William A.; Jackson, Brian P.

2005-01-01

Information on species-and stage-specific patterns of contaminant accumulation is generally lacking for amphibians, yet such information could provide valuable knowledge on how amphibians interact with contaminants. We assessed concentrations of As, Cd, Cu, Ni, Pb, Se, Sr, and Zn in whole bodies of larval, recently metamorphosed, and adult life stages in Bufo terrestris and Rana sphenocephala from a site that currently receives coal combustion waste (CCW) discharge, a site where CCW was formerly discharged that has undergone natural attenuation for 30 years, and a nearby reference site. For the majority of elements (As, Cd, Cu, Ni, Pb, Zn), concentrations were highest in larvae, but Se and Sr concentrations remained elevated in later life stages, likely because these elements are S and Ca analogs, respectively, and are thus retained throughout structural changes during metamorphosis. Element concentrations were generally higher in B. terrestris than in R. sphenocephala. Concentrations of As, Se, and Sr were up to 11-35 times higher in metamorphs emigrating from CCW-polluted wetlands compared to unpolluted wetlands, suggesting metamorphosed amphibians can transport trace elements from aquatic disposal basins to nearby uncontaminated terrestrial habitats. In addition, anurans utilizing naturally revegetated sites up to 30 years after CCW disposal ceases are exposed to trace elements, although to a lesser degree than sites where CCW is currently discharged. - Results suggest that metamorphosed amphibians can transport trace elements from aquatic disposal basins to non-contaminated habitats

Species-and stage-specific differences in trace element tissue concentrations in amphibians: implications for the disposal of coal-combustion wastes

Energy Technology Data Exchange (ETDEWEB)

Roe, John H. [University of Georgia, Savannah River Ecology Laboratory, P.O. Drawer E, Aiken, SC 29802 (United States); Hopkins, William A. [University of Georgia, Savannah River Ecology Laboratory, P.O. Drawer E, Aiken, SC 29802 (United States)]. E-mail: hopkins@srel.edu; Jackson, Brian P. [University of Georgia, Savannah River Ecology Laboratory, P.O. Drawer E, Aiken, SC 29802 (United States)

2005-07-15

Information on species-and stage-specific patterns of contaminant accumulation is generally lacking for amphibians, yet such information could provide valuable knowledge on how amphibians interact with contaminants. We assessed concentrations of As, Cd, Cu, Ni, Pb, Se, Sr, and Zn in whole bodies of larval, recently metamorphosed, and adult life stages in Bufo terrestris and Rana sphenocephala from a site that currently receives coal combustion waste (CCW) discharge, a site where CCW was formerly discharged that has undergone natural attenuation for 30 years, and a nearby reference site. For the majority of elements (As, Cd, Cu, Ni, Pb, Zn), concentrations were highest in larvae, but Se and Sr concentrations remained elevated in later life stages, likely because these elements are S and Ca analogs, respectively, and are thus retained throughout structural changes during metamorphosis. Element concentrations were generally higher in B. terrestris than in R. sphenocephala. Concentrations of As, Se, and Sr were up to 11-35 times higher in metamorphs emigrating from CCW-polluted wetlands compared to unpolluted wetlands, suggesting metamorphosed amphibians can transport trace elements from aquatic disposal basins to nearby uncontaminated terrestrial habitats. In addition, anurans utilizing naturally revegetated sites up to 30 years after CCW disposal ceases are exposed to trace elements, although to a lesser degree than sites where CCW is currently discharged. - Results suggest that metamorphosed amphibians can transport trace elements from aquatic disposal basins to non-contaminated habitats.

Combustible structural composites and methods of forming combustible structural composites

Science.gov (United States)

Daniels, Michael A.; Heaps, Ronald J.; Steffler, Eric D.; Swank, W. David

2013-04-02

Combustible structural composites and methods of forming same are disclosed. In an embodiment, a combustible structural composite includes combustible material comprising a fuel metal and a metal oxide. The fuel metal is present in the combustible material at a weight ratio from 1:9 to 1:1 of the fuel metal to the metal oxide. The fuel metal and the metal oxide are capable of exothermically reacting upon application of energy at or above a threshold value to support self-sustaining combustion of the combustible material within the combustible structural composite. Structural-reinforcing fibers are present in the composite at a weight ratio from 1:20 to 10:1 of the structural-reinforcing fibers to the combustible material. Other embodiments and aspects are disclosed.

