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1

Infrared Imagery of Solid Rocket Exhaust Plumes  

NASA Technical Reports Server (NTRS)

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

Moran, Robert P.; Houston, Janice D.

2011-01-01

2

Implementation of microwave transmissions for rocket exhaust plume diagnostics  

NASA Astrophysics Data System (ADS)

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

Coutu, Nicholas George

3

Bipropellant rocket exhaust plume analysis on the Galileo spacecraft  

NASA Technical Reports Server (NTRS)

This paper describes efforts to quantify the contaminant flow field produced by 10 N thrust bipropellant rocket engines used on the Galileo spacecraft. The prediction of the composition of the rocket exhaust by conventional techniques is found to be inadequate to explain experimental observations of contaminant deposition on moderately cold (200 K) surfaces. It is hypothesized that low volatility contaminants are formed by chemical reactions which occur on the surfaces. The flow field calculations performed using the direct simulation Monte Carlo method give the expected result that the use of line-of-sight plume shields may have very little effect on the flux of vapor phase contaminant species to a surface, especially if the plume shields are located so close to the engine that the interaction of the plume with the shield is in the transition flow regime. It is shown that significant variations in the exhaust plume composition caused by nonequilibrium effects in the flow field lead to very low concentrations of species which have high molecular weights in the more rarefied regions of the flow field. Recommendations for the design of spacecraft plume shields and further work are made.

Guernsey, C. S.; Mcgregor, R. D.

1986-01-01

4

Environmental Effects of Space Shuttle Solid Rocket Motor Exhaust Plumes  

NASA Technical Reports Server (NTRS)

The deposition of NOx and HCl in the stratosphere from the space shuttle solid rocket motors (SRM) and exhaust plume is discussed. A detailed comparison between stratospheric deposition rates using the baseline SRM propellant and an alternate propellant, which replaces ammonium perchlorate by ammonium nitrate, shows the total NOx deposition rate to be approximately the same for each propellant. For both propellants the ratio of the deposition rates of NOx to total chlorine-containing species is negligibly small. Rocket exhaust ground cloud transport processes in the troposphere are also examined. A brief critique of the multilayer diffusion models (presently used for predicting pollutant deposition in the troposphere) is presented, and some detailed cloud rise calculations are compared with data for Titan 3C launches. The results show that, when launch time meteorological data are used as input, the model can reasonably predict measured cloud stabilization heights.

Hwang, B.; Pergament, H. S.

1976-01-01

5

General Computer Program for Calculation of Radiation from Inhomogeneous, Nonisobaric, Nonisothermal Rocket Exhaust Plume  

NASA Technical Reports Server (NTRS)

Computer program evaluates radiation from an axisymmetric gas body with water vapor, carbon dioxide, carbon monoxide, and solid carbon particles as radiating constituents, and hydrogen as a nonradiating constituent. The program provides a convenient method of evaluating a great many problems of radiation from rocket exhaust plumes.

Dash, M. J.; Huffaker, R. M.

1968-01-01

6

Determination of Combustion Product Radicals in a Hydrocarbon Fueled Rocket Exhaust Plume  

NASA Technical Reports Server (NTRS)

The identification of metallic effluent materials in a rocket engine exhaust plume indicates the health of the engine. Since 1989, emission spectroscopy of the plume of the Space Shuttle Main Engine (SSME) has been used for ground testing at NASA's Stennis Space Center (SSC). This technique allows the identification and quantification of alloys from the metallic elements observed in the plume. With the prospect of hydrocarbon-fueled rocket engines, such as Rocket Propellant 1 (RP-1) or methane (CH4) fueled engines being considered for use in future space flight systems, the contributions of intermediate or final combustion products resulting from the hydrocarbon fuels are of great interest. The effect of several diatomic molecular radicals, such as Carbon Dioxide , Carbon Monoxide, Molecular Carbon, Methylene Radical, Cyanide or Cyano Radical, and Nitric Oxide, needs to be identified and the effects of their band systems on the spectral region from 300 nm to 850 nm determined. Hydrocarbon-fueled rocket engines will play a prominent role in future space exploration programs. Although hydrogen fuel provides for higher engine performance, hydrocarbon fuels are denser, safer to handle, and less costly. For hydrocarbon-fueled engines using RP-1 or CH4 , the plume is different from a hydrogen fueled engine due to the presence of several other species, such as CO2, C2, CO, CH, CN, and NO, in the exhaust plume, in addition to the standard H2O and OH. These species occur as intermediate or final combustion products or as a result of mixing of the hot plume with the atmosphere. Exhaust plume emission spectroscopy has emerged as a comprehensive non-intrusive sensing technology which can be applied to a wide variety of engine performance conditions with a high degree of sensitivity and specificity. Stennis Space Center researchers have been in the forefront of advancing experimental techniques and developing theoretical approaches in order to bring this technology to a more mature stage.

Langford, Lester A.; Allgood, Daniel C.; Junell, Justin C.

2007-01-01

7

An experimental and computational study of moderately underexpanded rocket exhaust plumes in a co-flowing hypersonic free stream  

SciTech Connect

Rocket plume exhaust structures are aerodynamically and thermochemically very complex and the prediction of plume properties such as temperature, velocity, pressure, chemical species concentrations and turbulence properties is a formidable task as there are no definitive models for viscous and chemical effects. Contemporary computational techniques are still in their infancy and cannot yet reliably predict plume properties. Only through validation of computer codes using experimental data, can computational models be developed to the point where they can be confidently used as design and predictive tools. The motivation for this study was to acquire well defined data for rocket plumes at low altitude hypersonic flight conditions so that the above issues could be investigated.

Morris, N.; Buttsworth, D.; Jones, T.; Brescianini, C. [Univ. of Oxford (United Kingdom)]|[Macquarie Univ., Sydney (Australia)

1995-09-01

8

On-board Optical Spectrometry for Detection of Mixture Ratio and Eroded Materials in Rocket Engine Exhaust Plume  

NASA Technical Reports Server (NTRS)

Optical spectrometry can provide means to characterize rocket engine exhaust plume impurities due to eroded materials, as well as combustion mixture ratio without any interference with plume. Fiberoptic probes and cables were designed, fabricated and installed on Space Shuttle Main Engines (SSME), allowing monitoring of the plume spectra in real time with a Commercial of the Shelf (COTS) fiberoptic spectrometer, located in a test-stand control room. The probes and the cables survived the harsh engine environments for numerous hot-fire tests. When the plume was seeded with a nickel alloy powder, the spectrometer was able to successfully detect all the metallic and OH radical spectra from 300 to 800 nanometers.

Barkhoudarian, Sarkis; Kittinger, Scott

2006-01-01

9

Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines  

NASA Technical Reports Server (NTRS)

The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

1975-01-01

10

Monte Carlo simulation of solid rocket exhaust plumes at high altitude  

NASA Astrophysics Data System (ADS)

The simulation of high altitude exhaust plumes from solid propellant rockets involves numerous complex physical processes which are not adequately understood. The work presented in this thesis aims at advancing the current state of modeling capabilities for these flows by better handling gas-solid interactions, particle rotation and shape effects, widely varying Knudsen number regimes, and radiation transport. First, using the direct simulation Monte Carlo (DSMC) method as a basis, condensed-phase particles are incorporated into the simulation of a rarefied gas flow. An existing method for the determination of momentum and energy exchange rates between a locally free molecular gas and a solid sphere is extended to nonspherical or rotating particles, by accounting for two-way coupling between the particles and gas. A nonequilibrium crystallization model for liquid Al2O3 droplets is also presented. A new near-equilibrium flow scheme is introduced for efficient gas phase simulation, which is shown to be well suited for DSMC-continuum hybrid two phase flow simulation. A number of Monte Carlo methods have recently been proposed for the simulation of near-equilibrium flows in a manner similar to DSMC. Based on existing methods for the ellipsoidal statistical Bhatnagar-Gross-Krook (ES-BGK) model of the Boltzmann equation, improved procedures are developed to enforce momentum and energy conservation, and to allow for rotational-translational energy exchange in a diatomic gas. In addition, a Monte Carlo ray trace (MCRT) model is developed for plume radiation analysis. Emission, absorption and anisotropic scattering are considered for non-gray condensed phase particles in a flowfield of arbitrary optical thickness. To evaluate the overall performance of the proposed schemes, simulations are performed for a representative solid rocket plume flow. Limited comparisons are made between calculated UV radiance values and measured values from a flight experiment, and relatively good agreement is found. A series of parametric studies involving simulations of this same flow is used to evaluate the influence of physical processes and input parameters related to gas-particle interaction, particle radiation, and the presence of soot. Particle accommodation coefficients and absorption index values are found to significantly influence results, while effects of soot and Al2O3 particle rotation are shown to be negligible.

Burt, Jonathan Matthew

11

Modification of Roberts' Theory for Rocket Exhaust Plumes Eroding Lunar Soil  

NASA Technical Reports Server (NTRS)

In preparation for the Apollo program, Leonard Roberts developed a remarkable analytical theory that predicts the blowing of lunar soil and dust beneath a rocket exhaust plume. Roberts' assumed that the erosion rate is determined by the "excess shear stress" in the gas (the amount of shear stress greater than what causes grains to roll). The acceleration of particles to their final velocity in the gas consumed a portion of the shear stress. The erosion rate continues to increase until the excess shear stress is exactly consumed, thus determining the erosion rate. He calculated the largest and smallest particles that could be eroded based on forces at the particle scale, but the erosion rate equation assumes that only one particle size exists in the soil. He assumed that particle ejection angles are determined entirely by the shape of the terrain, which acts like a ballistic ramp, the particle aerodynamics being negligible. The predicted erosion rate and particle upper size limit appeared to be within an order of magnitude of small-scale terrestrial experiments, but could not be tested more quantitatively at the time. The lower particle size limit and ejection angle predictions were not tested.

Metzger, Philip T.; Lane, John E.; Immer, Christopher D.

2008-01-01

12

Modeling The Interaction Between a Rocket Plume, Scoured Regolith, and a Plume Deflection Fence  

E-print Network

Modeling The Interaction Between a Rocket Plume, Scoured Regolith, and a Plume Deflection Fence A in the near vicinity of the engine. As the rocket exhaust expands further towards vacuum, continuum As a lunar lander approaches the surface, the rocket engine exhaust plume strikes the ground causing dust

Lightsey, Glenn

13

Transmittance and Radiance Computations for Rocket Engine Plume Environments  

NASA Technical Reports Server (NTRS)

Emission and absorption characteristics of several atmospheric and combustion species have been studied and are presented with reference to rocket engine plume environments. The effects of clous, rain, and fog on plume radiance/transmittance has also been studied.Preliminary results for the radiance from the exhaust plume of the space shuttle main engine are shown and discussed.

Tejwani, Gopal D.

2003-01-01

14

Rocket plume temperature measurement by wire welded thermocouples  

NASA Astrophysics Data System (ADS)

The plume of solid rocket motor is a high velocity flow with high temperature. Temperature distribution in the plume is of great interest for analyzing the compatibility of rocket weapon system. The high temperature exhausted flow field would cause damage on certain equipment and loading vehicles. An instantaneous temperature field with sharp step is established by the exhausted flow field of rocket motor. The increasing rate of the step depends on the flow velocity at cross section of nozzle exit. To perform an accurate measurement of temperature inside the flow field, a thermocouple must be sturdy enough to endure the flow impingement. In the meantime, the thermocouple must have a short time constant to trace the temperature fluctuation in flow field and a small size to avoid disturbing the flow field severely. The dynamic performance of the thermocouples used in exhausted flow temperature measurement must be evaluated before the experiment. The thermocouple which can be used in measuring the temperature distribution in rocket plume was presented in this paper. A NAMNAC (R) self-renew-erode thermocouples with a nominal time constant of 10 microseconds was used as a reference in a dynamic calibration test for this kind of thermocouple. The thermocouple could trace the temperature increase in the exhausted flow perfectly. This kind of thermocouples was used in several real tests of rocket motors, such as the temperature in free exhausted flow field of a stationary rocket motor test, the stagnate temperature in a shock flow field during the launching of a rocket, and the temperature in a launch tube.

Xu, Qiang

2006-05-01

15

Rocket plume properties measured in space simulators  

NASA Technical Reports Server (NTRS)

Molecular sink facility and 25-foot space simulator have been used to distinguish nature of exhaust plumes from nozzles with relatively large internal boundary layer flow. Plume density has been measured by electron beam/photomultiplier system.

Stephens, J. B.; Herrera, J. G.

1973-01-01

16

Rocket Engine Plume Diagnostics at Stennis Space Center  

NASA Technical Reports Server (NTRS)

The Stennis Space Center has been at the forefront of development and application of exhaust plume spectroscopy to rocket engine health monitoring since 1989. Various spectroscopic techniques, such as emission, absorption, FTIR, LIF, and CARS, have been considered for application at the engine test stands. By far the most successful technology h a been exhaust plume emission spectroscopy. In particular, its application to the Space Shuttle Main Engine (SSME) ground test health monitoring has been invaluable in various engine testing and development activities at SSC since 1989. On several occasions, plume diagnostic methods have successfully detected a problem with one or more components of an engine long before any other sensor indicated a problem. More often, they provide corroboration for a failure mode, if any occurred during an engine test. This paper gives a brief overview of our instrumentation and computational systems for rocket engine plume diagnostics at SSC. Some examples of successful application of exhaust plume spectroscopy (emission as well as absorption) to the SSME testing are presented. Our on-going plume diagnostics technology development projects and future requirements are discussed.

Tejwani, Gopal D.; Langford, Lester A.; VanDyke, David B.; McVay, Gregory P.; Thurman, Charles C.

2003-01-01

17

Effect of Soot Particles on Supersonic Rocket Plume Properties  

NASA Astrophysics Data System (ADS)

Plumes from hydrocarbon-fueled rockets usually contain some amount of soot. In spite of the small amount, such soot particles can play a critical role in the characteristics of the infrared radiation emission since soot radiates a continuous, near-blackbody spectrum. The contribution of the soot to the plume radiation depends on the amount of soot, the physical properties of the particles, their concentration, and their temperature distribution in the flow field. The trajectories of solid particles and their temperatures can differ from those of the gas due to the particle mechanical and thermal inertia. CFD FLUENT code for solving two-phase Navier-Stokes equations coupled with chemical reactions and soot particle combustion was applied for exhaust plume simulations. Exhaust plumes with soot mass loading of 2% were simulated for three altitudes of 2 km, 8 km and 16 km. Radial distributions of the cloud particle density were obtained for different distances downstream the exhaust nozzle. As a result of the particle deceleration at the boundary layer inside the nozzle the particle concentration increased at the plume periphery. The particle temperature was higher than the gaseous temperature of the plume. The temperature difference between the soot particle and gas along corresponding trajectories was about 5-10%. The infrared radiation from the plumes with carbon soot was calculated. Its intensity was found to be dependent on the particle distribution in the plume.

Gaissinski, Igor; Levy, Yeshayahou; Lev, Mikhael; Sherbaum, Valery

2012-06-01

18

Optical studies of rocket exhaust trails and artificial noctilucent clouds produced by Soyuz rocket launches  

NASA Astrophysics Data System (ADS)

Detailed tracing of an exhaust plume from a rocket's initial trajectory is a scientifically and diagnostically useful technique. It can provide detailed information on the atmosphere's mean winds, wind shears, turbulent regime, and physical state over a wide altitude range from 50 to 200 km. We analyze Soyuz rocket exhaust plumes from Plesetsk on 21 May 2009 and 27 June 2011, which uncovered significantly different atmospheric states and underlying dynamics. The first case showed highly dynamical conditions in the mesosphere, characterized by vortex structures, wind shears, and small-scale turbulent eddies. The estimated turbulent energy dissipation rates ranged 330-460 mW kg-1. A characteristic balloon-shaped trail was observed at altitudes between 105 and 160 km, having rapid expansion rates of 500-800 m s-1 over the time period of 2 min which can be explained by complex gas dynamic processes in the rocket wake involving the collision of shock waves. In the second case, we show evidence that the rocket exhaust trail persisted without any changes during its motion from Plesetsk via Denmark to the UK for 9 h, indicating extremely stable atmospheric conditions. This case introduces a new state of the summer mesosphere—remarkably quiet conditions, probably never observed before. The rocket plumes studied, related to the initial rocket trajectory, are essentially twilight phenomena as seen from the ground using wideband spectrum cameras, that is, the Sun should be below the horizon by 6°. For the first time, we analyze the dynamics of rocket exhaust products at the initial trajectory in the mesosphere and lower thermosphere using detailed photographic imaging taken from the ground.

Dalin, P.; Perminov, V.; Pertsev, N.; Dubietis, A.; Zadorozhny, A.; Smirnov, A.; Mezentsev, A.; Frandsen, S.; Grønne, J.; Hansen, O.; Andersen, H.; McEachran, I.; McEwan, T.; Rowlands, J.; Meyerdierks, H.; Zalcik, M.; Connors, M.; Schofield, I.; Veselovsky, I.

2013-07-01

19

Exhaust Nozzle Plume and Shock Wave Interaction  

NASA Technical Reports Server (NTRS)

Fundamental research for sonic boom reduction is needed to quantify the interaction of shock waves generated from the aircraft wing or tail surfaces with the exhaust plume. Both the nozzle exhaust plume shape and the tail shock shape may be affected by an interaction that may alter the vehicle sonic boom signature. The plume and shock interaction was studied using Computational Fluid Dynamics simulation on two types of convergent-divergent nozzles and a simple wedge shock generator. The nozzle plume effects on the lower wedge compression region are evaluated for two- and three-dimensional nozzle plumes. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the deflected lower plume boundary. The sonic boom pressure signature of the wedge is modified by the presence of the plume, and the computational predictions show significant (8 to 15 percent) changes in shock amplitude.

