539 research outputs found
Review of Combustion Stability Characteristics of Swirl Coaxial Element Injectors
Liquid propellant rocket engine injectors using coaxial elements where the center liquid is swirled have become more common in the United States over the past several decades, although primarily for technology or advanced development programs. Currently, only one flight engine operates with this element type in the United States (the RL10 engine), while the element type is very common in Russian (and ex-Soviet) liquid propellant rocket engines. In the United States, the understanding of combustion stability characteristics of swirl coaxial element injectors is still very limited, despite the influx of experimental and theoretical information from Russia. The empirical and theoretical understanding is much less advanced than for the other prevalent liquid propellant rocket injector element types, the shear coaxial and like-on-like paired doublet. This paper compiles, compares and explores the combustion stability characteristics of swirl coaxial element injectors tested in the United States, dating back to J-2 and RL-10 development, and extending to very recent programs at the NASA MSFC using liquid oxygen and liquid methane and kerosene propellants. Included in this study are several other relatively recent design and test programs, including the Space Transportation Main Engine (STME), COBRA, J-2X, and the Common Extensible Cryogenic Engine (CECE). A presentation of the basic data characteristics is included, followed by an evaluation by several analysis techniques, including those included in Rocket Combustor Interactive Design and Analysis Computer Program (ROCCID), and methodologies described by Hewitt and Bazarov
Analyses of Longitudinal Mode Combustion Instability in J-2X Gas Generator Development
The National Aeronautics and Space Administration (NASA) and Pratt & Whitney Rocketdyne are developing a liquid oxygen/liquid hydrogen rocket engine for future upper stage and trans-lunar applications. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. The contract for development was let to Pratt & Whitney Rocketdyne in 2006. Over the past several years, development of the gas generator for the J-2X engine has progressed through a variety of workhorse injector, chamber, and feed system configurations on the component test stand at the NASA Marshall Space Flight Center (MSFC). Several of the initial configurations resulted in combustion instability of the workhorse gas generator assembly at a frequency near the first longitudinal mode of the combustion chamber. In this paper, several aspects of these combustion instabilities are discussed, including injector, combustion chamber, feed system, and nozzle influences. To ensure elimination of the instabilities at the engine level, and to understand the stability margin, the gas generator system has been modeled at the NASA MSFC with two techniques, the Rocket Combustor Interaction Design and Analysis (ROCCID) code and a lumped-parameter MATLAB(TradeMark) model created as an alternative calculation to the ROCCID methodology. To correctly predict the instability characteristics of all the chamber and injector geometries and test conditions as a whole, several inputs to the submodels in ROCCID and the MATLAB(TradeMark) model were modified. Extensive sensitivity calculations were conducted to determine how to model and anchor a lumped-parameter injector response, and finite-element and acoustic analyses were conducted on several complicated combustion chamber geometries to determine how to model and anchor the chamber response. These modifications and their ramification for future stability analyses of this type are discussed
АСУ распылением жидкого топлива на примере мазута
АСУ подачи мазута в камеру сгорания топлива ПГ с учетом необходимого воздуха и контролированием температуры горения в топке.ACS to supply fuel oil to the combustion chamber of the SG with necessary air and to control the combustion temperature of furnace
Comparison of Single-Element and Multi-Element Oxygen/RP-1 Oxidizer-Rich Staged-Combustion Injector Hot-Fire Test Results
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 designed, fabricated, assembled and hot-fire tested an oxygen/hydrocarbon propellant multi-element integrated test article that included an oxidizer-rich oxygen/hydrocarbon propellant preburner and a staged-combustion main injector. Also as part of this project, the Air Force Research Laboratory fabricated single-element main injectors of the same designs as used in the NASA multi-element injectors, and tested them in a staged-combustion integrated test article that used an oxidizer-rich oxygen/hydrogen propellant preburner. Final results of the multi-element and single-element staged-combustion main injector test programs are described in companion papers at this JANNAF meeting. The design, development, and preliminary test results of these main injectors have also been described in previous JANNAF papers. The main injector element designs were all based on relatively conventional gas-centered swirl coaxial injector element configurations such as used in Russian RD-170 and NK-33 engines, and planned for use in future U.S.-built experimental engine systems such as the Hydrocarbon Boost program demonstration engine. Four different elements were tested in both the multi-element and single-element main injectors, at similar combustion chamber pressures, chamber contraction ratios, and mixture ratios. Variations of the element features included recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. This paper compares the hydraulics, combustion performance, stability, and compatibility characteristics of the single-element and multi-element injectors operated at similar conditions. The single-element hardware is shown to have captured a significant level of the operability of the multi-element hardware
Hot-Fire Test Results of Liquid Oxygen/RP-2 Multi-Element Oxidizer-Rich Preburners