Numerical investigation into premixed hydrogen combustion within two-stage porous media burner of 1 kW solid oxide fuel cell system

Energy Technology Data Exchange (ETDEWEB)

Yen Tzu-Hsiang; Chen Bao-Dong [Refining and Manufacturing Research Institute, CPC Corporation, Chia-Yi City 60036, Taiwan (China); Hong Wen-Tang; Tsai Yu-Ching; Wang Hung-Yu; Huang Cheng-Nan; Lee Chien-Hsiung [Institute of Nuclear Energy Research Atomic Energy Council, Taoyuan County 32546, Taiwan (China)

2010-07-01

Numerical simulations are performed to analyze the combustion of the anode off-gas / cathode off-gas mixture within the two-stage porous media burner of a 1 kW solid oxide fuel cell (SOFC) system. In performing the simulations, the anode gas is assumed to be hydrogen and the combustion of the gas mixture is modeled using a turbulent flow model. The validity of the numerical model is confirmed by comparing the simulation results for the flame barrier temperature and the porous media temperature with the corresponding experimental results. Simulations are then performed to investigate the effects of the hydrogen content and the burner geometry on the temperature distribution within the burner and the corresponding operational range. It is shown that the maximum flame temperature increases with an increasing hydrogen content. In addition, it is found that the burner has an operational range of 1.2--6.5 kW when assigned its default geometry settings (i.e. a length and diameter of 0.17 m and 0.06 m, respectively), but increases to 2--9 kW and 2.6--11.5 kW when the length and diameter are increased by a factor of 1.5, respectively. Finally, the operational range increases to 3.5--16.5 kW when both the diameter and the length of the burner are increased by a factor of 1.5.

Performance analysis of single stage libr-water absorption machine operated by waste thermal energy of internal combustion engine: Case study

Science.gov (United States)

Sharif, Hafiz Zafar; Leman, A. M.; Muthuraman, S.; Salleh, Mohd Najib Mohd; Zakaria, Supaat

2017-09-01

Combined heating, cooling, and power is also known as Tri-generation. Tri-generation system can provide power, hot water, space heating and air -conditioning from single source of energy. The objective of this study is to propose a method to evaluate the characteristic and performance of a single stage lithium bromide-water (LiBr-H2O) absorption machine operated with waste thermal energy of internal combustion engine which is integral part of trigeneration system. Correlations for computer sensitivity analysis are developed in data fit software for (P-T-X), (H-T-X), saturated liquid (water), saturated vapor, saturation pressure and crystallization temperature curve of LiBr-H2O Solution. Number of equations were developed with data fit software and exported into excel work sheet for the evaluation of number of parameter concerned with the performance of vapor absorption machine such as co-efficient of performance, concentration of solution, mass flow rate, size of heat exchangers of the unit in relation to the generator, condenser, absorber and evaporator temperatures. Size of vapor absorption machine within its crystallization limits for cooling and heating by waste energy recovered from exhaust gas, and jacket water of internal combustion engine also presented in this study to save the time and cost for the facilities managers who are interested to utilize the waste thermal energy of their buildings or premises for heating and air conditioning applications.

Additive Manufacturing a Liquid Hydrogen Rocket Engine

Science.gov (United States)

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

2016-01-01

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

The Spanish national programme of balloons and sounding rockets

International Nuclear Information System (INIS)

Casas, J.; Pueyo, L.

1978-01-01

The main points of the Spanish scientific programme are briefly described: CONIE/NASA cooperative project on meteorological sounding rocket launchings; ozonospheric programme; CONIE/NASA/CNES cooperative ionospheric sounding rocket project; D-layer research; rocket infrared dayglow measurements; ultraviolet astronomy research; cosmic ray research. The schedule of sounding rocket launchings at El Arenosillo station during 1977 is given

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

Science.gov (United States)

Evans, Brian

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

Mixture preparation and combustion in an optically-accessible HCCI, diesel engine; La preparation du melange et de la combustion dans un moteur Diesel, HCCI a acces optique

Energy Technology Data Exchange (ETDEWEB)

Kashdan, J.; Bruneaux, G. [Institut Francais du Petrole, 92 - Rueil-Malmaison (France)