Castner, Raymond S.; Elmiligui, Alaa; Cliff, Susan

2013-01-01

20

Atmospheric scavenging of solid rocket exhaust effluents  

NASA Technical Reports Server (NTRS)

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

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

1978-01-01

21

A study of exhaust plume interactions with external flow by the hydraulic analogy  

E-print Network

visualization experiments were conducted utilizing the hydraulic analogy. The first experiment simulated the plume interactions at different rocket stagnation pressures and different external flow Mach numbers. The rocket nozzle was designed for an exit speed.... Most rocket nozzles are designed to have exit pressures equal to atmospheric pressure at low altitudes. When a rocket travels into the upper atmosphere, the pressure of the exhaust gasses are higher than the surrounding atmosphere. This large...

Lawton, Stephen Hayes

2012-06-07

22

Propagation of light through ship exhaust plumes  

NASA Astrophysics Data System (ADS)

Looking through the atmosphere, it is sometimes difficult to see the details of an object. Effects like scintillation and blur are the cause of these difficulties. Exhaust plumes of e.g. a ship can cause extreme scintillation and blur, making it even harder to see the details of what lies behind the plume. Exhaust plumes come in different shapes, sizes, and opaqueness and depending on atmospheric parameters like wind speed and direction, as well as engine settings (power, gas or diesel, etc.). A CFD model is used to determine the plume's flow field outside the stack on the basis of exhaust flow properties, the interaction with the superstructure of the ship, the meteorological conditions and the interaction of ship's motion and atmospheric wind fields. A modified version of the NIRATAM code performs the gas radiation calculations and provides the radiant intensity of the (hot) exhaust gases and the transmission of the atmosphere around the plume is modeled with MODTRAN. This allows assessing the irradiance of a sensor positioned at some distance from the ship and its plume, as function of the conditions that influence the spatial distribution and thermal properties of the plume. Furthermore, an assessment can be made of the probability of detecting objects behind the plume. This plume module will be incorporated in the TNO EOSTAR-model, which provides estimates of detection range and image quality of EO-sensors under varying meteorological conditions.

van Iersel, M.; Mack, A.; van Eijk, A. M. J.; Schleijpen, H. M. A.

2014-10-01

23

Improvements in rocket engine nozzle and high altitude plume computations  

NASA Technical Reports Server (NTRS)

The characteristics of rocket exhaust flow fields are very complex, and many phenomena are involved. Previously, it was necessary to use a multitude of codes to treat a nozzle/plume flow in detail. In connection with both computational and economic standpoints, however, it is desirable to have a single code which can treat all the dominant phenomena in a rocket nozzle/plume flow field. The present investigation has the objective to describe a nozzle plume flowfield code which has capabilities that do not presently exist in a single computer code. The RAMP code considered by Penny et al. (1976) was used as a basis in the development of the new code. The basic RAMP employs modular construction and provides a two-phase, reacting gas, supersonic inviscid nozzle/plume solution. Other capabilities needed, which in most cases already exist in other computer codes, were incorporated into the RAMP code to enhance its usefulness. Attention is given to results of plume calculations for bipropellant and solid propellant motors.

Smith, S. D.

1983-01-01

24

An expert system for spectroscopic analysis of rocket engine plumes  

NASA Technical Reports Server (NTRS)

The expert system described in this paper analyzes spectral emissions of rocket engine exhaust plumes and shows major promise for use in engine health diagnostics. Plume emission spectroscopy is an important tool for diagnosing engine anomalies, but it is time-consuming and requires highly skilled personnel. The expert system was created to alleviate such problems. The system accepts a spectral plot in the form of wavelength vs intensity pairs and finds the emission peaks in the spectrum, lists the elemental emitters present in the data and deduces the emitter that produced each peak. The system consists of a conventional language component and a commercially available inference engine that runs on an Apple Macintosh computer. The expert system has undergone limited preliminary testing. It detects elements well and significantly decreases analysis time.

Reese, Greg; Valenti, Elizabeth; Alphonso, Keith; Holladay, Wendy

1991-01-01

25

Exhaust Plume Measurements of the VASIMR VX-200  

NASA Astrophysics Data System (ADS)

Recent progress is discussed in the development of an advanced RF electric propulsion concept: the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) VX-200 engine, a 200 kW flight-technology prototype. Results from high power Helicon only and Helicon with ICRH experiments are performed on the VX-200 using argon plasma. Recent measurements of axial plasma density and potential profiles, magnetic field-line shaping, charge exchange, and force measurements taken in the plume of the VX-200 exhaust are made within a new 125 cubic meter cryo-pumped vacuum chamber and are presented in the context of RF plasma thruster physics.

Longmier, Benjamin; Bering, Edgar, III; Squire, Jared; Glover, Tim; Chang-Diaz, Franklin; Brukardt, Michael

2008-11-01

26

Analysis of the measured effects of the principal exhaust effluents from solid rocket motors  

NASA Technical Reports Server (NTRS)

The feasibility of conducting environmental chamber tests using a small rocket motor to study the physical processes which occur when the exhaust products from solid motors mix with the ambient atmosphere was investigated. Of particular interest was the interaction between hydrogen chloride, aluminum oxide, and water vapor. Several types of instruments for measuring HCl concentrations were evaluated. Under some conditions it was noted that acid aerosols were formed in the ground cloud. These droplets condensed on Al2O3 nuclei and were associated with the rocket exhaust cooling during the period of plume rise to stabilization. Outdoor firings of the solid rocket motors of a 6.4 percent scaled model of the space shuttle were monitored to study the interaction of the exhaust effluents with vegetation downwind of the test site. Data concerning aluminum oxide particles produced by solid rocket motors were evaluated.

Dawbarn, R.; Kinslow, M.; Watson, D. J.

1980-01-01

27

Rocket exhaust ground cloud/atmospheric interactions  

NASA Technical Reports Server (NTRS)

An attempt to identify and minimize the uncertainties and potential inaccuracies of the NASA Multilayer Diffusion Model (MDM) is performed using data from selected Titan 3 launches. The study is based on detailed parametric calculations using the MDM code and a comparative study of several other diffusion models, the NASA measurements, and the MDM. The results are discussed and evaluated. In addition, the physical/chemical processes taking place during the rocket cloud rise are analyzed. The exhaust properties and the deluge water effects are evaluated. A time-dependent model for two aerosol coagulations is developed and documented. Calculations using this model for dry deposition during cloud rise are made. A simple model for calculating physical properties such as temperature and air mass entrainment during cloud rise is also developed and incorporated with the aerosol model.

Hwang, B.; Gould, R. K.

1978-01-01

28

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

PubMed Central

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

2014-01-01

29

Assessment of analytical techniques for predicting solid propellant exhaust plumes and plume impingement environments  

NASA Technical Reports Server (NTRS)

An analysis of experimental nozzle, exhaust plume, and exhaust plume impingement data is presented. The data were obtained for subscale solid propellant motors with propellant Al loadings of 2, 10 and 15% exhausting to simulated altitudes of 50,000, 100,000 and 112,000 ft. Analytical predictions were made using a fully coupled two-phase method of characteristics numerical solution and a technique for defining thermal and pressure environments experienced by bodies immersed in two-phase exhaust plumes.

Tevepaugh, J. A.; Smith, S. D.; Penny, M. M.

1977-01-01

30

Monopropellant thruster exhaust plume contamination measurements  

NASA Technical Reports Server (NTRS)

The potential spacecraft contaminants in the exhaust plume of a 0.89N monopropellant hydrazine thruster were measured in an ultrahigh quartz crystal microbalances located at angles of approximately 0 deg, + 15 deg and + or - 30 deg with respect to the nozzle centerline. The crystal temperatures were controlled such that the mass adhering to the crystal surface at temperatures of from 106 K to 256 K could be measured. Thruster duty cycles of 25 ms on/5 seconds off, 100 ms on/10 seconds off, and 200 ms on/20 seconds off were investigated. The change in contaminant production with thruster life was assessed by subjecting the thruster to a 100,000 pulse aging sequence and comparing the before and after contaminant deposition rates. The results of these tests are summarized, conclusions drawn, and recommendations given.

Baerwald, R. K.; Passamaneck, R. S.

1977-01-01

31

Radiation\\/convection coupling in rocket motors and plumes  

Microsoft Academic Search

The three commonly used propellant systems - H2\\/O2, RP-1\\/O2, and solid propellants - primarily radiate as molecular emitters, non-scattering small particles, and scattering larger particles, respectively. Present technology has accepted the uncoupling of the radiation analysis from that of the flowfield. This approximation becomes increasingly inaccurate as one considers plumes, interior rocket chambers, and nuclear rocket propulsion devices. This study

R. C. Farmer; A. J. Saladino

1993-01-01

32

Numerical study on the influence of aluminum on infrared radiation signature of exhaust plume  

NASA Astrophysics Data System (ADS)

The infrared radiation signature of exhaust plume from solid propellant rockets has been widely mentioned for its important realistic meaning. The content of aluminum powder in the propellants is a key factor that affects the infrared radiation signature of the plume. The related studies are mostly on the conical nozzles. In this paper, the influence of aluminum on the flow field of plume, temperature distribution, and the infrared radiation characteristics were numerically studied with an object of 3D quadrate nozzle. Firstly, the gas phase flow field and gas-solid multi phase flow filed of the exhaust plume were calculated using CFD method. The result indicates that the Al203 particles have significant effect on the flow field of plume. Secondly, the radiation transfer equation was solved by using a discrete coordinate method. The spectral radiation intensity from 1000-2400 cm-1 was obtained. To study the infrared radiation characteristics of exhaust plume, an exceptional quadrate nozzle was employed and much attention was paid to the influences of Al203 particles in solid propellants. The results could dedicate the design of the divert control motor in such hypervelocity interceptors or missiles, or be of certain meaning to the improvement of ingredients of solid propellants.

Zhang, Wei; Ye, Qing-qing; Li, Shi-peng; Wang, Ning-fei

2013-09-01

33

ASRM plume deflector analysis program. [advanced solid rocket motor  

NASA Technical Reports Server (NTRS)

This paper presents analytical conclusions resulting from subscale solid rocket motor tests and flowfield modeling for a plume deflector. Loads, flow characteristics, and corresponding material behavior were predicted or observed and will be used in final design of the deflector. The efforts resulted in quantifiable size reductions and lower cost material selections, which will significantly reduce the deflector cost while meeting performance requirements.

Dawson, Michael C.; Douglas, Freddie, III; Orlin, Peter A.

1992-01-01

34

Wavelength-Agile Optical Sensor for Exhaust Plume and Cryogenic Fluid Interrogation  

NASA Technical Reports Server (NTRS)

Two optical sensors developed in UW-Madison labs were evaluated for their potential to characterize rocket engine exhaust plumes and liquid oxygen (LOX) fluid properties. The plume sensor is based on wavelength-agile absorption spectroscopy A device called a chirped white pulse emitter (CWPE) is used to generate the wavelength agile light, scanning, for example, 1340 - 1560 nm every microsecond. Properties of the gases in the rocket plume (for example temperature and water mole fraction) can be monitored using these wavelength scans. We have performed preliminary tests in static gas cells, a laboratory GOX/GH2 thrust chamber, and a solid-fuel hybrid thrust chamber, and these initial tests demonstrate the potential of the CWPE for monitoring rocket plumes. The LOX sensor uses an alternative to wavelength agile sensing: two independent, fixed-wavelength lasers are combined into a single fiber. One laser is absorbed by LOX and the other not: by monitoring the differential transmission the LOX concentration in cryogenic feed lines can be inferred. The sensor was successful in interrogating static LOX pools in laboratory tests. Even in ice- and bubble-laden cryogenic fluids, LOX concentrations were measured to better than 1% with a 3 microsec time constant.

Sanders, Scott T.; Chiaverini, Martin J.; Gramer, Daniel J.

2004-01-01

35

The effects of the exhaust plume on the lightning triggering conditions for launch vehicles  

NASA Technical Reports Server (NTRS)

Apollo 12 and Atlas Centaur 67 are two launch vehicles that have experienced triggered lightning strikes. Serious consequences resulted from the events; in the case of Atlas Centaur 67, the vehicle and the payload were lost. These events indicate that it is necessary to develop launch rules which would prevent such occurrences. In order to develop valid lightning related rules, it is necessary to understand the effects of the plume. Some have assumed that the plume can be treated as a perfect conductor, and have computed electric field enhancement factors on that basis. The authors have looked at the plume, and believe that these models are not correct, because they ignore the fluid motion of the conducting plates. The authors developed a model which includes this flow character. In this model, the external field is excluded from the plume as it would be for any good conductor, but, in addition, the charge must distribute so that the charge density is zero at some location in the exhaust. When this condition is included in the calculation of triggering enhancement factors, they can be two to three times larger than calculated by other methods which include a conductive plume but don't include the correct boundary conditions. Here, the authors review the relevant features of rocket exhausts for the triggered lightning problem, present an approach for including flowing conductive gases, and present preliminary calculations to demonstrate the effect that the plume has on enhancement factors.

Eriksen, Frederick J.; Rudolph, Terence H.; Perala, Rodney A.

1991-01-01

36

UVPI imaging from the LACE satellite: The Strypi rocket plume  

NASA Astrophysics Data System (ADS)

The Ultraviolet Plume Instrument (UVPI) is a small plume-tracking instrument that was flown on the Naval Research Laboratory's Low-power Atmospheric Compensation Experiment (LACE) satellite. The UVPI plume camera has a narrow field of view (0.180 deg by 0.135 deg), and it observes sources through any of four filters with passbands of 195 to 295 nm, 220 to 320 nm, 235 to 350 nm, and 300 to 320 nm. The Strypi rocket was launched from Hawaii on 18 Feb. 1991. The second (Antares) and the third (Star 27) stages reached 110 km altitude and were successfully detected and tracked by the UVPI from a range of 450 to 550 km. The spectral radiance and spectral radiant intensities of the missile plumes were extracted from these images for the four passbands.

Smathers, H. W.; Horan, D. M.; Cardon, J. G.; Malaret, E. R.; Singh, M.

1993-09-01

37

Atmospheric scavenging of hydrochloric acid. [from rocket exhaust  

NASA Technical Reports Server (NTRS)

The scavenging of hydrogen chloride from a solid rocket exhaust cloud was investigated. Water drops were caused to fall through a confined exhaust cloud and then analyzed to determine the amount of HCl captured during fall. Bubblers were used to measure HCl concentration within the chamber. The measured chamber HCl concentration, together with the measured HCl deposition on the chamber walls, accounted for 81 to 94% of the theoretical HCl. It was found that the amount of HCl captured was approximately one-half of that predicted by the Frossling correlation. No effect of humidity was detected through a range of 69-98% R.H.. The scavenging of HCl from a solid rocket exhaust cloud was calculated using an idealized Kennedy Space Center rain cycle. Results indicate that this cycle would reduce the cloud HCl concentration to 20.6% if its value in the absence of rain.

Knutson, E. O.; Fenton, D. L.

1975-01-01

38

Effects of Rocket Exhaust on Lunar Soil Reflectance Properties  

NASA Astrophysics Data System (ADS)

The Apollo, Surveyor, and Luna spacecraft descent engine plumes affected the regolith at and surrounding their landing sites. Owing to the lack of rapid weathering processes on the Moon, surface alterations are still visible as photometric anomalies in Lunar Reconnaissance Orbiter Camera (LROC) Narrow Angle Camera (NAC) images. These areas are interpreted as disturbance of the regolith by rocket exhaust during descent of the spacecraft, which we refer to as "blast zones" (BZs). The BZs consist of an area of lower reflectance (LR-BZ) compared to the surroundings that extends up to a few meters out from the landers, as well as a broader halo of higher reflectance (HR-BZ) that extends tens to hundreds of meters out from the landers. We use phase-ratio images for each landing site to determine the spatial extent of the disturbed regions and to quantify differences in reflectance and backscattering characteristics within the BZs compared to nearby undisturbed regolith. We also compare the reflectance changes and BZ dimensions at the Apollo sites with those at Luna and Surveyor sites. We seek to determine the effects of rocket exhaust in terms of erosion and particle redistribution, as well as the cause(s) of the reflectance variations, i.e., physical changes at the regolith surface. When approximated as an ellipse, the average Apollo BZ area is ~29,000 m2 (~175 ± 60 m by 200 ± 27 m) which is 10x larger than the average Luna BZ, and over 100x larger than the average Surveyor BZ. Moreover, BZ area scales roughly with lander mass (as a proxy for thrust). The LR-BZs are evident at the Apollo sites, especially where astronaut bioturbation has roughened the soil, leading to a 2-14% reduction in reflectance at ~30° phase. The LR-BZs at the Luna and Surveyor sites are less evident and may be mostly confined to the area below the landers. The average normalized reflectance in the HR-BZs for images with a 30° phase angle is 2-16% higher than in the undisturbed surrounding areas; this magnitude is the same, within uncertainty, for all sites, indicating a common process or combination of processes causing differences in reflectance properties of the regolith. Phase-ratio images and photometric data collected over a range of illumination geometries show that a greater separation in reflectance occurs between the HR-BZs and undisturbed areas with increasing phase angle and indicate that the HR-BZs are less backscattering than undisturbed areas. As working hypotheses, we consider the following possibilities to explain BZ reflectance phenomena: change in macroscopic roughness, microscopic modification of surface structure, redistribution of fines (excavation from LR-BZ and deposition in HR-BZ), change in compaction, contamination from fuel, and modification of maturity. The LR-BZ is affected by macroscopic disruption of the surface and increased shadowing. We infer that HR-BZ reflectance has been affected by scouring from particles entrained by exhaust gases with low-angle trajectories. Entrained particles with trajectories greater than a few degrees relative to horizontal travel well beyond the BZ boundary and do not contribute to BZ reflectance variations. Regolith particle interactions with surface soil within HR-BZs may destroy fine-scale surface structure (e.g., "fairy-castle") and decrease macroscopic roughness, contributing to a decrease in backscattering character within the HR-BZ.