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. To supply the oxidizer-rich combustion products to the main injector of the integrated test article, existing subscale preburner injectors from a previous NASA-funded oxidizer-rich staged combustion engine development program were utilized. For the integrated test article, existing and newly designed and fabricated 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. However, before one of the preburners was used in the integrated test article, it was first hot-fire tested at length to prove it could provide the hot exhaust gas mean temperature, thermal uniformity and combustion stability necessary to perform in the integrated test article experiment. This paper presents results from hot-fire testing of several preburner injectors in a representative combustion chamber with a sonic throat. Hydraulic, combustion performance, exhaust gas thermal uniformity, and combustion stability data are presented. Results from combustion stability modeling of these test results are described in a companion paper at this JANNAF conference, while hot-fire test results of the preburner injector in the integrated test article are described in another companion paper
Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article
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 combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics
A Test of Interventions to Increase Adherence to Hypertensive Medical Regimens
Low rates of adherence to hypertensive therapy limit patients' securing the full benefits of treatment. While some factors related to adherence have been identified, research on the effectiveness of interventions to increase adherence levels is sparse. The present study was designed to assess the impact of a series of different interventions on a group of some 400 patients, all under the care of private physicians in a small community. A factorial design was employed to deliver four, sequential educational interventions, about four months apart, to randomly selected sub-groups. Interviews before and after each intervention provided information concerning self-reported adherence, health status, health beliefs, and personal characteristics. Pertinent medical records and pharmacy data were also obtained. The first intervention—printed material—did not significantly affect adherence. The second and fourth interventions—nurse telephone calls and social support—each increased medication taking and the third intervention—self-monitoring—led to better weight control. There was no cumulative impact of the interventions and different aspects of regimens were not signiticantly related to one another.Peer Reviewedhttp://deepblue.lib.umich.edu/bitstream/2027.42/67164/2/10.1177_109019818100800303.pd
Effect of Injector Exit Geometry on Atomization of a Liquid-Liquid Double Swirl Coaxial Injector using Non-Invasive Laser, Optical, and X-ray Techniques
The spray characteristics of a liquid-liquid double swirl coaxial injector were studied using non-invasive optical, laser, and X-ray diagnostics. Phase Doppler interferometry was used to characterize droplet statistics and non-dimensional droplet parameters over a range of inlet conditions and for various fluids allowing for a study on the role of fluid properties on atomization. Based on the atomization statistics and observed trends from high-speed images, a description of breakup regimes over a range of Reynolds and Weber numbers was created. Next, X-ray computed tomography scans revealed that the spray cone was distinctively non-uniform and comprised of several pockets of increased mass concentration which varied with injection pressure. Finally, a parametric study of injector exit geometry demonstrated that spray breakup time, breakup type, and sheet stability could be controlled with exit geometry. Implications for these data on injector stability and atomization efficiency are discussed considering the desired performance characteristics of liquid-liquid rocket injectors
Performance, Stability and Compatibility of Oxygen/RP-1 Multi-Element Oxidizer-Rich Staged-Combustion Injectors
In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting
Анализ и совершенствование технологии организации строительства зданий и сооружений магистральных газопроводов с использованием технической мелиорации грунтов
РЕФЕРАТ
Выпускная квалификационная работа 98 с., 8 рис., 6 табл., 21 источник.
Ключевые слова: Сооружение объектов магистральных газопроводов
Объектом исследования является (ются): Изучение технологии производства работ и элементов технологической системы, рациональное сочетание которых позволяет создавать эффективные конструктивно–технологические решения в результате термического укрепления грунтов с помощью мобильных СВЧ–генераторов.
Цель работы: Выбор оптимальных решений при сооружении оснований зданий посредством термического укрепления грунта на строительной площадке с помощью мобильных СВЧ – установок.
В процессе исследования проводились: Расчёт тепловых режимов обработки грунтовых массивов СВЧ – полем, расчёт керамического фундамента на прочность при изгибе, расчёт керамическогоREPORT
Final qualifying work 98 p., 8 fig., 6 tables., 21 source.
Keywords: Building of objects of main gas pipelines
A research object is (ются) : Study of technology of production of works and elements of the technological system, rational combination of that allows to create effective structurally-technological decisions as a result of the thermal strengthening of soils by means of mobile СВЧ-генераторов.
Aim of work : Choice of optimal decisions at building of grounds of building by means of the thermal strengthening of soil on a site area by means of mobile СВЧ - options.
In the process of research conducted: Calculation of the thermal modes of treatment of the ground arrays СВЧ - we weed, calculation of ceramic foundation on durability at a bend, calculation of ceramic foundation
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