2006-07-01

Planar laser-induced fluorescence (LIF) imaging techniques have been applied in order to study the mixture preparation and combustion process in a single cylinder, optically-accessible homogeneous charge, compression ignition (HCCI) engine. In particular, the influence of piston bowl geometry on the in-cylinder mixture distribution and subsequent combustion process has been investigated. A new optically-accessible piston design enabled the application of LIF diagnostics directly within the combustion chamber bowl. Firstly, laser-induced exciplex fluorescence (LIEF) was exploited in order to characterise the in-cylinder fuel spray and vapour distribution. Subsequently a detailed study of the two-stage HCCI combustion process was conducted by a combination of direct chemiluminescence imaging, laser-induced fluorescence (LIF) of the intermediate species formaldehyde (CH{sub 2}O) which is present during the cool flame and LIF of the OH radical which is subsequently present in the reaction and burned gas zones at higher temperature. Finally, spectrometry measurements were performed with the objective of determining the origin of the emitting species of the chemiluminescence signal. The experiments were performed on a single cylinder optical engine equipped with a direct-injection, common rail injection system and narrow angle injector. The experimental results presented reveal the significant role of the combustion chamber geometry on the mixture preparation and combustion characteristics for late HCCI injection strategies particularly in such cases where liquid impingement is unavoidable. Planar LIF 355 imaging revealed the presence of the intermediate species formaldehyde allowing the temporal and spatial detection of auto-ignition precursors prior to the signal observed by chemiluminescence in the early stages of the cool flame. Formaldehyde was then rapidly consumed at the start of main combustion which was marked not only by the increase in the main heat release

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Science.gov (United States)

Jones, Daniel R.; Allgood, Daniel C.; Saunders, Grady P.

2016-01-01

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Thermogravimetric investigation on the degradation properties and combustion performance of bio-oils.

Science.gov (United States)

Ren, Xueyong; Meng, Jiajia; Moore, Andrew M; Chang, Jianmin; Gou, Jinsheng; Park, Sunkyu

2014-01-01

The degradation properties and combustion performance of raw bio-oil, aged bio-oil, and bio-oil from torrefied wood were investigated through thermogravimetric analysis. A three-stage process was observed for the degradation of bio-oils, including devolatilization of the aqueous fraction and light compounds, transition of the heavy faction to solid, and combustion of carbonaceous residues. Pyrolysis kinetics parameters were calculated via the reaction order model and 3D-diffusion model, and combustion indexes were used to qualitatively evaluate the thermal profiles of tested bio-oils for comparison with commercial oils such as fuel oils. It was found that aged bio-oil was more thermally instable and produced more combustion-detrimental carbonaceous solid. Raw bio-oil and bio-oil from torrefied wood had comparable combustion performance to fuel oils. It was considered that bio-oil has a potential to be mixed with or totally replace the fuel oils in boilers. Copyright © 2013 Elsevier Ltd. All rights reserved.

A Mechanistic Investigation of Nitrogen Evolution and Corrosion with Oxy-Combustion

Energy Technology Data Exchange (ETDEWEB)

Dale Tree; Andrew Mackrory; Thomas Fletcher

2008-12-31

A premixed, staged, down-fired, pulverized coal reactor and a flat flame burner were used to study the evolution of nitrogen in coal contrasting differences in air and oxy-combustion. In the premixed reactor, the oxidizer was staged to produce a fuel rich zone followed by a burnout zone. The initial nominal fuel rich zone stoichiometric ratio (S.R.) of 0.85 selected produced higher NO reductions in the fuel rich region under oxy-combustion conditions. Air was found to be capable of similar NO reductions when the fuel rich zone was at a much lower S.R. of 0.65. At a S.R. of 0.85, oxy-combustion was measured to have higher CO, unburned hydrocarbons, HCN and NH{sub 3} in the fuel rich region than air at the same S.R. There was no measured difference in the initial formation of NO. The data suggest devolatilization and initial NO formation is similar for the two oxidizers when flame temperatures are the same, but the higher CO{sub 2} leads to higher concentrations of CO and nitrogen reducing intermediates at a given equivalence ratio which increases the ability of the gas phase to reduce NO. These results are supported by flat flame burner experiments which show devolatilization of nitrogen from the coal and char to be similar for air and oxy-flame conditions at a given temperature. A model of premixed combustion containing devolatilization, char oxidation and detailed kinetics captures most of the trends seen in the data. The model suggests CO is high in oxy-combustion because of dissociation of CO{sub 2}. The model also predicts a fraction (up to 20%, dependent on S.R.) of NO in air combustion can be formed via thermal processes with the source being nitrogen from the air while in oxy-combustion equilibrium drives a reduction in NO of similar magnitude. The data confirm oxy-combustion is a superior oxidizer to air for NO control because NO reduction can be achieved at higher S.R. producing better char burnout in addition to NO from recirculated flue gas being reduced

Combustion Stratification for Naphtha from CI Combustion to PPC

KAUST Repository

Vallinayagam, R.