Clegg, R. N.; Jolliff, B. L.; Robinson, M. S.; Hapke, B. W.; Plescia, J. B.

2012-12-01

39

Numerical Simulation of Rocket Exhaust Interaction with Lunar Soil  

NASA Technical Reports Server (NTRS)

This technology development originated from the need to assess the debris threat resulting from soil material erosion induced by landing spacecraft rocket plume impingement on extraterrestrial planetary surfaces. The impact of soil debris was observed to be highly detrimental during NASA s Apollo lunar missions and will pose a threat for any future landings on the Moon, Mars, and other exploration targets. The innovation developed under this program provides a simulation tool that combines modeling of the diverse disciplines of rocket plume impingement gas dynamics, granular soil material liberation, and soil debris particle kinetics into one unified simulation system. The Unified Flow Solver (UFS) developed by CFDRC enabled the efficient, seamless simulation of mixed continuum and rarefied rocket plume flow utilizing a novel direct numerical simulation technique of the Boltzmann gas dynamics equation. The characteristics of the soil granular material response and modeling of the erosion and liberation processes were enabled through novel first principle-based granular mechanics models developed by the University of Florida specifically for the highly irregularly shaped and cohesive lunar regolith material. These tools were integrated into a unique simulation system that accounts for all relevant physics aspects: (1) Modeling of spacecraft rocket plume impingement flow under lunar vacuum environment resulting in a mixed continuum and rarefied flow; (2) Modeling of lunar soil characteristics to capture soil-specific effects of particle size and shape composition, soil layer cohesion and granular flow physics; and (3) Accurate tracking of soil-borne debris particles beginning with aerodynamically driven motion inside the plume to purely ballistic motion in lunar far field conditions. In the earlier project phase of this innovation, the capabilities of the UFS for mixed continuum and rarefied flow situations were validated and demonstrated for lunar lander rocket plume flow impingement under lunar vacuum conditions. Applications and improvements to the granular flow simulation tools contributed by the University of Florida were tested against Earth environment experimental results. Requirements for developing, validating, and demonstrating this solution environment were clearly identified, and an effective second phase execution plan was devised. In this phase, the physics models were refined and fully integrated into a production-oriented simulation tool set. Three-dimensional simulations of Apollo Lunar Excursion Module (LEM) and Altair landers (including full-scale lander geometry) established the practical applicability of the UFS simulation approach and its advanced performance level for large-scale realistic problems.

Liever, Peter; Tosh, Abhijit; Curtis, Jennifer

2012-01-01

40

Thermal radiation model for solid rocket booster plumes  

NASA Technical Reports Server (NTRS)

The Monte Carlo method is used to model the thermal radiation field of the plumes for the dual solid rocket boosters astride the Space Shuttle launch configuration. The model accounts for axial and radial variations in radiative properties of the plumes. The plumes are considered to be composed of a dispersion of aluminum oxide (Al2O3) particles immersed in the gaseous products of combustion. The principal emitting gases are taken to be CO, CO2, H2O, and HCl. The thermal model is based on local thermodynamic equilibrium. Scattering of radiant energy by Al2O3 particles may be treated as isotropic or anisotropic. Sample radiant heating rates to the base region of the Space Shuttle are shown. Space Shuttle geometries are simulated as combinations of quadric surfaces.

Watson, G. H.; Lee, A. L.

1977-01-01

41

Ecological effects and environmental fate of solid rocket exhaust  

NASA Technical Reports Server (NTRS)

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

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

1974-01-01

42

Lidar for remote measurement of ozone in the exhaust plumes of launch vehicles.  

PubMed

Large quantities of chlorine and alumina particles are injected directly into the stratosphere by the current fleet of launch vehicles. Environmental concerns have been raised over the impact of the rocket exhaust on the ozone layer. Recently differential absorption lidar (DIAL) was selected for remote sensing of ozone density within the plumes of Titan IV launch vehicles. The application of DIAL to this very challenging problem is described, and an implementation of UV-ozone DIAL is discussed that holds promise for this application. PMID:21085408

Gelbwachs, J A

1996-05-20

43

Lidar for remote measurement of ozone in the exhaust plumes of launch vehicles  

NASA Astrophysics Data System (ADS)

Large quantities of chlorine and alumina particles are injected directly into the stratosphere by the current fleet of launch vehicles. Environmental concerns have been raised over the impact of the rocket exhaust on the ozone layer. Recently, differential absorption lidar (DIAL) was selected for remote sensing of ozone density within the plumes of Titan IV launch vehicles. The application of DIAL to this very challenging problem is described, and an implementation of UV-ozone DIAL is discussed that holds promise for this application.

Gelbwachs, Jerry A.

1996-05-01

44

Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes  

NASA Astrophysics Data System (ADS)

Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

Kinefuchi, K.; Funaki, I.; Shimada, T.; Abe, T.

2012-10-01

45

Measurement and Characterization of Space Shuttle Solid Rocket Motor Plume Acoustics  

NASA Technical Reports Server (NTRS)

Lift-off acoustic environments generated by the future Ares I launch vehicle are assessed by the NASA Marshall Space Flight Center (MSFC) acoustics team using several prediction tools. This acoustic environment is directly caused by the Ares I First Stage booster, powered by the five-segment Reusable Solid Rocket Motor (RSRMV). The RSRMV is a larger-thrust derivative design from the currently used Space Shuttle solid rocket motor, the Reusable Solid Rocket Motor (RSRM). Lift-off acoustics is an integral part of the composite launch vibration environment affecting the Ares launch vehicle and must be assessed to help generate hardware qualification levels and ensure structural integrity of the vehicle during launch and lift-off. Available prediction tools that use free field noise source spectrums as a starting point for generation of lift-off acoustic environments are described in the monograph NASA SP-8072: "Acoustic Loads Generated by the Propulsion System." This monograph uses a reference database for free field noise source spectrums which consist of subscale rocket motor firings, oriented in horizontal static configurations. The phrase "subscale" is appropriate, since the thrust levels of rockets in the reference database are orders of magnitude lower than the current design thrust for the Ares launch family. Thus, extrapolation is needed to extend the various reference curves to match Ares-scale acoustic levels. This extrapolation process yields a subsequent amount of uncertainty added upon the acoustic environment predictions. As the Ares launch vehicle design schedule progresses, it is important to take every opportunity to lower prediction uncertainty and subsequently increase prediction accuracy. Never before in NASA s history has plume acoustics been measured for large scale solid rocket motors. Approximately twice a year, the RSRM prime vendor, ATK Launch Systems, static fires an assembled RSRM motor in a horizontal configuration at their test facility in Utah. The remaining RSRM static firings will take place on elevated terrain, with the nozzle exit plume being mostly undeflected and the landscape allowing placement of microphones within direct line of sight to the exhaust plume. These measurements will help assess the current extrapolation process by direct comparison between subscale and full scale solid rocket motor data.

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

2009-01-01

46

Radiation/convection coupling in rocket motors and plumes  

NASA Technical Reports Server (NTRS)

The three commonly used propellant systems - H2/O2, RP-1/O2, and solid propellants - primarily radiate as molecular emitters, non-scattering small particles, and scattering larger particles, respectively. Present technology has accepted the uncoupling of the radiation analysis from that of the flowfield. This approximation becomes increasingly inaccurate as one considers plumes, interior rocket chambers, and nuclear rocket propulsion devices. This study will develop a hierarchy of methods which will address radiation/convection coupling in all of the aforementioned propulsion systems. The nature of the radiation/convection coupled problem is that the divergence of the radiative heat flux must be included in the energy equation and that the local, volume-averaged intensity of the radiation must be determined by a solution of the radiative transfer equation (RTE). The intensity is approximated by solving the RTE along several lines of sight (LOS) for each point in the flowfield. Such a procedure is extremely costly; therefore, further approximations are needed. Modified differential approximations are being developed for this purpose. It is not obvious which order of approximations are required for a given rocket motor analysis. Therefore, LOS calculations have been made for typical rocket motor operating conditions in order to select the type approximations required. The results of these radiation calculations, and the interpretation of these intensity predictions are presented herein.

Farmer, R. C.; Saladino, A. J.

1993-01-01

47

ASRM subscale plume deflector testing. [advanced solid rocket motor  

NASA Technical Reports Server (NTRS)

This paper reports the results of the scale model (1/22) testing of candidate refractory materials to be used as surface coatings for a solid rocket motor plume deflector structure. Five ROM tests were conducted to acquire data to support the selection, thickness determination, and placement of the materials. All data acquisition was achieved through nonintrusive methods. The tests demonstrated that little or no reductions in performance of the full-scale deflector would be experienced if the most economical materials were selected for construction.

Douglas, Freddie, III; Dawson, Michael C.; Orlin, Peter A.

1992-01-01

48

Plume spectrometry for liquid rocket engine health monitoring  

NASA Technical Reports Server (NTRS)

An investigation of Space Shuttle Main Engine (SSME) testing failures identified optical events which appeared to be precursors of those failures. A program was therefore undertaken to detect plume trace phenomena characteristic of the engine and to design a monitoring system, responsive to excessive activity in the plume, capable of delivering a warning of an anomalous condition. By sensing the amount of extraneous material entrained in the plume and considering engine history, it may be possible to identify wearing of failing components in time for a safe shutdown and thus prevent a catastrophic event. To investigate the possibilities of safe shutdown and thus prevent a monitor to initiate the shutdown procedure, a large amount of plume data were taken from SSME firings using laboratory instrumentation. Those data were used to design a more specialized instrument dedicated to rocket plume diagnostics. The spectral wavelength range of the baseline data was about 220 nanometers (nm) to 15 micrometer with special attention given to visible and near UV. The data indicates that a satisfactory design will include a polychromator covering the range of 250 nM to 1000 nM, along with a continuous coverage spectrometer, each having a resolution of at least 5A degrees. The concurrent requirements for high resolution and broad coverage are normally at odds with one another in commercial instruments, therefore necessitating the development of special instrumentation. The design of a polychromator is reviewed herein, with a detailed discussion of the continuous coverage spectrometer delayed to a later forum. The program also requires the development of applications software providing detection, variable background discrimination, noise reduction, filtering, and decision making based on varying historical data.

Powers, William T.; Sherrell, F. G.; Bridges, J. H., III; Bratcher, T. W.

1988-01-01

49

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

SciTech Connect

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

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

1993-06-01

50

Plume flowfield analysis of the shuttle primary Reaction Control System (RCS) rocket engine  

NASA Technical Reports Server (NTRS)

A solution was generated for the physical properties of the Shuttle RCS 4000 N (900 lb) rocket engine exhaust plume flowfield. The modeled exhaust gas consists of the five most abundant molecular species, H2, N2, H2O, CO, and CO2. The solution is for a bare RCS engine firing into a vacuum; the only additional hardware surface in the flowfield is a cylinder (=engine mount) which coincides with the nozzle lip outer corner at X = 0, extends to the flowfield outer boundary at X = -137 m and is coaxial with the negative symmetry axis. Continuum gas dynamic methods and the Direct Simulation Monte Carlo (DSMC) method were combined in an iterative procedure to produce a selfconsistent solution. Continuum methods were used in the RCS nozzle and in the plume as far as the P = 0.03 breakdown contour; the DSMC method was used downstream of this continuum flow boundary. The DSMC flowfield extends beyond 100 m from the nozzle exit and thus the solution includes the farfield flow properties, but substantial information is developed on lip flow dynamics and thus results are also presented for the flow properties in the vicinity of the nozzle lip.

Hueser, J. E.; Brock, F. J.

1990-01-01

51

Nuclear thermal rocket plume interactions with spacecraft. Final report  

SciTech Connect

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

Mauk, B.H. [Johns Hopkins Univ., Laurel, MD (United States); Gatsonis, N.A.; Buzby, J.; Yin, X. [Worcester Polytechnic Inst., MA (United States). Mechanical Engineering Department

1997-05-01

52

Development and Validation of a Computational Model for Predicting the Behavior of Plumes from Large Solid Rocket Motors  

NASA Technical Reports Server (NTRS)

Exhaust plumes from large solid rocket motors fired at ATK's Promontory test site carry particulates to high altitudes and typically produce deposits that fall on regions downwind of the test area. As populations and communities near the test facility grow, ATK has become increasingly concerned about the impact of motor testing on those surrounding communities. To assess the potential impact of motor testing on the community and to identify feasible mitigation strategies, it is essential to have a tool capable of predicting plume behavior downrange of the test stand. A software package, called PlumeTracker, has been developed and validated at ATK for this purpose. The code is a point model that offers a time-dependent, physics-based description of plume transport and precipitation. The code can utilize either measured or forecasted weather data to generate plume predictions. Next-Generation Radar (NEXRAD) data and field observations from twenty-three historical motor test fires at Promontory were collected to test the predictive capability of PlumeTracker. Model predictions for plume trajectories and deposition fields were found to correlate well with the collected dataset.

Wells, Jason E.; Black, David L.; Taylor, Casey L.

2013-01-01

53

Effects of rocket exhaust products in the thermosphere and ionsphere  

SciTech Connect

This paper reviews the current state of understanding of the problem of ionospheric F-layer depletions produced by chemical effects of the exhaust gases from large rockets, with particular emphasis on the Heavy Lift Launch Vehicles (HLLV) proposed for use in the construction of solar power satellites. The currently planned HLLV flight profile calls for main second-stage propulsion confined to altitudes below 124 km, and a brief orbit circularization maneuver at apogee. The second stage engines deposit 9 x 10/sup 31/ H/sub 2/O and H/sub 2/ molecules between 74 and 124 km. Model computations show that they diffuse gradually into the ionospheric F region, where they lead to weak but widespread and persistent depletions of ionization and continuous production of H atoms. The orbit circularization burn deposits 9 x 10/sup 29/ exhaust molecules at about 480-km altitude. These react rapidly with the F2 region 0/sup +/ ions, leading to a substantial (factor-of-three) reduction in plasma density, which extends over a 1000- by 2000-km region and persists for four to five hours. For purposes of computer model verification, a computation is included representing the Skylab I launch, for which observational data exist. The computations and data are compared, and the computer model is described.

Zinn, J.; Sutherland, C.D.

1980-02-01

54

Calculation of Free-Atom Fractions in Hydrocarbon-Fueled Rocket Engine Plume  

NASA Technical Reports Server (NTRS)

Free atom fractions (Beta) of nine elements are calculated in the exhaust plume of CH4- oxygen and RP-1-oxygen fueled rocket engines using free energy minimization method. The Chemical Equilibrium and Applications (CEA) computer program developed by the Glenn Research Center, NASA is used for this purpose. Data on variation of Beta in both fuels as a function of temperature (1600 K - 3100 K) and oxygen to fuel ratios (1.75 to 2.25 by weight) is presented in both tabular and graphical forms. Recommendation is made for the Beta value for a tenth element, Palladium. The CEA computer code was also run to compare with experimentally determined Beta values reported in literature for some of these elements. A reasonable agreement, within a factor of three, between the calculated and reported values is observed. Values reported in this work will be used as a first approximation for pilot rocket engine testing studies at the Stennis Space Center for at least six elements Al, Ca, Cr, Cu, Fe and Ni - until experimental values are generated. The current estimates will be improved when more complete thermodynamic data on the remaining four elements Ag, Co, Mn and Pd are added to the database. A critique of the CEA code is also included.

Verma, Satyajit

2006-01-01

55

Effect of contamination on the optical properties of transmitting and reflecting materials exposed to a MMH/N2O4 rocket exhaust  

NASA Technical Reports Server (NTRS)

The changes are presented in spectral transmittance, and reflectance due to exposure of various optical materials to the exhaust plume of a 5-pound thrust bipropellant rocket. The engine was fired in a pulsed mode for a total exposure of 223.7 second. Spectral optical properties were measured in air before and after exposure to the exhaust plume in vacuum. The contaminating layer resulted in both absorption and scattering effects which caused changes as large as 30-50% for transmitting elements and 15% for mirrors in the near ultraviolet wavelengths. The changes in spectral properties of materials exposed to the exhaust plume for 44 and 223.7 seconds are compared and found to be similar.

Bowman, R. L.; Spisz, E. W.; Jack, J. R.