2017-03-28

This study demonstrates the combustion stratification from conventional compression ignition (CI) combustion to partially premixed combustion (PPC). Experiments are performed in an optical CI engine at a speed of 1200 rpm for diesel and naphtha (RON = 46). The motored pressure at TDC is maintained at 35 bar and fuelMEP is kept constant at 5.1 bar to account for the difference in fuel properties between naphtha and diesel. Single injection strategy is employed and the fuel is injected at a pressure of 800 bar. Photron FASTCAM SA4 that captures in-cylinder combustion at the rate of 10000 frames per second is employed. The captured high speed video is processed to study the combustion homogeneity based on an algorithm reported in previous studies. Starting from late fuel injection timings, combustion stratification is investigated by advancing the fuel injection timings. For late start of injection (SOI), a direct link between SOI and combustion phasing is noticed. At early SOI, combustion phasing depends on both intake air temperature and SOI. In order to match the combustion phasing (CA50) of diesel, the intake air temperature is increased to 90°C for naphtha. The combustion stratification from CI to PPC is also investigated for various level of dilution by displacing oxygen with nitrogen in the intake. The start of combustion (SOC) was delayed with the increase in dilution and to compensate for this, the intake air temperature is increased. The mixture homogeneity is enhanced for higher dilution due to longer ignition delay. The results show that high speed image is initially blue and then turned yellow, indicating soot formation and oxidation. The luminosity of combustion images decreases with early SOI and increased dilution. The images are processed to generate the level of stratification based on the image intensity. The level of stratification is same for diesel and naphtha at various SOI. When O concentration in the intake is decreased to 17.7% and 14

Hybrid rocket engine, theoretical model and experiment

Science.gov (United States)

Chelaru, Teodor-Viorel; Mingireanu, Florin

2011-06-01

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

RX LAPAN Rocket data Program With Dbase III Plus

International Nuclear Information System (INIS)

Sauman

2001-01-01

The components data rocket RX LAPAN are taken from workshop product and assembling rocket RX. In this application software, the test data are organized into two data files, i.e. test file and rocket file. Besides [providing facilities to add, edit and delete data, this software provides also data manipulation facility to support analysis and identification of rocket RX failures and success

The Norwegian sounding rocket programme 1978-81

International Nuclear Information System (INIS)

Landmark, B.

1978-01-01

The Norwegian sounding rocket programme is reasonably well defined up to and including the winter of 1981/82. All the projects have been planned and will be carried out in international cooperation. Norwegian scientists so far plan to participate in a number of 24 rocket payloads over the period. Out of these 18 will be launched from the Andoya rocket range, 3 from Esrange and 3 from the siple station in the antarctic. (author)

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

Science.gov (United States)

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

1995-05-01

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

An analysis of the orbital distribution of solid rocket motor slag

Science.gov (United States)

Horstman, Matthew F.; Mulrooney, Mark

2009-01-01

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

Preliminary Design Optimization For A Supersonic Turbine For Rocket Propulsion

Science.gov (United States)

Papila, Nilay; Shyy, Wei; Griffin, Lisa; Huber, Frank; Tran, Ken; McConnaughey, Helen (Technical Monitor)

2000-01-01

In this study, we present a method for optimizing, at the preliminary design level, a supersonic turbine for rocket propulsion system application. Single-, two- and three-stage turbines are considered with the number of design variables increasing from 6 to 11 then to 15, in accordance with the number of stages. Due to its global nature and flexibility in handling different types of information, the response surface methodology (RSM) is applied in the present study. A major goal of the present Optimization effort is to balance the desire of maximizing aerodynamic performance and minimizing weight. To ascertain required predictive capability of the RSM, a two-level domain refinement approach has been adopted. The accuracy of the predicted optimal design points based on this strategy is shown to he satisfactory. Our investigation indicates that the efficiency rises quickly from single stage to 2 stages but that the increase is much less pronounced with 3 stages. A 1-stage turbine performs poorly under the engine balance boundary condition. A portion of fluid kinetic energy is lost at the turbine discharge of the 1-stage design due to high stage pressure ratio and high-energy content, mostly hydrogen, of the working fluid. Regarding the optimization technique, issues related to the design of experiments (DOE) has also been investigated. It is demonstrated that the criteria for selecting the data base exhibit significant impact on the efficiency and effectiveness of the construction of the response surface.