1973-01-01

56

Active chlorine and nitric oxide formation from chemical rocket plume afterburning  

NASA Technical Reports Server (NTRS)

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

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

1994-01-01

57

Laser optogalvanic spectroscopy of neon in a discharge plasma and modeling and analysis of rocket plume RF-line emissions  

NASA Astrophysics Data System (ADS)

The Optogalvanic Effect (OGE) of neon in a hollow cathode discharge lamp has been investigated both experimentally and theoretically. A tunable dye laser was tuned to several 1si -- 2pj neon transitions and the associated time--resolved optogalvanic (OG) spectral waveforms recorded corresponding to the DeltaJ = DeltaK = 0, +/-1 selection rules and modeled using a semi-empirical model. Decay rate constants, amplitudes and the instrumentation time constants were recorded following a good least-squares fit (between the experimental and the theoretical OG data) using the Monte Carlo technique and utilizing both the search and random walk methods. Dominant physical processes responsible for the optogalvanic effect have been analyzed, and the corresponding populations of the laser-excited level and collisional excited levels determined. The behavior of the optogalvanic signal waveform as a function of time, together with the decay rate constants as a function of the discharge current and the instrumentation time constant as a function of current have been studied in detail. The decay times of the OG signals and the population redistributions were also determined. Fairly linear relationships between the decay rate constant and the discharge current, as well as between the instrumental time constant and the discharge current, have been observed. The decay times and the electron collisional rate parameters of the 1s levels involved in the OG transitions have been obtained with accuracy. The excitation temperature of the discharge for neon transitions grouped with the same 1s level have been determined and found to be fairly constant for the neon transitions studied. The experimental optogalvanic effort in the visible region of the electromagnetic spectrum has been complemented by a computation-intensive modeling investigation of rocket plumes in the microwave region. Radio frequency lines of each of the plume species identified were archived utilizing the HITRAN and other databases (e.g. JPL/NASA and Cologne), together with other appropriate spectroscopic data. Hydrazine fuel was selected as the rocket propellant of choice and the plume codes were run by the JHU-APL research group. A representative monopropellant hydrazine plume has been determined to provide exhaust temperature, pressure, velocity, and species number density inputs for model development. A MATLAB code has been developed for computing broadside line-of-sight (LOS) intensities due to line emissions involving ammonia and other plume species. Initially, we assumed Local Thermodynamic Equilibrium (LTE) and included self-absorption contributions due to plume opacity, together with collisional and Doppler broadening, as well as the Doppler shift due to the plume radial velocity towards and away from a stationary detector. The recorded code output was MATLAB coded and an assortment of plume parameters computed, such as the volume emission rate, the absorption coefficient, optical depth and species radiance line-by-line. These parameters were computed both manually utilizing a spread sheet and then automated using the Matlab code. The volume emissions, along with other plume properties, were plotted as a function of the axial distance in the plume for several Radio Frequency (RF) transitions involving various significant plume species. Plume properties, such as the temperature, pressure, number density, and plume particulate speed emanating from the nozzle where analyzed and modeled as the plume drifts away from the rocket nozzle. Both the axial and radial distance dependences were investigated with respect to the various plume properties and parameters. Population distribution of the species (number density) dependence on the plume temperature was investigated and modeled line-by-line for each of the plume species studied at the nozzle exit plane and beyond. In addition, volume emission and absorption coefficients have been analyzed and modeled and solutions to the Radiative Transfer Equation (RTE) applied line-by-line and the radiance determined accurately in the micro

Ogungbemi, Kayode I.

58

Atmospheric measurements of the physical evolution of aircraft exhaust plumes.  

PubMed

Drawing from a series of field measurement activities including the Alternative Aviation Fuels Experiments (AAFEX1 and AAFEX2), we present experimental measurements of particle number, size, and composition-resolved mass that describe the physical and chemical evolution of aircraft exhaust plumes on the time scale of 5 s to 2-3 min. As the plume ages, the particle number emission index initially increases by a factor of 10-50, due to gas-to-particle formation of a nucleation/growth mode, and then begins to fall with increased aging. Increasing the fuel sulfur content causes the initial increase to occur more rapidly. The contribution of the nucleation/growth mode to the overall particle number density is most pronounced at idle power and decreases with increasing engine power. Increasing fuel sulfur content, but not fuel aromatic content causes the nucleation/growth mode to dominate the particle number emissions at higher powers than for a fuel with "normal" sulfur and aromatic content. Particle size measurements indicate that the observed particle number emissions trends are due to continuing gas-to-particle conversion and coagulation growth of the nucleation/growth mode particles, processes which simultaneously increase particle mass and reduce particle number density. Measurements of nucleation/growth mode mass are consistent with the interpretation of particle number and size data and suggest that engine exit plane measurements may underestimate the total particle mass by as much as a factor of between 5 and 10. PMID:23356965

Timko, M T; Fortner, E; Franklin, J; Yu, Z; Wong, H-W; Onasch, T B; Miake-Lye, R C; Herndon, S C

2013-04-01

59

Assessment of analytical and experimental techniques utilized in conducting plume technology tests 575 and 593. [exhaust flow simulation (wind tunnel tests) of scale model Space Shuttle Orbiter  

NASA Technical Reports Server (NTRS)

Since exhaust plumes affect vehicle base environment (pressure and heat loads) and the orbiter vehicle aerodynamic control surface effectiveness, an intensive program involving detailed analytical and experimental investigations of the exhaust plume/vehicle interaction was undertaken as a pertinent part of the overall space shuttle development program. The program, called the Plume Technology program, has as its objective the determination of the criteria for simulating rocket engine (in particular, space shuttle propulsion system) plume-induced aerodynamic effects in a wind tunnel environment. The comprehensive experimental program was conducted using test facilities at NASA's Marshall Space Flight Center and Ames Research Center. A post-test examination of some of the experimental results obtained from NASA-MSFC's 14 x 14-inch trisonic wind tunnel is presented. A description is given of the test facility, simulant gas supply system, nozzle hardware, test procedure and test matrix. Analysis of exhaust plume flow fields and comparison of analytical and experimental exhaust plume data are presented.

Baker, L. R.; Sulyma, P. R.; Tevepaugh, J. A.; Penny, M. M.

1976-01-01

60

Detection of massive negative chemiions in the exhaust plume of a jet aircraft in flight  

NASA Astrophysics Data System (ADS)

Gaseous negative ions were mass spectrometrically measured in the exhaust plume of a jet aircraft in flight. Using a quadrupole mass spectrometer operated in a high-pass mode, it was found that by far most of the ions had mass numbers > 450 amu (atomic mass units) and number densities which markedly exceeded the number densities of ambient atmospheric ions. The latter were observed outside the exhaust plume and had mostly mass numbers < 200 amu. Both their large numbers and large concentrations strongly suggest that the massive ions observed inside the plume are chemiions which were produced by the jet engines. The low fuel sulfur content (22 µg/g) suggests that the massive ions consist at least partly of species other than sulfuric acid. By interaction with exhaust gases these chemiions experienced rapid chemical transformation and growth in the early exhaust plume already at plume ages < 0.4 s.

Arnold, F.; Curtius, J.; Sierau, B.; Bürger, V.; Busen, R.; Schumann, U.

61

In situ plume radiance measurements from the bow shock ultraviolet 2 rocket flight  

Microsoft Academic Search

The ultraviolet spectrum (200-400 nm) of the plumes generated by the second- and third-stage engines of a Strypi XI rocket and of the Mach 17 re-entry bow shock were obtained by a sounding rocket experiment launched from the Barking Sands Research Range (Kauai, Hawaii) on February 18, 1991 at 14:30 GMT. The re-entry optical data were obtained as the payload

Peter W. Erdman; Edward C. Zipf; Patrick Espy; Carl Howlett; Carol Christou; Deborah A. Levin; Robert J. Collins; Graham V. Candler

1993-01-01

62

UVPI imaging from the LACE satellite: The Strypi rocket plume  

Microsoft Academic Search

The Ultraviolet Plume Instrument (UVPI) is a small plume-tracking instrument that was flown on the Naval Research Laboratory's Low-power Atmospheric Compensation Experiment (LACE) satellite. The UVPI plume camera has a narrow field of view (0.180 deg by 0.135 deg), and it observes sources through any of four filters with passbands of 195 to 295 nm, 220 to 320 nm, 235

H. W. Smathers; D. M. Horan; J. G. Cardon; E. R. Malaret; M. Singh

1993-01-01

63

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

NASA Technical Reports Server (NTRS)

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

Everson, John; Nelson, H. F.

1993-01-01

64

Effect of bipropellant plume exhaust effluents on spaceborne optical instruments  

Microsoft Academic Search

Analytical tools together with a good data base are necessary to predict the transport of plume contaminants and their effects on spacecraft surfaces. The present paper describes an assessment of bipropellant thrusters, the production and transport of plume contaminants from these thrusters, and the use of the JPL contamination analysis program to assess the effects of plume contamination on the

C. R. Maag; J. A. Jeffery; J. M. Millard

1980-01-01

65

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

NASA Technical Reports Server (NTRS)

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

Susko, M.

1979-01-01

66

Particulate emissions in the exhaust plume from commercial jet aircraft under cruise conditions  

Microsoft Academic Search

In situ measurements of total concentration, size distribution, and hydration properties of jet engine exhaust from a range of commercial transports is reported. Significant concentration enhancements (above ambient background) for aircraft exhaust particulates is reported permitting the detection of not only newly formed but also aged plumes, even in the presence of considerable ambient pollution. Two types of particle size

D. E. Hagen; P. D. Whitefield; H. Schlager

1996-01-01

67

Hot rocket plume experiment - Survey and conceptual design. [of rhenium-iridium bipropellants  

NASA Technical Reports Server (NTRS)

Attention is given to a space-borne engine plume experiment study to fly an experiment which will both verify and quantify the reduced contamination from advanced rhenium-iridium earth-storable bipropellant rockets (hot rockets) and provide a correlation between high-fidelity, in-space measurements and theoretical plume and surface contamination models. The experiment conceptual design is based on survey results from plume and contamination technologists throughout the U.S. With respect to shuttle use, cursory investigations validate Hitchhiker availability and adaptability, adequate remote manipulator system (RMS) articulation and dynamic capability, acceptable RMS attachment capability, adequate power and telemetry capability, and adequate flight altitude and attitude/orbital capability.

Millard, Jerry M.; Luan, Taylor W.; Dowdy, Mack W.

1992-01-01

68

Recent Advances in Studies of Ionospheric Modification Using Rocket Exhaust (Invited)  

Microsoft Academic Search

Rocket exhaust interacts with the ionosphere to produce a wide range of disturbances. A ten second burn of the Orbital Maneuver Subsystem (OMS) engines on the Space Shuttle deposits over 1 Giga Joule of energy into the upper atmosphere. The exhaust vapors travel at speeds between 4.7 and 10.7 km\\/s coupling momentum into the ions by both collisions and charge

P. A. Bernhardt

2009-01-01

69

Effect of bipropellant plume exhaust effluents on spaceborne optical instruments  

NASA Astrophysics Data System (ADS)

Analytical tools together with a good data base are necessary to predict the transport of plume contaminants and their effects on spacecraft surfaces. The present paper describes an assessment of bipropellant thrusters, the production and transport of plume contaminants from these thrusters, and the use of the JPL contamination analysis program to assess the effects of plume contamination on the Galileo spacecraft. It is shown that, in the case of the Galileo mission, contamination from the liquid engines has been effectively reduced to nothing by the use of predictive tools. Plume shields together with precise scan platform stowage have been designed to protect the optical instruments.

Maag, C. R.; Jeffery, J. A.; Millard, J. M.

1980-01-01

70

UV, VISIBLE, AND INFRARED SPECTRAL EMISSIONS IN HYBRID ROCKET PLUMES  

Microsoft Academic Search

A survey was made of the spectral emissions from a 2 x 10 inch labscale hybrid rocket motor system. The emissions in the Ultraviolet-Visible (300-750 nm), Near Infrared (750-1100 nm), and Mid Infrared (2-16 ?m) regions were studied. Baseline emissions were found to consist of the sodium and potassium atomic lines, present due to the use of silica phenolic insulators,

M. Keith Hudson; Robert B. Shanks; Dallas H. Snider; Diana M. Lindquist; Chris Luchini; Sterling Rooke

71

Ecological effects and environmental fate of solid rocket exhaust. Annual report  

Microsoft Academic Search

Studies on the chemical, chemical-physical, and biological reactions and the toxic effects of solid rocket emissions within selected ecosystems at Kennedy Space Center cover laboratory conditions of exhaust and field studies of natural plant populations in three diverse ecosystems. The latter failed to reveal any structural damage at the concentration levels tested (5 to 100 ppM HCl). The effect, if

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

1974-01-01

72

Program listing for the REEDM (Rocket Exhaust Effluent Diffusion Model) computer program  

NASA Technical Reports Server (NTRS)

The program listing for the REEDM Computer Program is provided. A mathematical description of the atmospheric dispersion models, cloud-rise models, and other formulas used in the REEDM model; vehicle and source parameters, other pertinent physical properties of the rocket exhaust cloud and meteorological layering techniques; user's instructions for the REEDM computer program; and worked example problems are contained in NASA CR-3646.

Bjorklund, J. R.; Dumbauld, R. K.; Cheney, C. S.; Geary, H. V.

1982-01-01

73

Measurement and Characterization of Space Shuttle Solid Rocket Motor Plume Acoustics  

NASA Technical Reports Server (NTRS)

NASA's current models to predict lift-off acoustics for launch vehicles are currently being updated using several numerical and empirical inputs. One empirical input comes from free-field acoustic data measured at three Space Shuttle Reusable Solid Rocket Motor (RSRM) static firings. The measurements were collected by a joint collaboration between NASA - Marshall Space Flight Center, Wyle Labs, and ATK Launch Systems. For the first time NASA measured large-thrust solid rocket motor plume acoustics for evaluation of both noise sources and acoustic radiation properties. Over sixty acoustic free-field measurements were taken over the three static firings to support evaluation of acoustic radiation near the rocket plume, far-field acoustic radiation patterns, plume acoustic power efficiencies, and apparent noise source locations within the plume. At approximately 67 m off nozzle centerline and 70 m downstream of the nozzle exit plan, the measured overall sound pressure level of the RSRM was 155 dB. Peak overall levels in the far field were over 140 dB at 300 m and 50-deg off of the RSRM thrust centerline. The successful collaboration has yielded valuable data that are being implemented into NASA's lift-off acoustic models, which will then be used to update predictions for Ares I and Ares V liftoff acoustic environments.

Kenny, Robert Jeremy

2009-01-01

74

A computer simulation of the afterburning processes occurring within solid rocket motor plumes in the troposphere  

NASA Technical Reports Server (NTRS)

As part of a continuing study of the environmental effects of solid rocket motor (SRM) operations in the troposphere, a numerical model was used to simulate the afterburning processes occurring in solid rocket motor plumes and to predict the quantities of potentially harmful chemical species which are created. The calculations include the effects of finite-rate chemistry and turbulent mixing. It is found that the amount of NO produced is much less than the amount of HCl present in the plume, that chlorine will appear predominantly in the form of HCl although some molecular chlorine is present, and that combustion is complete as is evident from the predominance of carbon dioxide over carbon monoxide.

Gomberg, R. I.; Stewart, R. B.

1976-01-01

75

Use of a Microphone Phased Array to Determine Noise Sources in a Rocket Plume  

NASA Technical Reports Server (NTRS)

A 70-element microphone phased array was used to identify noise sources in the plume of a solid rocket motor. An environment chamber was built and other precautions were taken to protect the sensitive condenser microphones from rain, thunderstorms and other environmental elements during prolonged stay in the outdoor test stand. A camera mounted at the center of the array was used to photograph the plume. In the first phase of the study the array was placed in an anechoic chamber for calibration, and validation of the indigenous Matlab(R) based beamform software. It was found that the "advanced" beamform methods, such as CLEAN-SC was partially successful in identifying speaker sources placed closer than the Rayleigh criteria. To participate in the field test all equipments were shipped to NASA Marshal Space Flight Center, where the elements of the array hardware were rebuilt around the test stand. The sensitive amplifiers and the data acquisition hardware were placed in a safe basement, and 100m long cables were used to connect the microphones, Kulites and the camera. The array chamber and the microphones were found to withstand the environmental elements as well as the shaking from the rocket plume generated noise. The beamform map was superimposed on a photo of the rocket plume to readily identify the source distribution. It was found that the plume made an exceptionally long, >30 diameter, noise source over a large frequency range. The shock pattern created spatial modulation of the noise source. Interestingly, the concrete pad of the horizontal test stand was found to be a good acoustic reflector: the beamform map showed two distinct source distributions- the plume and its reflection on the pad. The array was found to be most effective in the frequency range of 2kHz to 10kHz. As expected, the classical beamform method excessively smeared the noise sources at lower frequencies and produced excessive side-lobes at higher frequencies. The "advanced" beamform routine CLEAN-SC created a series of lumped sources which may be unphysical. We believe that the present effort is the first-ever attempt to directly measure noise source distribution in a rocket plume.

Panda, J.; Mosher, R.

2010-01-01

76

Passive ranging of dynamic rocket plumes using infrared and visible oxygen attenuation  

NASA Astrophysics Data System (ADS)

Atmospheric oxygen absorption bands in observed spectra of boost phase missiles can be used to accurately estimate range from sensor to target. One method is to compare observed values of band averaged absorption to radiative transfer models. This is most effective using bands where there is a single absorbing species. This work compares spectral attenuation of two oxygen absorption bands in the near-infrared (NIR) and visible (Vis) spectrum, centered at 762 nm and 690 nm, to passively determine range. Spectra were observed from a static test of a full-scale solid rocket motor at a 900m range. The NIR O2 band provided range estimates accurate to within 3%, while the Vis O2 band had a range error of 15%. A Falcon 9 rocket launch at an initial range of 13km was also tracked and observed for 90 seconds after ignition. The NIR O2 band provided in-flight range estimates accurate to within 2% error for the first 30 seconds of tracked observation. The Vis O2 band also provided accurate range estimates with an error of approximately 4%. Rocket plumes are expected to be significantly brighter at longer wavelengths, but absorption in the NIR band is nearly ten times stronger than the Vis band, causing saturation at shorter path lengths. An atmospheric band is considered saturated when all the in-band frequencies emitted from the rocket plume are absorbed before reaching the sensor.

Vincent, R. Anthony; Hawks, Michael R.

2011-05-01

77

Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines: 1. Heterogeneous condensation of combustion products  

NASA Astrophysics Data System (ADS)

Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines during last stages of Proton, Molniya, and Start launchers operating in the upper atmospheric with different types of fuels is considered. Particle heating is taken into account with emission of latent heat of condensation and energy loss due to radiation and heat exchange with combustion products. Using the solution of the heat balance and condensed particle mass equations, the temporal change in the temperature and thickness of the condensate layer is obtained. Practically, no condensation of water vapor and carbon dioxide in the jet exhaust of a Start launcher occurs. In plumes of Proton and Molniya launchers, the condensation of water vapor and carbon dioxide can start at distances of 120-170 m and 450-650 m from the engine nozzle, respectively. In the course of condensation, the thickness of the "water" layer on particles can exceed 100 Å, and the thickness of carbon dioxide can exceed 60 Å.