A 3D numerical study of LO2/GH2 supercritical combustion in the ONERA-Mascotte Test-rig configuration

Science.gov (United States)

Benmansour, Abdelkrim; Liazid, Abdelkrim; Logerais, Pierre-Olivier; Durastanti, Jean-Félix

2016-02-01

Cryogenic propellants LOx/H2 are used at very high pressure in rocket engine combustion. The description of the combustion process in such application is very complex due essentially to the supercritical regime. Ideal gas law becomes invalid. In order to try to capture the average characteristics of this combustion process, numerical computations are performed using a model based on a one-phase multi-component approach. Such work requires fluid properties and a correct definition of the mixture behavior generally described by cubic equations of state with appropriated thermodynamic relations validated against the NIST data. In this study we consider an alternative way to get the effect of real gas by testing the volume-weighted-mixing-law with association of the component transport properties using directly the NIST library data fitting including the supercritical regime range. The numerical simulations are carried out using 3D RANS approach associated with two tested turbulence models, the standard k-Epsilon model and the realizable k-Epsilon one. The combustion model is also associated with two chemical reaction mechanisms. The first one is a one-step generic chemical reaction and the second one is a two-step chemical reaction. The obtained results like temperature profiles, recirculation zones, visible flame lengths and distributions of OH species are discussed.

Theoretical and experimental studies on emissions from wood combustion

Energy Technology Data Exchange (ETDEWEB)

Skreiberg, Oeyvind

1997-12-31

This thesis discusses experiments on emissions from wood log combustion and single wood particle combustion, both caused by incomplete combustion and emissions of nitric and nitrous oxide, together with empirical and kinetic NO{sub x} modelling. Experiments were performed in three different wood stoves: a traditional stove, a staged air stove and a stove equipped with a catalytic afterburner. Ideally, biomass fuel does not give a net contribution to the greenhouse effect. However, incomplete combustion was found to result in significant greenhouse gas emissions. Empirical modelling showed the excess air ratio and the combustion chamber temperature to be the most important input variables controlling the total fuel-N to NO{sub x} conversion factor. As the result of an international round robin test of a wood stove equipped with a catalytic afterburner, particle emission measurements were found to be the best method to evaluate the environmental acceptability of the tested stove, since the particle emission level was least dependent of the national standards, test procedures and calculation procedures used. In batch single wood particle combustion experiments on an electrically heated small-scale fixed bed reactor the fuel-N to NO conversion factor varied between 0.11-0.86 depending on wood species and operating conditions. A parameter study and homogeneous kinetic modelling on a plug flow reactor showed that, depending on the combustion compliance in question, there is an optimum combination of primary excess air ratio, temperature and residence time that gives a maximum conversion of fuel-N to N{sub 2}. 70 refs., 100 figs., 26 tabs.

Assessment of exposure-response functions for rocket-emission toxicants

National Research Council Canada - National Science Library

Subcommittee on Rocket-Emission Toxicants, National Research Council

... aborted launch that results in a rocket being destroyed near the ground. Assessment of Exposure-Response Functions for Rocket-Emmission Toxicants evaluates the model and the data used for three rocket emission toxicants...

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

Science.gov (United States)

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

2004-08-01

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

Scaled Rocket Testing in Hypersonic Flow

Science.gov (United States)

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

2015-01-01

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

Effect of two-stage injection on combustion and emissions under high EGR rate on a diesel engine by fueling blends of diesel/gasoline, diesel/n-butanol, diesel/gasoline/n-butanol and pure diesel

International Nuclear Information System (INIS)

Zheng, Zunqing; Yue, Lang; Liu, Haifeng; Zhu, Yuxuan; Zhong, Xiaofan; Yao, Mingfa

2015-01-01

Highlights: • Two-stage injection using diesel blended fuel at high EGR (46%) was studied. • Blending fuels induce retarded pilot heat release and have less effect on MPRR. • Effects of injection parameters of blended fuels on emissions are similar to diesel. • Different fuels have little influence on post combustion heat release. • Small quantity post injection close to main results in better efficiency and emissions. - Abstract: The effect of two-stage injection on combustion and emission characteristics under high EGR (46%) condition were experimentally investigated. Four different fuels including pure diesel and blended fuels of diesel/gasoline, diesel/n-butanol, diesel/gasoline/n-butanol were tested. Results show that blending gasoline or/and n-butanol in diesel improves smoke emissions while induces increase in maximum pressure rise rate (MPRR). Adopting pilot injection close to main injection can effectively reduce the peak of premixed heat release rate and MPRR. However, for fuels blends with high percentage of low cetane number fuel, the effect of pilot fuel on ignition can be neglected and the improvement of MPRR is not that obvious. Pilot-main interval presents more obvious effect on smoke than pilot injection rate does, and the smoke emissions decrease with increasing pilot-main interval. A longer main-post interval results in a lower post heat release rate and prolonged combustion duration. While post injection rate has little effect on the start of ignition for post injection. The variation in fuel properties caused by blending gasoline or/and n-butanol into diesel does not impose obvious influence on post combustion. The smoke emission increases first and then declines with retard of post injection timing. Compared to diesel, the smoke emissions of blended fuels are more sensitive to the variation of post injection strategy

Integrated Composite Rocket Nozzle Extension, Phase I

Data.gov (United States)

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

Ash liberation from included minerals during combustion of pulverized coal: the relationship with char structure and burnout

Energy Technology Data Exchange (ETDEWEB)

Wu, H.; Wall, T.; Liu, G.; Bryant, G. [University of Newcastle, Callaghan, NSW (Australia). CRC for Black Coal Utilization and Dept. of Chemical Engineering

1999-12-01

In this study, the float fraction ({lt} specific gravity of 2.0) of a size cut (63-90 {mu}m) bituminous coal was combusted in a drop tube furnace (DTF) at a gas temperature of 1300{degree}C under an atmosphere of air, to investigate the ash liberation at five coal burnoff levels (35.5%, 54.3%, 70.1%, 87.1% and 95.6%). The data indicated that char structure determines the ash liberation at different burnoff levels. Fragmentation of porous char was found to be the determinative mechanism for formation of fine ash during the early and middle stages of char combustion, while coalescence of included mineral matter determines the coarse ash formed in the later stages of combustion. The investigation confirmed that the char morphology and structure play a key role in determining char fragmentation, char burnout history, and the ash liberation during combustion. 35 refs., 5 figs., 2 tabs.

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

Science.gov (United States)

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

1998-01-01

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

IEA combustion agreement : a collaborative task on alternative fuels in combustion

International Nuclear Information System (INIS)

Larmi, M.

2009-01-01

The focus of the alternative fuels in combustion task of the International Energy Agency is on high efficiency engine combustion, furnace combustion, and combustion chemistry. The objectives of the task are to develop optimum combustion for dedicated fuels by fully utilizing the physical and chemical properties of synthetic and renewable fuels; a significant reduction in carbon dioxide, NOx and particulate matter emissions; determine the minimum emission levels for dedicated fuels; and meet future emission standards of engines without or with minimum after-treatment. This presentation discussed the alternative fuels task and addressed issues such as synthetic fuel properties and benefits. The anticipated future roadmap was presented along with a list of the synthetic and renewable engine fuels to be studied, such as neat oxygenates like alcohols and ethers, biogas/methane and gas combustion, fuel blends, dual fuel combustion, high cetane number diesel fuels like synthetic Fischer-Tropsch diesel fuel and hydrogenated vegetable oil, and low CN number fuels. Implementation examples were also discussed, such as fuel spray studies in optical spray bombs; combustion research in optical engines and combustion chambers; studies on reaction kinetics of combustion and emission formation; studies on fuel properties and ignition behaviour; combustion studies on research engines; combustion optimization; implementing the optimum combustion in research engines; and emission measurements. Overall milestone examples and the overall schedule of participating countries were also presented. figs.

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

Science.gov (United States)

Hueter, Uwe; Turner, James

1998-01-01

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

Design, Fabrication and Test of a Full Scale Copper Tubular Combustion Chamber

National Research Council Canada - National Science Library

Cooley, Christine

2002-01-01

This paper presents the design fabrication and test of a full scale copper tubular combustion chamber as an enabling technology for future application in a high thrust upper-stage expander-cycle engine...