Platov, Yu. V.; Semenov, A. I.; Filippov, B. V.

2014-01-01

78

Stratospheric aircraft exhaust plume and wake chemistry studies  

NASA Astrophysics Data System (ADS)

This report documents progress to date in an ongoing study to analyze and model emissions leaving a proposed High Speed Civil Transport (HSCT) from when the exhaust gases leave the engine until they are deposited at atmospheric scales in the stratosphere. Estimates are given for the emissions, summarizing relevant earlier work (CIAP) and reviewing current propulsion research efforts. The chemical evolution and the mixing and vortical motion of the exhaust are analyzed to track the exhaust and its speciation as the emissions are mixed to atmospheric scales. The species tracked include those that could be heterogeneously reactive on the surfaces of the condensed solid water (ice) particles and on exhaust soot particle surfaces. Dispersion and reaction of chemical constituents in the far wake are studied with a Lagrangian air parcel model, in conjunction with a radiation code to calculate the net heating/cooling. Laboratory measurements of heterogeneous chemistry of aqueous sulfuric acid and nitric acid hydrates are also described. Results include the solubility of HCl in sulfuric acid which is a key parameter for modeling stratospheric processing. We also report initial results for condensation of nitric acid trihydrate from gas phase H2O and HNO3.

Miake-Lye, R. C.; Martinez-Sanchez, M.; Brown, R. C.; Kolb, C. E.; Worsnop, D. R.; Zahniser, M. S.; Robinson, G. N.; Rodriguez, J. M.; Ko, M. K. W.; Shia, R.-L.

1992-10-01

79

Factors to Consider in Designing Aerosol Inlet Systems for Engine Exhaust Plume Sampling  

NASA Technical Reports Server (NTRS)

This document consists of viewgraphs of charts and diagrams of considerations to take when sampling the engine exhaust plume. It includes a chart that compares the emissions from various fuels, a diagram and charts of the various processes and conditions that influence the particulate size and concentration,

Anderson, Bruce

2004-01-01

80

Computational models for the analysis of three-dimensional internal and exhaust plume flowfields  

NASA Technical Reports Server (NTRS)

This paper describes computational procedures developed for the analysis of three-dimensional supersonic ducted flows and multinozzle exhaust plume flowfields. The models/codes embodying these procedures cater to a broad spectrum of geometric situations via the use of multiple reference plane grid networks in several coordinate systems. Shock capturing techniques are employed to trace the propagation and interaction of multiple shock surfaces while the plume interface, separating the exhaust and external flows, and the plume external shock are discretely analyzed. The computational grid within the reference planes follows the trace of streamlines to facilitate the incorporation of finite-rate chemistry and viscous computational capabilities. Exhaust gas properties consist of combustion products in chemical equilibrium. The computational accuracy of the models/codes is assessed via comparisons with exact solutions, results of other codes and experimental data. Results are presented for the flows in two-dimensional convergent and divergent ducts, expansive and compressive corner flows, flow in a rectangular nozzle and the plume flowfields for exhausts issuing out of single and multiple rectangular nozzles.

Dash, S. M.; Delguidice, P. D.

1977-01-01

81

Calculation of exhaust plume structure and emissions of the ER 2 aircraft in the stratosphere  

NASA Astrophysics Data System (ADS)

Calculations are presented for the NASA ER 2 aircraft at high altitude to simulate measurements taken of its own emissions during a wake-crossing event. Results are presented for a Mach = 0.71 case in the lower stratosphere with an engine NOx emission index of 4.6 corresponding to the measured value. A series of codes was used in the analysis to calculate the flow field and chemical kinetics, from the engine combustor out to a distance of about 20.2 km (97 s). Initial plume properties were calculated with a two-dimensional computational fluid dynamics (CFD) code with finite rate chemistry. The results of the plume code initialized a three-dimensional parabolized Navier-Stokes (PNS) reacting flow solution, where the plume dynamics interacting with the aircraft wake were calculated out to the region of plume breakup. Results show that the early shape and mixing rate of the engine exhaust plume are dominated by the presence of the aircraft vortex wake. Model results for NOY emissions compare well to in situ measurements taken in the field. Calculated exhaust species evolutions predict several species ratios in good agreement with field data. The mixing rate of the engine plume was also predicted to be consistent with dilution measured in the field.

Anderson, M. R.; Miake-Lye, R. C.; Brown, R. C.; Kolb, C. E.

1996-02-01

82

In situ plume radiance measurements from the bow shock ultraviolet 2 rocket flight  

NASA Astrophysics Data System (ADS)

The ultraviolet spectrum (200-400 nm) of the plumes generated by the second- and third-stage engines of a Strypi XI rocket and of the Mach 17 re-entry bow shock were obtained by a sounding rocket experiment launched from the Barking Sands Research Range (Kauai, Hawaii) on February 18, 1991 at 14:30 GMT. The re-entry optical data were obtained as the payload descended from 120 to 65 km with a vehicle velocity of 5.1 km/s. The intensities of the vacuum ultraviolet resonance radiation emitted by atomic oxygen and hydrogen in the bow shock at 130.4 and 121.5 nm, respectively, were also measured. Complementary Langmuir probe measurements provided data on the total plasma density and electron temperature in the boundary layer.

Erdman, Peter W.; Zipf, Edward C.; Espy, Patrick; Howlett, Carl; Christou, Carol; Levin, Deborah A.; Collins, Robert J.; Candler, Graham V.

1993-10-01

83

Exhausted Plume Flow Field Prediction Near the Afterbody of Hypersonic Flight Vehicles in High Altitudes  

NASA Technical Reports Server (NTRS)

A two-dimensional computer code to solve the Burnett equations has been developed which computes the flow interaction between an exhausted plume and hypersonic external flow near the afterbody of a flight vehicle. This Burnett-2D code extends the capability of Navier-Stokes solver (RPLUS2D code) to include high-order Burnett source terms and slip-wall conditions for velocity and temperature. Higher-order Burnett viscous stress and heat flux terms are discretized using central-differencing and treated as source terms. Blocking logic is adopted in order to overcome the difficulty of grid generation. The computation of exhaust plume flow field is divided into two steps. In the first step, the thruster nozzle exit conditions are computed which generates inflow conditions in the base area near the afterbody. Results demonstrated that at high altitudes, the computations of nozzle exit conditions must include the effects of base flow since significant expansion exists in the base region. In the second step, Burnett equations were solved for exhaust plume flow field near the afterbody. The free stream conditions are set at an altitude equal to 80km and the Mach number is equal to 5.0. The preliminary results show that the plume expansion, as altitude increases, will eventually cause upstream flow separation.

Chou, Lynn Chen; Mach, Kervyn D.; Deng, Zheng-Tao; Liaw, Goang-Shin

1995-01-01

84

Computational models for the viscous/inviscid analysis of jet aircraft exhaust plumes. [predicting afterbody drag  

NASA Technical Reports Server (NTRS)

Computational models which analyze viscous/inviscid flow processes in jet aircraft exhaust plumes are discussed. These models are component parts of an NASA-LaRC method for the prediction of nozzle afterbody drag. Inviscid/shock processes are analyzed by the SCIPAC code which is a compact version of a generalized shock capturing, inviscid plume code (SCIPPY). The SCIPAC code analyzes underexpanded jet exhaust gas mixtures with a self-contained thermodynamic package for hydrocarbon exhaust products and air. A detailed and automated treatment of the embedded subsonic zones behind Mach discs is provided in this analysis. Mixing processes along the plume interface are analyzed by two upgraded versions of an overlaid, turbulent mixing code (BOAT) developed previously for calculating nearfield jet entrainment. The BOATAC program is a frozen chemistry version of BOAT containing the aircraft thermodynamic package as SCIPAC; BOATAB is an afterburning version with a self-contained aircraft (hydrocarbon/air) finite-rate chemistry package. The coupling of viscous and inviscid flow processes is achieved by an overlaid procedure with interactive effects accounted for by a displacement thickness type correction to the inviscid plume interface.

Dash, S. M.; Pergament, H. S.; Thorpe, R. D.

1980-01-01

85

An overview of in-flight plume diagnostics for rocket engines  

NASA Technical Reports Server (NTRS)

An overview and progress report of the work performed or sponsored by LeRC toward the development of in-flight plume spectroscopy technology for health and performance monitoring of liquid propellant rocket engines are presented. The primary objective of this effort is to develop technology that can be utilized on any flight engine. This technology will be validated by a hardware demonstration of a system capable of being retrofitted onto the Space Shuttle Main Engines for spectroscopic measurements during flight. The philosophy on system definition and status on the development of instrumentation, optics, and signal processing with respect to implementation on a flight engine are discussed.

Madzsar, G. C.; Bickford, R. L.; Duncan, D. B.

1992-01-01

86

Space Shuttle Solid Rocket Motor Plume Pressure and Heat Rate Measurements  

NASA Technical Reports Server (NTRS)

The Solid Rocket Booster (SRB) Main Flame Deflector (MFD) at Launch Complex 39A was instrumented with sensors to measure heat rates, pressures, and temperatures on the last three Space Shuttle launches. Because the SRB plume is hot and erosive, a robust Tungsten Piston Calorimeter was developed to compliment the measurements made by off-the-shelf sensors. Witness materials were installed and their melting and erosion response to the Mach 2 / 4500 F / 4-second duration plume was observed. The data show that the specification document used for the design of the MFD thermal protection system over-predicted heat rates by a factor of 3 and under-predicted pressures by a factor of 2. These findings will be used to baseline NASA Computational Fluid Dynamics models and develop innovative MFD designs for the Space Launch System (SLS) before this vehicle becomes operational in 2017.

vonEckroth, Wulf; Struchen, Leah; Trovillion, Tom; Perez, Ravael; Nereolich, Shaun; Parlier, Chris

2012-01-01

87

A fast sampling device for the mass spectrometric analysis of liquid rocket engine exhaust  

NASA Technical Reports Server (NTRS)

The design of a device to obtain compositional data on rocket exhaust by direct sampling of reactive flow exhausts into a mass spectrometer is presented. Sampling at three stages differing in pressure and orifice angle and diameter is possible. Results of calibration with pure gases and gas mixtures are erratic and of unknown accuracy for H2, limiting the usefulness of the apparatus for determining oxidizer/fuel ratios from combustion product analysis. Deposition effects are discussed, and data obtained from rocket exhaust spectra are analyzed to give O/F ratios and mixture ratio distribution. The O/F ratio determined spectrometrically is insufficiently accurate for quantitative comparison with cold flow data. However, a criterion for operating conditions with improved mixing of fuel and oxidizer which is consistent with cold flow results may be obtained by inspection of contour plots. A chemical inefficiency in the combustion process when oxidizer is in excess is observed from reactive flow measurements. Present results were obtained with N2O4/N2H4 propellants.

Ryason, P. R.

1975-01-01

88

Characterization of rocket propellant combustion products: Description of sampling and analysis methods for rocket exhaust characterization studies  

SciTech Connect

A systematic approach has been developed and experimentally validated for the sampling and chemical characterization of the rocket motor exhaust generated from the firing of scaled down test motors at the US Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama. The overall strategy was to sample and analyze major exhaust constituents in near real time, while performing off-site analyses of samples collected for the determination of trace constituents of the particulate and vapor phases. Initial interference studies were performed using atmospheric pressure burns of 1 g quantities of propellants in small chambers at Oak Ridge National Laboratory. Carbon monoxide and carbon dioxide were determined using non-dispersive infrared instrumentation. Hydrogen cyanide, hydrogen chloride, and ammonia determinations were made using ion selective electrode technology. Oxides of nitrogen were determined using chemiluminescence instrumentation. Airborne particulate mass concentration was determined using infrared forward scattering measurements and a tapered element oscillating microbalance, as well as conventional gravimetry. Particulate phase metals were determined by collection on Teflon membrane filters, followed by inductively coupled plasma and atomic absorption analysis. Particulate phase polynuclear aromatic hydrocarbons (PAH) and nitro-PAH were collected using high volume sampling on a two stage filter. Target species were extracted, and quantified by gas chromatography/mass spectrometry (GC/MS). Vapor phase species were collected on multi-sorbent resin traps, and subjected to thermal desorption GC/MS for analysis. 11 refs., 1 fig., 1 tab.

Jenkins, R.A.

1990-06-07

89

Abatement of an aircraft exhaust plume using aerodynamic baffles.  

PubMed

The exhaust jet from a departing commercial aircraft will eventually rise buoyantly away from the ground; given the high thrust/power (i.e., momentum/buoyancy) ratio of modern aero-engines, however, this is a slow process, perhaps requiring ? 1 min or more. Supported by theoretical and wind tunnel modeling, we have experimented with an array of aerodynamic baffles on the surface behind a set of turbofan engines of 124 kN thrust. Lidar and point sampler measurements show that, as long as the intervention takes place within the zone where the Coanda effect holds the jet to the surface (i.e., within about 70 m in this case), then quite modest surface-mounted baffles can rapidly lift the jet away from the ground. This is of potential benefit in abating both surface concentrations and jet blast downstream. There is also some modest acoustic benefit. By distributing the aerodynamic lift and drag across an array of baffles, each need only be a fraction of the height of a single blast fence. PMID:23343109

Bennett, Michael; Christie, Simon M; Graham, Angus; Garry, Kevin P; Velikov, Stefan; Poll, D Ian; Smith, Malcolm G; Mead, M Iqbal; Popoola, Olalekan A M; Stewart, Gregor B; Jones, Roderic L

2013-03-01

90

Hydrochloric acid aerosol and gaseous hydrogen chloride partitioning in a cloud contaminated by solid rocket exhaust  

NASA Technical Reports Server (NTRS)

Partitioning of hydrogen chloride between hydrochloric acid aerosol and gaseous HCl in the lower atmosphere was experimentally investigated in a solid rocket exhaust cloud diluted with humid ambient air. Airborne measurements were obtained of gaseous HCl, total HCl, relative humidity and temperature to evaluate the conditions under which aerosol formation occurs in the troposphere in the presence of hygroscopic HCl vapor. Equilibrium predictions of HCl aerosol formation accurately predict the measured HCl partitioning over a range of total HCl concentrations from 0.6 to 16 ppm.

Sebacher, D. I.; Bendura, R. J.; Wornom, D. E.

1980-01-01

91

Temperature, Pressure, and Infrared Image Survey of an Axisymmetric Heated Exhaust Plume  

NASA Technical Reports Server (NTRS)

The focus of this research is to numerically predict an infrared image of a jet engine exhaust plume, given field variables such as temperature, pressure, and exhaust plume constituents as a function of spatial position within the plume, and to compare this predicted image directly with measured data. This work is motivated by the need to validate computational fluid dynamic (CFD) codes through infrared imaging. The technique of reducing the three-dimensional field variable domain to a two-dimensional infrared image invokes the use of an inverse Monte Carlo ray trace algorithm and an infrared band model for exhaust gases. This report describes an experiment in which the above-mentioned field variables were carefully measured. Results from this experiment, namely tables of measured temperature and pressure data, as well as measured infrared images, are given. The inverse Monte Carlo ray trace technique is described. Finally, experimentally obtained infrared images are directly compared to infrared images predicted from the measured field variables.

Nelson, Edward L.; Mahan, J. Robert; Birckelbaw, Larry D.; Turk, Jeffrey A.; Wardwell, Douglas A.; Hange, Craig E.

1996-01-01

92

Simulation of the evolution of particle size distributions in a vehicle exhaust plume with unconfined dilution by ambient air.  

PubMed

Over the past several years, numerous studies have linked ambient concentrations of particulate matter (PM) to adverse health effects, and more recent studies have identified PM size and surface area as important factors in determining the health effects of PM. This study contributes to a better understanding of the evolution of particle size distributions in exhaust plumes with unconfined dilution by ambient air. It combines computational fluid dynamics (CFD) with an aerosol dynamics model to examine the effects of different streamlines in an exhaust plume, ambient particle size distributions, and vehicle and wind speed on the particle size distribution in an exhaust plume. CFD was used to calculate the flow field and gas mixing for unconfined dilution of a vehicle exhaust plume, and the calculated dilution ratios were then used as input to the aerosol dynamics simulation. The results of the study show that vehicle speed affected the particle size distribution of an exhaust plume because increasing vehicle speed caused more rapid dilution and inhibited coagulation. Ambient particle size distributions had an effect on the smaller sized particles (approximately 10 nm range under some conditions) and larger sized particles (>2 microm) of the particle size distribution. The ambient air particle size distribution affects the larger sizes of the exhaust plume because vehicle exhaust typically contains few particles larger than 2 microm. Finally, the location of a streamline in the exhaust plume had little effect on the particle size distribution; the particle size distribution along any streamline at a distance x differed by less than 5% from the particle size distributions along any other streamline at distance x. PMID:15887887

Jiang, Pengzhi; Lignell, David O; Kelly, Kerry E; Lighty, JoAnn S; Sarofim, Adel F; Montgomery, Christopher J

2005-04-01

93

The effect of rocket plume contamination on the optical properties of transmitting and reflecting materials.  

NASA Technical Reports Server (NTRS)

The preliminary results of plume contamination from a 5-pound thrust single-doublet, bipropellant rocket engine on the transmittance of quartz and the reflectance of a silicon monoxide overcoated aluminum mirror have been presented. Changes in quartz transmittance were found to be significant and were due to both absorption and scattering effects. Contaminant absorption effects were predominant at the short wavelengths and scattering effects were greatest in the visible wavelengths. Measured changes in mirror reflectance were due primarily to contaminant absorption. Scattering effects were found to be as much as 9% of the total reflected energy from the mirror. There were no noticeable chemical or erosion effects on either the quartz or the front surface mirror.

Jack, J. R.; Spisz, E. W.; Cassidy, J. F.