Easier Analysis With Rocket Science

Science.gov (United States)

2003-01-01

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

Internal combustion engine using premixed combustion of stratified charges

Science.gov (United States)

Marriott, Craig D [Rochester Hills, MI; Reitz, Rolf D [Madison, WI

2003-12-30

During a combustion cycle, a first stoichiometrically lean fuel charge is injected well prior to top dead center, preferably during the intake stroke. This first fuel charge is substantially mixed with the combustion chamber air during subsequent motion of the piston towards top dead center. A subsequent fuel charge is then injected prior to top dead center to create a stratified, locally richer mixture (but still leaner than stoichiometric) within the combustion chamber. The locally rich region within the combustion chamber has sufficient fuel density to autoignite, and its self-ignition serves to activate ignition for the lean mixture existing within the remainder of the combustion chamber. Because the mixture within the combustion chamber is overall premixed and relatively lean, NO.sub.x and soot production are significantly diminished.

Formaldehyde and acetaldehyde emissions from residential wood combustion in Portugal

Science.gov (United States)

Cerqueira, Mário; Gomes, Luís; Tarelho, Luís; Pio, Casimiro

2013-06-01

A series of experiments were conducted to characterize formaldehyde and acetaldehyde emissions from residential combustion of common wood species growing in Portugal. Five types of wood were investigated: maritime pine (Pinus pinaster), eucalyptus (Eucalyptus globulus), cork oak (Quercus suber), holm oak (Quercus rotundifolia) and pyrenean oak (Quercus pyrenaica). Laboratory experiments were performed with a typical wood stove used for domestic heating in Portugal and operating under realistic home conditions. Aldehydes were sampled from diluted combustion flue gas using silica cartridges coated with 2,4-dinitrophenylhydrazine and analyzed by high performance liquid chromatography with diode array detection. The average formaldehyde to acetaldehyde concentration ratio (molar basis) in the stove flue gas was in the range of 2.1-2.9. Among the tested wood types, pyrenean oak produced the highest emissions for both formaldehyde and acetaldehyde: 1772 ± 649 and 1110 ± 454 mg kg-1 biomass burned (dry basis), respectively. By contrast, maritime pine produced the lowest emissions: 653 ± 151 and 371 ± 162 mg kg-1 biomass (dry basis) burned, respectively. Aldehydes were sampled separately during distinct periods of the holm oak wood combustion cycles. Significant variations in the flue gas concentrations were found, with higher values measured during the devolatilization stage than in the flaming and smoldering stages.

Water Impact Prediction Tool for Recoverable Rockets

Science.gov (United States)

Rooker, William; Glaese, John; Clayton, Joe

2011-01-01

Reusing components from a rocket launch can be cost saving. NASA's space shuttle system has reusable components that return to the Earth and impact the ocean. A primary example is the Space Shuttle Solid Rocket Booster (SRB) that descends on parachutes to the Earth after separation and impacts the ocean. Water impact generates significant structural loads that can damage the booster, so it is important to study this event in detail in the design of the recovery system. Some recent examples of damage due to water impact include the Ares I-X First Stage deformation as seen in Figure 1 and the loss of the SpaceX Falcon 9 First Stage.To ensure that a component can be recovered or that the design of the recovery system is adequate, an adequate set of structural loads is necessary for use in failure assessments. However, this task is difficult since there are many conditions that affect how a component impacts the water and the resulting structural loading that a component sees. These conditions include the angle of impact with respect to the water, the horizontal and vertical velocities, the rotation rate, the wave height and speed, and many others. There have been attempts to simulate water impact. One approach is to analyze water impact using explicit finite element techniques such as those employed by the LS-Dyna tool [1]. Though very detailed, this approach is time consuming and would not be suitable for running Monte Carlo or optimization analyses. The purpose of this paper is to describe a multi-body simulation tool that runs quickly and that captures the environments a component might see. The simulation incorporates the air and water interaction with the component, the component dynamics (i.e. modes and mode shapes), any applicable parachutes and lines, the interaction of winds and gusts, and the wave height and speed. It is capable of quickly conducting Monte Carlo studies to better capture the environments and genetic algorithm optimizations to reproduce a

Thermogravimetric and Kinetic Analysis of Raw and Torrefied Biomass Combustion

Directory of Open Access Journals (Sweden)