1972-01-01

94

The effect of rocket plume contamination on the optical properties of transmitting and reflecting materials  

NASA Technical Reports Server (NTRS)

The preliminary results of plume contamination from a 5-pound thrust single-doublet, bipropellant rocket engine on the transmittance of quartz and the reflectance of a silicon monoxide overcoated aluminum mirror are presented. Changes in quartz transmittance were found to be significant and were due to both absorption and scattering effects. Contaminant absorption effects were predominant at the short wavelengths and scattering effects were greatest in the visible wavelengths. Measured changes in mirror reflectance were due primarily to contaminant absorption. Scattering effects were found to be as much as 9 percent of the total reflected energy from the mirror. There were no noticeable chemical or erosion effects on either the quartz or the front surface mirror.

Jack, J. R.; Spisz, E. W.; Cassidy, J. F.

1971-01-01

95

Particulate emissions in the exhaust plume from commercial jet aircraft under cruise conditions  

NASA Astrophysics Data System (ADS)

In situ measurements of total concentration, size distribution, and hydration properties of jet engine exhaust from a range of commercial transports is reported. Significant concentration enhancements (above ambient background) for aircraft exhaust particulates is reported permitting the detection of not only newly formed but also aged plumes, even in the presence of considerable ambient pollution. Two types of particle size distributions are found in the near-field (˜8 km behind source) exhaust plume from jet aircraft operating under cruise conditions. One type exhibits the form of the Junge distribution with exponential coefficient -2.4. The second exhibits the Junge distribution form in the small-particle region, below about 50 nm, followed by a larger-particle mode between 0.1 and 0.2 ?m. Neither of these observed types of distributions exhibit the sharp drop-off in particle concentrations at the small-particle end of the spectrum that was found in ground-based engine tests. Binary nucleation of sulfuric acid aerosols and heterogeneous nucleation on ion clusters are postulated for particles in this size range. This is supported by the finding of significant numbers of particles having high soluble mass fractions. These data are compared with those taken in ground test cells and those reported by other investigators.

Hagen, D. E.; Whitefield, P. D.; Schlager, H.

1996-08-01

96

Dynamic Analysis of a Building Under Rocket Engine Plume Acoustic Load  

NASA Technical Reports Server (NTRS)

Studies have been performed to develop finite-element modeling and simulation techniques to predict the dynamic structural response of Building 4010 to the acoustic load from the plume of high-thrust rocket motors. The building is the Test Control Center and general office space for the E-complex at Stennis Space Center. It is a large single span; light-structured building located approximately 1,000 feet from the E-1 test stand. A three-dimensional shell/beam combined model of the building was built using Pro/Engineer platform and imported into Pro/Mechanica for analysis. An Equivalent Shell technique was developed to simplify the highly complex building structure so that the calculation is more efficient and accurate. A deterministic approach was used for the dynamic analysis. A pre-stressed modal analysis was performed to simulate the weight stiffening of the structure, through which about 200 modes ranging from 0 to 35 Hz were identified. In an initial dynamic frequency analysis, the maximum response over the model was found. Then the complete 3-D distributions of the displacement, as well as the stresses, were calculated through a final frequency analysis. The results were compared to a strain gage and accelerometer recordings from rocket engine tests and showed reasonable agreement.

Qian, Z.; VanDyke, D.; Wright, S.; Redmond, M.

2001-01-01

97

Method and apparatus for suppressing ignition overpressure in solid rocket propulsion systems  

NASA Technical Reports Server (NTRS)

The transient overpressure wave produced upon ignition of a solid rocket booster is suppressed by providing within the launch platform, a plurality of pipes and spray heads disposed around the periphery of the exhaust gas plume near its upper end and spraying water into the upper end of the plume during ignition. A large amount of water, preferably equivalent in mass of exhaust products being ejected, is sprayed into the plume in a direction generally perpendicular to plume flow.

Guest, S. H.; Jones, J. H. (inventors)

1982-01-01

98

First direct sulfuric acid detection in the exhaust plume of a jet aircraft in flight  

NASA Astrophysics Data System (ADS)

Sulfuric acid (SA) was for the first time directly detected in the exhaust plume of a jet aircraft in flight. The measurements were made by a novel aircraft-based VACA (Volatile Aerosol Component Analyzer) instrument of MPI-K Heidelberg while the research aircraft Falcon was chasing another research aircraft ATTAS. The VACA measures the total SA in the gas and in volatile submicron aerosol particles. During the chase the engines of the ATTAS alternatively burned sulfur-poor and sulfur-rich fuel. In the sulfur-rich plume very marked enhancements of total SA were observed of up to 1300 pptv which were closely correlated with ?CO2 and ?T and were far above the local ambient atmospheric background-level of typically 15-50 pptv. Our observations indicate a lower limit for the efficiency ? for fuel-sulfur conversion to SA of 0.34 %.

Curtius, J.; Sierau, B.; Arnold, F.; Baumann, R.; Busen, R.; Schulte, P.; Schumann, U.

99

A Collimated Retarding Potential Analyzer for the Study of Magnetoplasma Rocket Plumes  

NASA Technical Reports Server (NTRS)

A gridded retarding potential analyzer (RPA) has been developed to characterize the magnetized plasma exhaust of the 10 kW Variable Specific Impulse Magnetoplasma Rocket (VX-10) experiment at NASA's Advanced Space Propulsion Laboratory. In this system, plasma is energized through coupling of radio frequency waves at the ion cyclotron resonance (ICR). The particles are subsequently accelerated in a magnetic nozzle to provide thrust. Downstream of the nozzle, the RPA's mounting assembly enables the detector to make complete axial and radial scans of the plasma. A multichannel collimator can be inserted into the RPA to remove ions with pitch angles greater than approximately 1 deg. A calculation of the general collimator transmission as a function over velocity space is presented, which shows the instrument's sensitivity in detecting changes in both the parallel and perpendicular components of the ion energy. Data from initial VX-10 ICRH experiments show evidence of ion heating.

Glover, T. W.; Chan, A. A.; Chang-Diaz, F. R.; Kittrell, C.

2003-01-01

100

Ionospheric effects of rocket exhaust products (HEAO-C, Skylab and SPS-HLLV)  

SciTech Connect

This paper reviews the current state of our understanding of the problem of ionospheric F-layer depletions produced by chemical effects of the exhaust gases from large rockets, with particular emphasis on the Heavy Lift Launch Vehicles (HLLV) proposed for use in the construction of solar power satellites. The currently planned HLLV flight profile calls for main second-stage propulsion confined to altitudes below 124 km, and a brief orbit-circularization maneuver at apogee. The second-stage engines deposit 9 x 10/sup 31/ H/sub 2/O and H/sub 2/ molecules between 56 and 124 km. Model computations show that they diffuse gradually into the ionospheric F region, where they lead to weak but widespread and persistent depletions of ionization and continuous production of H atoms. The orbit-circularization burn deposits 9 x 10/sup 29/ exhaust molecules at about 480-km altitude. These react rapidly with the F2 region 0/sup +/ ions, leading to a substantial (factor-of-three) reduction in plasma density, which extends over a 1000- by 2000-km region and persists for four to five hours. Also described are experimental airglow and incoherent-scatter radar measurements performed in conjunction with the 1979 launch of satellite HEAO-C, together with prelaunch and post-launch computations of the ionospheric effects. Several improvements in the model have been driven by the experimental observations. The computer model is described in some detail.

Zinn, J; Sutherland, D; Stone, S N; Duncan, L M; Behnke, R

1980-10-01

101

Rockets  

NSDL National Science Digital Library

This teacher's guide for rocketry presents the history, scientific principles and mathematics of rockets through problem-solving and cooperative learning activities. These activities demonstrate the physical principles behind the operation of rockets.

102

Effects of nozzle exit geometry and pressure ratio on plume shape for nozzles exhausting into quiescent air  

NASA Technical Reports Server (NTRS)

The effects of varying the exit geometry on the plume shapes of supersonic nozzles exhausting into quiescent air at several exit-to-ambient pressure ratios are given. Four nozzles having circular throat sections and circular, elliptical and oval exit cross sections were tested and the exit plume shapes are compared at the same exit-to-ambient pressure ratios. The resulting mass flows were calculated and are also presented.

Scallion, William I.

1991-01-01

103

Theoretical and Experimental Investigation of Lunar and Martian Regolith Simulant Dynamic Response to Rocket Plume Impingement  

NASA Astrophysics Data System (ADS)

An investigation of rocket plume impingement on the regolith of the Moon and Mars is being conducted both theoretically and experimentally. Experimental results (1)and data from the Apollo landings inspired a theoretical model at ORBITEC : the ABL (Ablating Boundary Layer) model that assumes that regolith erosion and entrainment occurs in the thin boundary layer. The resulting crater streamlines itself with curve formed by extremization of the Lagrangian : L = (Z')^2+ Z^2 where Z(r) and Z(r)' are a depth variable and its radial derivative respectively. The actual depth profile z (r) in this model is derived from the formula z=Log ( 1+ Z/Zo) where Zo is a constant. For light soils the model reduces to z˜ Z/Zo and cantenary profiles result, exponential density profiles (2) give conoidal craters. (1) Experimental tests of the ABL model performed at Duke have shown good agreement. Further theoretical modeling and experimental data will be presented. (1) Metzger P., Lane, J., Immer C. and Clements, S. '6^th International Conference on Case Histories in Geotechnicla Engineeering , Arlinton VA August 11-16, 2008. (2) Bresson L. M., Moran C. J., and Assoline, S. Soil Sci. Soc. of Am. Jou, 2004, vol. 68, 4, pp. 1169-1176.

Brandenburg, John; Behringer, Robert; Clarke, Abraham

2009-11-01

104

Application of a Gaussian multilayer diffusion model to characterize dispersion of vertical HCl column density in rocket exhaust clouds  

NASA Technical Reports Server (NTRS)

Solid rocket exhaust cloud dispersion cases, based on seven meteorological regimes for overland advection in the Cape Canaveral, Florida, area, are examined for launch vehicle environmental impacts. They include a space shuttle case and all seven meteorological cases for the Titan 3, which exhausts 60% less HC1. The C(HC1) decays are also compared with recent in cloud peak HC1 data from eight Titan 3 launches. It is stipulated that while good overall agreement provides validation of the model, its limitations are considerable and a dynamics model is needed to handle local convective situations.

Pellett, G. L.; Staton, W. L.

1981-01-01

105

Coupled turbulence and aerosol dynamics modeling of vehicle exhaust plumes using the CTAG model  

NASA Astrophysics Data System (ADS)

This paper presents the development and evaluation of an environmental turbulent reacting flow model, the Comprehensive Turbulent Aerosol Dynamics and Gas Chemistry (CTAG) model. CTAG is designed to simulate transport and transformation of multiple air pollutants, e.g., from emission sources to ambient background. For the on-road and near-road applications, CTAG explicitly couples the major turbulent mixing processes, i.e., vehicle-induced turbulence (VIT), road-induced turbulence (RIT) and atmospheric boundary layer turbulence with gas-phase chemistry and aerosol dynamics. CTAG's transport model is referred to as CFD-VIT-RIT. This paper presents the evaluation of the CTAG model in simulating the dynamics of individual plumes in the “tailpipe-to-road” stage, i.e., VIT behind a moving van and aerosol dynamics in the wake of a diesel car by comparing the modeling results against the respective field measurements. Combined with sensitivity studies, we analyze the relative roles of VIT, sulfuric acid induced nucleation, condensation of organic compounds and presence of soot-mode particles in capturing the dynamics of exhaust plumes as well as their implications in vehicle emission controls.

Wang, Yan Jason; Zhang, K. Max

2012-11-01

106

Experimental measurements of the ground cloud growth during the 11 February 1974, Titan-Centaur launch at Kennedy Space Center. [(measurement of rocket exhaust from rocket launching)  

NASA Technical Reports Server (NTRS)

The Titan-Centaur was launched from Kennedy Space Center on February 11, 1974 at 0948 eastern daylight time. Ground level effluent measurements were obtained from the solid rocket motors for comparison with NASA diffusion models for predicting effluent ground level concentrations and cloud behavior. The results obtained provide a basis for an evaluation of such key model inputs such as cloud rise rate, stabilization altitude, crosswind growth, volume expansion, and cloud trajectory. Ground level effluent measurements were limited because of changing meteorological conditions, incorrect instrument location, and operational problems. Based on the measurement results, operational changes are defined. Photographs of the ground exhaust clouds are shown. The chemical composition of the exhaust gases was analyzed and is given.

Stewart, R. B.; Sentell, R. J.; Gregory, G. L.

1976-01-01

107

Fluid Gravity Engineering Rocket motor flow analysis  

E-print Network

Fluid Gravity Engineering Capability · Rocket motor flow analysis -Internal (performance) -External response · Rocket motor plume IR assessments -Plume extent, composition and temperature -Radiation

Anand, Mahesh

108

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Vectored Nozzles  

NASA Technical Reports Server (NTRS)

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions were due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed how the shock wave formed at the nozzle lip interacts with the nozzle boat-tail expansion wave. An experiment was conducted in the 1- by 1-foot Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center. Results show how the shock generated at the nozzle lip affects the near field pressure signature, and thereby the potential sonic boom contribution for a nozzle at vector angles from 3 to 8 . The experiment was based on the NASA F-15 nozzle used in the Lift and Nozzle Change Effects on Tail Shock experiment, which possessed a large external boat-tail angle. In this case, the large boat-tail angle caused a dramatic expansion, which dominated the near field pressure signature. The impact of nozzle vector angle and nozzle pressure ratio are summarized.

Castner, Raymond

2012-01-01

109

Spectroscopic studies of the exhaust plume of a quasi-steady MPD accelerator. Ph.D. Thesis  

NASA Technical Reports Server (NTRS)

Spectroscopic and photographic investigations are reported that reveal a complex azimuthal species structure in the exhaust plume of a quasi-steady argon MPD accelerator. Over a wide range of operating conditions the injected argon remains collimated in discrete jets which are azimuthally in line with the six propellant injector orifices. The regions between these argon jets, including the central core of the exhaust flow, are occupied by impurities such as carbon, hydrogen and oxygen ablated from the Plexiglas back plate of the arc chamber. The features of this plume structure are found to be dependent on the arc current and mass flow rate. It is found that nearly half the observed velocity is attained in an acceleration region well downstream of the region of significant electromagnetic interaction. Recombination calculations show that the ionization energy is essentially frozen.

Bruckner, A. P.

1972-01-01

110

Apollo 12 Lunar Module exhaust plume impingement on Lunar Surveyor III  

NASA Astrophysics Data System (ADS)

Understanding plume impingement by retrorockets on the surface of the Moon is paramount for safe lunar outpost design in NASA's planned return to the Moon for the Constellation Program. Visual inspection, Scanning Electron Microscopy, and surface scanned topology have been used to investigate the damage to the Lunar Surveyor III spacecraft that was caused by the Apollo 12 Lunar Module's close proximity landing. Two parts of the Surveyor III craft returned by the Apollo 12 astronauts, Coupons 2050 and 2051, which faced the Apollo 12 landing site, show that a fine layer of lunar regolith coated the materials and was subsequently removed by the Apollo 12 Lunar Module landing rocket. The coupons were also pitted by the impact of larger soil particles with an average of 103 pits/cm 2. The average entry size of the pits was 83.7 ?m (major diameter) × 74.5 ?m (minor diameter) and the average estimated penetration depth was 88.4 ?m. Pitting in the surface of the coupons correlates to removal of lunar fines and is likely a signature of lunar material imparting localized momentum/energy sufficient to cause cracking of the paint. Comparison with the lunar soil particle size distribution and the optical density of blowing soil during lunar landings indicates that the Surveyor III spacecraft was not exposed to the direct spray of the landing Lunar Module, but instead experienced only the fringes of the spray of soil. Had Surveyor III been exposed to the direct spray, the damage would have been orders of magnitude higher.

Immer, Christopher; Metzger, Philip; Hintze, Paul E.; Nick, Andrew; Horan, Ryan

2011-02-01

111

Sampling Particles In Hot Gas Plumes  

NASA Technical Reports Server (NTRS)

Sampling darts and launching apparatus built to collect particles in vertical plume of hot gas. In original application, hot gas plume is rocket-engine exhaust during test firing. Dart passes made at various heights, depending on launch angle and launch-gas pressure. Adaptable to variety of terrestrial uses like research on particulate emissions of volcanoes or determining origin of building fire while still burning.

Taylor, James F.; Sambamurthi, Jay

1994-01-01

112

Rocket plume spectrometry: A system permitting engine condition monitoring, as applied to the technology test bed engine  

NASA Technical Reports Server (NTRS)

The appearance of visible objects in the exhaust plume of space shuttle main engines (SSME) during test firings is discussed. A program was undertaken to attempt to identify anomalous material resulting from wear, normal or excessive, of internal parts, allowing time monitoring of engine condition or detection of failure precursors. Measurements were taken during test firings at Stennis Space Center and at the Santa Suzanna facility in California. The results indicated that a system having high spectral resolution, a fast time response, and a wide spectral range was required to meet all requirements, thus two special systems have been designed and built. One is the Optical Plume Anomaly Detector (OPAD). The other instrument, which is described in this report, is the superspectrometer, an optical multichannel analyzer having 8,192 channels covering the spectral band 250 to 1,000 nm.

Powers, W. T.

1989-01-01

113

Sulfuric acid measurements in the exhaust plume of a jet aircraft in flight: Implications for the sulfuric acid formation efficiency  

NASA Astrophysics Data System (ADS)

Sulfuric acid concentrations were measured in the exhaust plume of a B737-300 aircraft in flight. The measurements were made onboard of the German research aircraft Falcon using the Volatile Aerosol Component Analyzer (VACA). The VACA measures total H2SO4, which is the sum of gaseous H2SO4 and aerosol H2SO4. Measurements took place at distances of 25-200 m behind the B737 corresponding to plume ages of about 0.1-1 seconds. The fuel sulfur content (FSC) of the fuel burned by the B737 engines was alternatively 2.6 and 56 mg sulfur per kilogram fuel (ppmm). H2SO4 concentrations measured in the plume for the 56 ppmm sulfur case were up to ~600 pptv. The average concentration of H2SO4 measured in the ambient atmosphere outside the aircraft plume was 88 pptv, the maximum ambient atmospheric H2SO4 was ~300 pptv. Average efficiencies ??CO2 = 3.3 +/- 1.8% and ??T = 2.9 +/- 1.6% for fuel sulfur conversion to sulfuric acid were inferred when relating the H2SO4 data to measurements of the plume tracers ?CO2 and ?T.

Curtius, J.; Arnold, F.; Schulte, P.

2002-04-01

114

Measurements of HONO, NO, NOy and SO2 in aircraft exhaust plumes at cruise  

NASA Astrophysics Data System (ADS)

Measurements of gaseous nitrogen and sulfur oxide emissions in young aircraft exhaust plumes give insight into chemical oxidation processes inside aircraft engines. Particularly, the OH-induced formation of nitrous acid (HONO) from nitrogen oxide (NO) and sulfuric acid (H2SO4) from sulfur dioxide (SO2) inside the turbine which is highly uncertain, need detailed analysis to address the climate impact of aviation. We report on airborne in situ measurements at cruise altitudes of HONO, NO, NOy, and SO2 in 9 wakes of 8 different types of modern jet airliners, including for the first time also an A380. Measurements of HONO and SO2 were made with an ITCIMS (Ion Trap Chemical Ionization Mass Spectrometer) using a new ion-reaction scheme involving SF5- reagent ions. The measured molar ratios HONO/NO and HONO/NOy with averages of 0.038 ± 0.010 and 0.027 ± 0.005 were found to decrease systematically with increasing NOx emission-index (EI NOx). We calculate an average EI HONO of 0.31 ± 0.12 g NO2 kg-1. Using reliable measurements of HONO and NOy, which are less adhesive than H2SO4 to the inlet walls, we derive the OH-induced conversion fraction of fuel sulfur to sulfuric acid $\\varepsilon$ with an average of 2.2 ± 0.5 %. $\\varepsilon$ also tends to decrease with increasing EI NOx, consistent with earlier model simulations. The lowest HONO/NO, HONO/NOy and $\\varepsilon$ was observed for the largest passenger aircraft A380.

Jurkat, T.; Voigt, C.; Arnold, F.; Schlager, H.; Kleffmann, J.; Aufmhoff, H.; Schäuble, D.; Schaefer, M.; Schumann, U.

2011-05-01

115

Phenomenology of soil erosion due to rocket exhaust on the Moon and the Mauna Kea lunar test site  

NASA Astrophysics Data System (ADS)

The soil-blowing phenomena observed in the Apollo lunar missions have not previously been described in the literature in sufficient detail to elucidate the physical processes and to support the development of physics-based modeling of the plume effects. In part, this is because previous laboratory experiments have used overly simplistic model soils that fail to produce many of the phenomena seen in lunar landings, some of which therefore went unrecognized. Here, the Apollo descent videos, terrain photography, and ascent videos are interpreted with the assistance of field experiments using a more complex regolith. Rocket thruster firings were performed upon the tephra of a lunar test site on Mauna Kea in Hawaii. This tephra possesses embedded rocks, large fractions of gravel and dust, some cohesion, and natural geological lamination. This produced more realistic plume phenomenology. The relevant phenomena include the relationship of dust liberation with overall soil erosion rate, terrain bed forms created by the plume, dust tails associated with the exhumation and blowing of rocks, bed load transport, the removal of discrete layers of soil hypothesized to be the stratigraphic units corresponding to impact events, the total mass of ejected soil during a landing, and the brightening of the regolith around the landing site. This analysis provides insight into the erosion processes and nature of the regolith. This paper also synthesizes theory, experiment, simulation, and observational data to produce a clearer picture of the physical processes of lunar soil erosion.

Metzger, Philip T.; Smith, Jacob; Lane, John E.

2011-06-01

116

U. S. Air Force approach to plume contamination  

NASA Astrophysics Data System (ADS)

Exhaust products from rocket engine firings can produce undesirable effects on sensitive satellite surfaces, such as optical systems, solar cells, and thermal control surfaces. The Air Force has an objective of minimizing the effect of rocket plume contamination on space-craft mission effectiveness. Plume contamination can result from solid rocket motors, liquid propellant engines, and electric thrusters. To solve the plume contamination problem, the Air Force Rocket Propulsion Laboratory (AFRPL) has developed a plume contamination computer model which predicts the production, transport, and deposition of rocket exhaust products. In addition, an experimental data base is being obtained through ground-based vacuum chamber experiments and in-flight measurements with which to compare the analytical results. Finally, the experimental data is being used to verify and improve the analytical model. The plume contamination model, known as CONTAM, has been used to make contamination predictions for various engines. The experimental programs have yielded quantitative data, such as species concentrations and temperatures, in all regions of the plume. The result of the modelling and experimental programs will ultimately be computer models which can be used by the satellite designer to analyze and to minimize the effect plume contamination will have on a particular spacecraft system.

Furstenau, Ronald P.; McCay, T. Dwayne; Mann, David M.

1980-08-01

117

Influence of fuel sulfur on the composition of aircraft exhaust plumes: The experiments SULFUR 1-7  

NASA Astrophysics Data System (ADS)

The series of SULFUR experiments was performed to determine the aerosol particle and contrail formation properties of aircraft exhaust plumes for different fuel sulfur contents (FSC, from 2 to 5500 ?g/g), flight conditions, and aircraft (ATTAS, A310, A340, B707, B747, B737, DC8, DC10). This paper describes the experiments and summarizes the results obtained, including new results from SULFUR 7. The conversion fraction ? of fuel sulfur to sulfuric acid is measured in the range 0.34 to 4.5% for an older (Mk501) and 3.3 +/- 1.8% for a modern engine (CFM56-3B1). For low FSC, ? is considerably smaller than what is implied by the volume of volatile particles in the exhaust. For FSC >= 100 ?g/g and ? as measured, sulfuric acid is the most important precursor of volatile aerosols formed in aircraft exhaust plumes of modern engines. The aerosol measured in the plumes of various aircraft and models suggests ? to vary between 0.5 and 10% depending on the engine and its state of operation. The number of particles emitted from various subsonic aircraft engines or formed in the exhaust plume per unit mass of burned fuel varies from 2 × 1014 to 3 × 1015 kg-1 for nonvolatile particles (mainly black carbon or soot) and is of order 2 × 1017 kg-1 for volatile particles >1.5 nm at plume ages of a few seconds. Chemiions (CIs) formed in kerosene combustion are found to be quite abundant and massive. CIs contain sulfur-bearing molecules and organic matter. The concentration of CIs at engine exit is nearly 109 cm-3. Positive and negative CIs are found with masses partially exceeding 8500 atomic mass units. The measured number of volatile particles cannot be explained with binary homogeneous nucleation theory but is strongly related to the number of CIs. The number of ice particles in young contrails is close to the number of soot particles at low FSC and increases with increasing FSC. Changes in soot particles and FSC have little impact on the threshold temperature for contrail formation (less than 0.4 K).

Schumann, U.; Arnold, F.; Busen, R.; Curtius, J.; Kärcher, B.; Kiendler, A.; Petzold, A.; Schlager, H.; Schröder, F.; Wohlfrom, K.-H.

2002-08-01

118

Results of an investigation of jet plume effects on a 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 8 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82C), volume 1. [(an exhaust flow simulation)  

NASA Technical Reports Server (NTRS)

The primary test objective was to define the base pressure environment of the first and second stage mated vehicle in a supersonic flow field from Mach 2.60 through 3.50 with simulated rocket engine exhaust plumes. The secondary objective was to obtain the pressure environment of the Orbiter at various vent port locations at these same freestream conditions. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed Orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of test equipment and tested configurations are shown.

Hawthorne, P. J.

1976-01-01

119

Direct active measurements of movements of lunar dust: Rocket exhausts and natural effects contaminating and cleansing Apollo hardware on the Moon in 1969  

Microsoft Academic Search

Dust is the Number 1 environmental hazard on the Moon, yet its movements and adhesive properties are little understood. Matchbox-sized, 270-gram Dust Detector Experiments (DDEs) measured contrasting effects triggered by rocket exhausts of Lunar Modules (LM) after deployment 17 m and 130 m from Apollo 11 and 12 LMs. Apollo 11 Lunar Seismometer was contaminated, overheated and terminated after 21

Brian O'Brien

2009-01-01

120

Electromagnetic Effects in the Near Field Plume Exhaust of a Micro-Pulsed Plasma Thruster  

NASA Astrophysics Data System (ADS)

In this work we present a model of the near field plasma plume of a Pulsed Plasma Thruster (PPT). As a working example we consider a micro-PPT developed at the Air Force Research Laboratory. This is a miniaturized design of the axisymmetric PPT with a thrust in the 10 micro-N range that utilizes Teflon(TrademarkTrademark) as a propellant. The plasma plume is simulated using a hybrid fluid-PIC-DSMC approach. The plasma plume model is combined with Teflon(Trademark) ablation and plasma generation models that provide boundary conditions for the plume. This approach provides a consistent description of the plasma flow from the surface into the near plume. The magnetic field diffusion into the plume region is also considered and plasma acceleration by the electromagnetic mechanism is studied. Teflon(Trademark) ablation and plasma generation analyses show that file Teflon(Trademark) surface temperature and plasma parameters are strongly non-uniform in the radial direction. The plasma density near the propellant surface peaks at about 1024/cu m in the middle of the propellant face while the electron temperature peaks at about 4 eV near the electrodes. The plume simulation shows that a dense plasma focus is developed at a few millimeters from the thruster exit plane at the axis. This plasma focus exists during the entire pulse, but the plasma density in the focus decreases from about 2x1022/cu m at the beginning of the pulse down to 0. 3x1022/cu m at 5 microsec. The velocity phase is centered at about 20 km/s in the axial direction. At later stages of the pulse there are two ion populations with positive and negative radial velocity. Electron densities predicted by file plume model are compared with near field measurements using a Herriot Cell technique and very good agreement is obtained.

Keidar, Michael; Boyd, Iain D.; Antonsen, Eric; Spanjers, Gregory G.

2002-06-01

121

Digital filtering of plume emission spectra  

NASA Technical Reports Server (NTRS)

Fourier transformation and digital filtering techniques were used to separate the superpositioned spectral phenomena observed in the exhaust plumes of liquid propellant rocket engines. Space shuttle main engine (SSME) spectral data were used to show that extraction of spectral lines in the spatial frequency domain does not introduce error, and extraction of the background continuum introduces only minimal error. Error introduced during band extraction could not be quantified due to poor spectrometer resolution. Based on the atomic and molecular species found in the SSME plume, it was determined that spectrometer resolution must be 0.03 nm for SSME plume spectral monitoring.

Madzsar, George C.

1990-01-01

122

First gaseous ion composition measurements in the exhaust plume of a jet aircraft in flight: Implications for gaseous sulfuric acid, aerosols, and chemiions  

NASA Astrophysics Data System (ADS)

Mass spectrometric composition measurements of gaseous negative ions have been made in the exhaust plume of a commercial jet aircraft (Airbus A310) in flight at altitudes around 10.4 km and at two plume ages around 3.0 and 3.6 s. Negative ions observed inside the exhaust plume are mostly NO3-(HNO3)m and HSO4-(HNO3)m with m ?2. Outside the plume in the “background” atmosphere the same negative ion species with the same R = (HSO4-(HNO3)m)/(NO3-(HNO3)m) were observed. This indicates that the ions observed in the plume were entrained ambient atmospheric ions. By contrast no indications for negative chemiions (with masses ?1100 amu) produced by the airbus engines were found in the plume. Furthermore our measurements indicate a modest decrease of the total concentration of entrained negative ions in the plume compared to the ambient atmosphere outside the plume. This decrease may be due to ion-removal by ion-attachment to aerosol-particles and/or ion-recombination with positive chemiions. We propose that the observed entrained ions can serve as probes for important plume components including gaseous sulfuric acid, aerosol particles and chemiions. Making use of this analytical potential we infer upper limits for the gaseous sulfuric acid concentration, total aerosol surface area density, and positive chemiion concentration. We conclude that initially formed gaseous sulfuric acid must have experienced rapid gas-to-particle conversion already in the very early plume at plume ages < 1.6 s.

Arnold, F.; Wohlfrom, K.-H.; Klemm, M. W.; Schneider, J.; Gollinger, K.; Schumann, U.; Busen, R.

123

Ionospheric shock waves triggered by rockets  

NASA Astrophysics Data System (ADS)

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.

Lin, C. H.; Lin, J. T.; Chen, C. H.; Liu, J. Y.; Sun, Y. Y.; Kakinami, Y.; Matsumura, M.; Chen, W. H.; Liu, H.; Rau, R. J.

2014-09-01

124

Process-Hardened, Multi-Analyte Sensor for Characterizing Rocket Plume Constituents  

NASA Technical Reports Server (NTRS)

A multi-analyte sensor was developed that enables simultaneous detection of rocket engine combustion-product molecules in a launch-vehicle ground test stand. The sensor was developed using a pin-printing method by incorporating multiple sensor elements on a single chip. It demonstrated accurate and sensitive detection of analytes such as carbon dioxide, carbon monoxide, kerosene, isopropanol, and ethylene from a single measurement. The use of pin-printing technology enables high-volume fabrication of the sensor chip, which will ultimately eliminate the need for individual sensor calibration since many identical sensors are made in one batch. Tests were performed using a single-sensor chip attached to a fiber-optic bundle. The use of a fiber bundle allows placement of the opto-electronic readout device at a place remote from the test stand. The sensors are rugged for operation in harsh environments.

Goswami, Kisholoy

2011-01-01

125

HCl in rocket exhaust clouds - Atmospheric dispersion, acid aerosol characteristics, and acid rain deposition  

NASA Technical Reports Server (NTRS)

Both measurements and model calculations of the temporal dispersion of peak HCl (g + aq) concentration in Titan III exhaust clouds are found to be well characterized by one-term power-law decay expressions. The respective coefficients and decay exponents, however, are found to vary widely with meteorology. The HCl (g), HCl (g + aq), dewpoint, and temperature-pressure-altitude data for Titan III exhaust clouds are consistent with accurately calculated HCl/H2O vapor-liquid compositions for a model quasi-equilibrated flat surface aqueous aerosol. Some cloud evolution characteristics are also defined. Rapid and extensive condensation of aqueous acid clearly occurs during the first three min of cloud rise. Condensation is found to be intensified by the initial entrainment of relatively moist ambient air from lower levels, that is, from levels below eventual cloud stabilization. It is pointed out that if subsequent dilution air at stabilization altitude is significantly drier, a state of maximum condensation soon occurs, followed by an aerosol evaporation phase.

Pellett, G. L.; Sebacher, D. I.; Bendura, R. J.; Wornom, D. E.

1983-01-01

126

Three-dimensional reconstruction method on the PDE exhaust plume flow flame temperature field  

Microsoft Academic Search

Pulse detonation engine (referred to as PDE) has many advantage about simple structure, high efficiency thermal [1] cycling etc. In the future, it can be widely used in unmanned aircraft, target drone, luring the plane, the imaginary target, target missiles, long-range missiles and other military targets. However, because the exhaust flame of PDE is complicated [2], non-uniform temperature distribution and

Zhimin Zhang; Xiong Wan; Ningning Luo; Shujing Li

2010-01-01

127

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles  

NASA Technical Reports Server (NTRS)

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nose cone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1x1 SWT for Schlieren photography and comparison to CFD analysis.

Castner, Raymond S.

2009-01-01

128

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles  

NASA Technical Reports Server (NTRS)

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1 1 SWT for Schlieren photography and comparison to CFD analysis.

Castner, Raynold S.

2010-01-01

129

Three-dimensional reconstruction method on the PDE exhaust plume flow flame temperature field  

NASA Astrophysics Data System (ADS)

Pulse detonation engine (referred to as PDE) has many advantage about simple structure, high efficiency thermal [1] cycling etc. In the future, it can be widely used in unmanned aircraft, target drone, luring the plane, the imaginary target, target missiles, long-range missiles and other military targets. However, because the exhaust flame of PDE is complicated [2], non-uniform temperature distribution and mutation in real time, its 3-D temperature distribution is difficult to be measured by normal way. As a result, PDE is used in the military project need to face many difficulties and challenges. In order to analyze and improve the working performance of PDE, deep research on the detonation combustion process is necessary. However, its performance characteristic which is in non-steady-state, as well as high temperature, high pressure, transient combustion characteristics put forward high demands about the flow field parameters measurement. In this paper, the PDE exhaust flames temperature field is reconstructed based on the theory of radiation thermometry [3] and Emission Spectral Tomography (referred to as EST) [4~6] which is one branch of Optical CT. It can monitor the detonation wave temperature distribution out of the exhaust flames at different moments, it also provides authentication for the numerical simulation which directs towards PDE work performance, and then it provides the basis for improving the structure of PDE.

Zhang, Zhimin; Wan, Xiong; Luo, Ningning; Li, Shujing

2010-10-01

130

Modeling a VASIMR rocket at UVSC  

NASA Astrophysics Data System (ADS)

A Variable Specific Impulse Magnetohydrodynamic Rocket (VASIMR) takes advantage of magnetic mirrors to confine a plasma long enough to heat it with ion-cyclotron resonant heating before directing the flow through a magnetic nozzle to produce thrust. We are engaged in building a computational model for the flow of ions through the rocket to investigate optimal configuration of the mirrors and to consider the problem of plasma detachment from the magnetic field in the exhaust plume. A description of the model will be presented with preliminary results from the computation.