Kopczyński Marcin

2015-06-01

Full Text Available The use of torrefied biomass as a substitute for untreated biomass may decrease some technological barriers that exist in biomass co-firing technologies e.g. low grindability, high moisture content, low energy density and hydrophilic nature of raw biomass. In this study the TG-MS-FTIR analysis and kinetic analysis of willow (Salix viminalis L. and samples torrefied at 200, 220, 240, 260, 280 and 300 °C (TSWE 200, 220, 240, 260, 280 and 300, were performed. The TG-DTG curves show that in the case of willow and torrefied samples TSWE 200, 220, 240 and 260 there are pyrolysis and combustion stages, while in the case of TSWE 280 and 300 samples the peak associated with the pyrolysis process is negligible, in contrast to the peak associated with the combustion process. Analysis of the TG-MS results shows m/z signals of 18, 28, 29 and 44, which probably represent H2O, CO and CO2. The gaseous products were generated in two distinct ranges of temperature. H2O, CO and CO2 were produced in the 500 K to 650 K range with maximum yields at approximately 600 K. In the second range of temperature, 650 K to 800 K, only CO2 was produced with maximum yields at approximately 710 K as a main product of combustion process. Analysis of the FTIR shows that the main gaseous products of the combustion process were H2O, CO2, CO and some organics including bonds: C=O (acids, aldehydes and ketones, C=C (alkenes, aromatics, C-O-C (ethers and C-OH. Lignin mainly contributes hydrocarbons (3000-2800 cm−1, while cellulose is the dominant origin of aldehydes (2860-2770 cm−1 and carboxylic acids (1790-1650 cm−1. Hydrocarbons, aldehydes, ketones and various acids were also generated from hemicellulose (1790-1650 cm−1. In the kinetic analysis, the two-steps first order model (F1F1 was assumed. Activation energy (Ea values for the first stage (pyrolysis increased with increasing torrefaction temperature from 93 to 133 kJ/mol, while for the second stage (combustion it

Influence of the microwave irradiation dewatering on the combustion characteristics of Chinese brown coals

Science.gov (United States)

Ge, Lichao; Feng, Hongcui; Xu, Chang; Zhang, Yanwei; Wang, Zhihua

2018-02-01

This study investigates the influence of microwave irradiation on coal composition, pore structure, coal rank, and combustion characteristics of typical brown coals in China. Results show that the upgrading process significantly decreased the inherent moisture, and increased calorific value and fixed carbon content. After upgrading, pore distribution extended to micropore region, oxygen functional groups were reduced and destroyed, and the apparent aromaticity increased suggesting an improvement in the coal rank. Based on thermogravimetric analysis, the combustion processes of upgraded coals were delayed toward the high temperature region, and the temperatures of ignition, peak and burnout increased. Based on the average combustion rate and comprehensive combustion parameter, the upgraded coals performed better compared with raw brown coals and a high rank coal. In ignition and burnout segments, the activation energy increased but exhibited a decrease in the combustion stage.

Flow-Structural Interaction in Solid Rocket Motors

National Research Council Canada - National Science Library

Murdock, John

2004-01-01

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

Combustion Stratification for Naphtha from CI Combustion to PPC

KAUST Repository

Vallinayagam, R.; Vedharaj, S.; An, Yanzhao; Dawood, Alaaeldin; Izadi Najafabadi, Mohammad; Somers, Bart; Johansson, Bengt

2017-01-01

This study demonstrates the combustion stratification from conventional compression ignition (CI) combustion to partially premixed combustion (PPC). Experiments are performed in an optical CI engine at a speed of 1200 rpm for diesel and naphtha (RON

A new facility for advanced rocket propulsion research

Science.gov (United States)

Zoeckler, Joseph G.; Green, James M.; Raitano, Paul

1993-06-01

A new test facility was constructed at the NASA Lewis Research Center Rocket Laboratory for the purpose of conducting rocket propulsion research at up to 8.9 kN (2000 lbf) thrust, using liquid oxygen and gaseous hydrogen propellants. A laser room adjacent to the test cell provides access to the rocket engine for advanced laser diagnostic systems. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods, with rapid turnover between programs. These capabilities make the new test facility an important asset for basic and applied rocket propulsion research.

Fuel and combustion stratification study of Partially Premixed Combustion

OpenAIRE

Izadi Najafabadi, M.; Dam, N.; Somers, B.; Johansson, B.

2016-01-01

Relatively high levels of stratification is one of the main advantages of Partially Premixed Combustion (PPC) over the Homogeneous Charge Compression Ignition (HCCI) concept. Fuel stratification smoothens heat release and improves controllability of this kind of combustion. However, the lack of a clear definition of “fuel and combustion stratifications” is obvious in literature. Hence, it is difficult to compare stratification levels of different PPC strategies or other combustion concepts. T...

Development of flameless combustion; Desarrollo de la combustion sin flama