Matheson, Phil; Gray, William; Page, Leland

2004-10-01

131

Simulation of the evolution of aircraft exhaust plumes including detailed chemistry and segregation  

NASA Astrophysics Data System (ADS)

The Field Monte Carlo or Stochastic Fields (SF) method for turbulent reacting flows has been applied to the chemical evolution of the early part of a hot jet with bypass flow producing 7kN of thrust, using a 23 species chemical mechanism. This is done to broadly approximate a turbofan engine at idle thrust setting. Much of the chemistry was found to take place inside the core of the jet before mixing occurs, as there is no reactant gradient there, considering segregation makes little difference. Radical concentrations, however, were found to be changed. The reaction between NO and ambient O3, which is slow compared to the fast mixing timescale of the turbulent jet, is unaffected by segregation. The local Damköhler number was calculated based on an estimate of the chemical timescale and the local large-eddy timescale. It was found that only those species which had local Da greater than five were affected by segregation. In this work we have applied the SF method the early part of the plume, however the method developed here could equally be employed to study the plume over a longer distance.

Garmory, A.; Britter, R. E.; Mastorakos, E.

2008-04-01

132

High altitude chemically reacting gas particle mixtures. Volume 3: Computer code user's and applications manual. [rocket nozzle and orbital plume flow fields  

NASA Technical Reports Server (NTRS)

A users manual for the RAMP2 computer code is provided. The RAMP2 code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. The general structure and operation of RAMP2 are discussed. A user input/output guide for the modified TRAN72 computer code and the RAMP2F code is given. The application and use of the BLIMPJ module are considered. Sample problems involving the space shuttle main engine and motor are included.

Smith, S. D.

1984-01-01

133

Measurement of plasma parameters in the exhaust of a magnetoplasma rocket by gridded energy analyzer and emissive Langmuir probe  

NASA Astrophysics Data System (ADS)

The 10 kilowatt prototype of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) engine, abbreviated as VX-10, is designed to eject plasma at exhaust velocities of tens of kilometers per second. In this device, energy is imparted to the plasma ions by two mechanisms: ion cyclotron resonant heating (ICRH), and acceleration in an ambipolar electric field. Measurements from two different electrostatic probes are combined to determine how much each mechanism contributes to the total ion energy. The first probe is a gridded retarding potential analyzer (RPA) that incorporates a multi-channel collimator to obtain precise measurement of the ion and electron parallel energy distributions. The second is an emissive Langmuir probe that measures the DC and RF components of the plasma potential. The plasma potential obtained from the emitting probe allows calculation of the parallel velocity distribution once the parallel energy distribution is obtained from the energy analyzer data. Biasing the RPA housing is shown to minimize the plasma perturbation, as monitored by an auxiliary probe. When this minimization is done, the RPA measurements become compatible with the emissive probe's measurement of plasma potential. The collimated RPA and emissive probe have been used to examine the effects of a double dual half-turn (DDHT) antenna encircling the plasma. When power at the ion cyclotron frequency is applied, changes are seen in the saturation current and mean ion energy of the collimated RPA characteristic. The evolution of these changes as the RPA is moved downstream from the antenna is interpreted as firm evidence of ion cyclotron heating, albeit at absorbed energies of less than 1 electronvolt per ion. The emissive probe shows that, within experimental error, all of the increased ion energy is accounted for by an increase in the plasma potential that occurs when the ICRF power is applied. The combined RPA and emissive probe data also show that there is a jet of flowing plasma in the VX-10 when operated with the helicon source alone but that the signal from this jet is overwhelmed by a rapidly growing stationary plasma within the first second of the discharge.

Glover, Timothy Ward

2002-01-01

134

Characterization of rocket propellant combustion products. Chemical characterization and computer modeling of the exhaust products from four propellant formulations: Final report, September 23, 1987--April 1, 1990  

SciTech Connect

The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army`s Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.

1991-12-09

135

Navier-Stokes computations with finite-rate chemistry for LO2/LH2 rocket engine plume flow studies  

NASA Technical Reports Server (NTRS)

Computational fluid dynamics methods have been developed and applied to Space Shuttle Main Engine LO2/LH2 plume flow simulation/analysis of airloading and convective base heating effects on the vehicle at high flight velocities and altitudes. New methods are described which were applied to the simulation of a Return-to-Launch-Site abort where the vehicle would fly briefly at negative angles of attack into its own plume. A simplified two-perfect-gases-mixing approach is used where one gas is the plume and the other is air at 180-deg and 135-deg flight angle of attack. Related research has resulted in real gas multiple-plume interaction methods with finite-rate chemistry described herein which are applied to the same high-altitude-flight conditions of 0 deg angle of attack. Continuing research plans are to study Orbiter wake/plume flows at several Mach numbers and altitudes during ascent and then to merge this model with the Shuttle 'nose-to-tail' aerodynamic and SRB plume models for an overall 'nose-to-plume' capability. These new methods are also applicable to future launch vehicles using clustered-engine LO2/LH2 propulsion.

Dougherty, N. Sam; Liu, Baw-Lin

1991-01-01

136

Rocket motor exhaust products generated by the space shuttle vehicle during its launch phase (1976 design data)  

NASA Technical Reports Server (NTRS)

The principal chemical species emitted and/or entrained by the rocket motors of the space shuttle vehicle during the launch phase of its trajectory are considered. Results are presented for two extreme trajectories, both of which were calculated in 1976.

Bowyer, J. M.

1977-01-01

137

One-Dimensional Rocket Launch  

NSDL National Science Digital Library

A simulation of a 1-d rocket launch from the Earth's surface with graph of position versus time. Rocket parameters may be varied by typing new values for the initial mass of the fuel and the exhaust velocity.

Christian, Wolfgang; Belloni, Mario

2006-01-12

138

Payload dose rate from direct beam radiation and exhaust gas fission products. [for nuclear engine for rocket vehicles  

NASA Technical Reports Server (NTRS)

A study was made to determine the dose rate at the payload position in the NERVA System (1) due to direct beam radiation and (2) due to the possible effect of fission products contained in the exhaust gases for various amounts of hydrogen propellant in the tank. Results indicate that the gamma radiation is more significant than the neutron flux. Under different assumptions the gamma contribution from the exhaust gases was 10 to 25 percent of total gamma flux.

Capo, M. A.; Mickle, R.

1975-01-01

139

Parametric studies with an atmospheric diffusion model that assesses toxic fuel hazards due to the ground clouds generated by rocket launches  

NASA Technical Reports Server (NTRS)

Parametric studies were made with a multilayer atmospheric diffusion model to place quantitative limits on the uncertainty of predicting ground-level toxic rocket-fuel concentrations. Exhaust distributions in the ground cloud, cloud stabilized geometry, atmospheric coefficients, the effects of exhaust plume afterburning of carbon monoxide CO, assumed surface mixing-layer division in the model, and model sensitivity to different meteorological regimes were studied. Large-scale differences in ground-level predictions are quantitatively described. Cloud alongwind growth for several meteorological conditions is shown to be in error because of incorrect application of previous diffusion theory. In addition, rocket-plume calculations indicate that almost all of the rocket-motor carbon monoxide is afterburned to carbon dioxide CO2, thus reducing toxic hazards due to CO. The afterburning is also shown to have a significant effect on cloud stabilization height and on ground-level concentrations of exhaust products.

Stewart, R. B.; Grose, W. L.

1975-01-01

140

Investigation of jet exhaust and local Mach number effects on the Solid Rocket Booster (SRB) base pressures (fa19)  

NASA Technical Reports Server (NTRS)

Two single nozzles with flare angles of 10 and 20 degrees were tested at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48 in the presence of gaseous plumes. An attempt was made to determine the local Mach number above the flare by utilizing a pitot probe. This objective was only partially satisfied because the 20 degree flare separated the flow ahead of the flare for Mach numbers of 0.5 to 1.96. An accurate local Mach number could not be determined because of the separated flow. To meet the objective of a data base as a function of freestream Mach number, model surface and base pressures were obtained in the presence of gaseous plumes for a matrix of chamber pressures and temperatures at Mach numbers of 0.5, 0.9, 1.2, 1.46, 1.96 and 3.48.

Love, D. A.

1978-01-01

141

Site alteration effects from rocket exhaust impingment during a simulated Viking Mars landing. Part 1: Nozzle development and physical site alternation  

NASA Technical Reports Server (NTRS)

A potential interference problem for the Viking '75 scientific investigation of the Martian surface resulting from retrorocket exhaust plume impingement of the surface was investigated experimentally and analytically. It was discovered that the conventional bell nozzle originally planned for the Viking Lander retrorockets would produce an unacceptably large amount of physical disturbance to the landing site. An experimental program was subsequently undertaken to find and/or develop a nozzle configuration which would significantly reduce the site alteration. A multiple nozzle configuration, consisting of 18 small bell nozzles, was shown to produce a level of disturbance that was considered by the Viking Lander Science Teams to be acceptable on the basis of results from full-scale tests on simulated Martian soils.

Romine, G. L.; Reisert, T. D.; Gliozzi, J.

1973-01-01

142

A tandem mirror plasma source for hybrid plume plasma studies  

NASA Technical Reports Server (NTRS)

A tandem mirror device to be considered as a hot plasma source for the hybrid plume rocket concept is discussed. The hot plamsa from this device is injected into an exhaust duct, which will interact with an annular hypersonic layer of neutral gas. The device can be used to study the dynamics of the hybrid plume, and to verify the numerical predictions obtained with computer codes. The basic system design is also geared towards low weight and compactness, and high power density at the exhaust. The basic structure of the device consists of four major subsystems: (1) an electric power supply; (2) a low temperature, high density plasma gun, such as a stream gun, an MPD source or gas cell; (3) a power booster in the form of a tandem mirror machine; and (4) an exhaust nozzle arrangement. The configuration of the tandem mirror section is shown.

Yang, T. F.; Chang, F. R.; Miller, R. H.; Wenzel, K. W.; Krueger, W. A.

1985-01-01

143

STS-31 Discovery, OV-103, rockets through low-lying clouds after KSC liftoff  

NASA Technical Reports Server (NTRS)

STS-31 Discovery, Orbiter Vehicle (OV) 103, rides above the firey glow of the solid rocket boosters (SRBs) and space shuttle main engines (SSMEs) and a long trail of exhaust as it heads toward Earth orbit. Kennedy Space Center (KSC) Launch Complex (LC) Pad 39B is covered in an exhaust cloud moments after the liftoff of OV-103 at 8:33:51.0492 am (Eastern Daylight Time (EDT)). The exhaust plume pierces the low-lying clouds as OV-103 soars into the clear skies above. A nearby waterway appears in the foreground.

1990-01-01

144

Ablative Rocket Deflector Testing and Computational Modeling  

NASA Technical Reports Server (NTRS)

A deflector risk mitigation program was recently conducted at the NASA Stennis Space Center. The primary objective was to develop a database that characterizes the behavior of industry-grade refractory materials subjected to rocket plume impingement conditions commonly experienced on static test stands. The program consisted of short and long duration engine tests where the supersonic exhaust flow from the engine impinged on an ablative panel. Quasi time-dependent erosion depths and patterns generated by the plume impingement were recorded for a variety of different ablative materials. The erosion behavior was found to be highly dependent on the material s composition and corresponding thermal properties. For example, in the case of the HP CAST 93Z ablative material, the erosion rate actually decreased under continued thermal heating conditions due to the formation of a low thermal conductivity "crystallization" layer. The "crystallization" layer produced near the surface of the material provided an effective insulation from the hot rocket exhaust plume. To gain further insight into the complex interaction of the plume with the ablative deflector, computational fluid dynamic modeling was performed in parallel to the ablative panel testing. The results from the current study demonstrated that locally high heating occurred due to shock reflections. These localized regions of shock-induced heat flux resulted in non-uniform erosion of the ablative panels. In turn, it was observed that the non-uniform erosion exacerbated the localized shock heating causing eventual plume separation and reversed flow for long duration tests under certain conditions. Overall, the flow simulations compared very well with the available experimental data obtained during this project.

Allgood, Daniel C.; Lott, Jeffrey W.; Raines, Nickey

2010-01-01

145

Application of the SAHA equation to high temperature (greater than or equal to 6000 K) rocket exhaust  

NASA Astrophysics Data System (ADS)

Using the SAHA equations and spectroscopic constants, the computer program calculates the species populations of diatomic molecules, neutrals, ions, and electrons in a plasma. The code considers the equilibrium of a two-element, chemically reacting plasma, by calculating the partition function for each species. By rederiving theoretically, as the JANNAF tables do, the thermodynamic properties of a rocket propellant, but using extended excitation levels, the program can estimate (above 6000 K) performance of advanced propulsion concepts. This second edition code described considers in addition to mixtures of carbon (C), hydrogen (H), nitrogen (N), and oxygen (O), the species of argon (Ar), up through C(sub V), H(sub II), N(sub IV), O(sub V), and Ar(sub V); any diatomic combination of these elements, both neutral and singly ionized. The user can enter in spectroscopic data for his own elements. Program results agree with JANNAF tables at 6000 K, for dominant species; output graphs of nitrogen species densities (3000 K to 30,000 K) and air species mole fractions (3000 K to 10,000 K) match published data. The second edition also has the additional feature of being able to calculate rocket performance based upon the input of a specific quantity of energy when running the code. This is especially valuable for calculating the high temperature performance of laser, fusion, and antiproton thermal propulsion systems which add heat to a working fluid or propellant.

Nachtrieb, Robert T.

1993-03-01

146

Modelling exhaust plume mixing in the near eld of an aircraft F. Garnier, S. Brunet, L. Jacquin  

E-print Network

by the system water vapour-sulphuric acid on aerosol formation in the wake. For favourable ambient relative of accurate partial vapour pressures of dierent species (mainly water and sulphuric acid) in the wake vortex exhaust contains products resulting from combustion, usually designated as major species (CO2, CO, H2O

Paris-Sud XI, Université de

147

COMPARISON OF THE PARTICLE SIZE DISTRIBUTION OF HEAVY-DUTY DIESEL EXHAUST USING A DILUTION TAIL-PIPE SAMPLER AND IN-PLUME SAMPLER DURING ON-ROAD OPERATION  

EPA Science Inventory

The paper compares the particle size distribution of heavy-duty diesel exhaust using a dilution tail-pipe sampler and an in-plume sampler during on-road operation. EPA's On-road Diesel Emissions Characterization Facility, modified to incorporate particle measurement instrumentat...

148

Reusable rocket engine optical condition monitoring  

NASA Technical Reports Server (NTRS)

Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

Wyett, L.; Maram, J.; Barkhoudarian, S.; Reinert, J.

1987-01-01

149

Quick Access Rocket Exhaust Rig Testing of Coated GRCop-84 Sheets Used to Aid Coating Selection for Reusable Launch Vehicles  

NASA Technical Reports Server (NTRS)

The design of the next generation of reusable launch vehicles calls for using GRCop-84 copper alloy liners based on a composition1 invented at the NASA Glenn Research Center: Cu-8(at.%)Cr-4%Nb. Many of the properties of this alloy have been shown to be far superior to those of other conventional copper alloys, such as NARloy-Z. Despite this considerable advantage, it is expected that GRCop-84 will suffer from some type of environmental degradation depending on the type of rocket fuel utilized. In a liquid hydrogen (LH2), liquid oxygen (LO2) booster engine, copper alloys undergo repeated cycles of oxidation of the copper matrix and subsequent reduction of the copper oxide, a process termed "blanching". Blanching results in increased surface roughness and poor heat-transfer capabilities, local hot spots, decreased engine performance, and premature failure of the liner material. This environmental degradation coupled with the effects of thermomechanical stresses, creep, and high thermal gradients can distort the cooling channel severely, ultimately leading to its failure.

Raj, Sai V.; Robinson, Raymond C.; Ghosn, Louis J.

2005-01-01

150

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

NASA Technical Reports Server (NTRS)

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

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

1978-01-01

151

Nozzle exit exhaust products from space shuttle boost vehicle (November 1973 design)  

NASA Technical Reports Server (NTRS)

Principal exhaust species emitted at various altitudes for two trajectories of the space shuttle vehicle are presented. The exhaust composition is given for the nozzle exit plane on the basis of equilibrium chemistry. Afterburning of excess H, H2, and CO in the plume is accounted for. Species considered include HCl and Al2O3, which have been recognized as environmentally significant, as well as others such as H2O (produced by both the solid rocket motor and the orbiter main engine) which, although innocuous, may participate in subsequent chemical reactions in the atmosphere.

1975-01-01

152

Liquid propellant rockets.  

NASA Technical Reports Server (NTRS)

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

Dipprey, D. F.

1972-01-01

153

Delta 2 Explosion Plume Analysis Report  

NASA Technical Reports Server (NTRS)

A Delta II rocket exploded seconds after liftoff from Cape Canaveral Air Force Station (CCAFS) on 17 January 1997. The cloud produced by the explosion provided an opportunity to evaluate the models which are used to track potentially toxic dispersing plumes and clouds at CCAFS. The primary goal of this project was to conduct a case study of the dispersing cloud and the models used to predict the dispersion resulting from the explosion. The case study was conducted by comparing mesoscale and dispersion model results with available meteorological and plume observations. This study was funded by KSC under Applied Meteorology Unit (AMU) option hours. The models used in the study are part of the Eastern Range Dispersion Assessment System (ERDAS) and include the Regional Atmospheric Modeling System (RAMS), HYbrid Particle And Concentration Transport (HYPACT), and Rocket Exhaust Effluent Dispersion Model (REEDM). The primary observations used for explosion cloud verification of the study were from the National Weather Service's Weather Surveillance Radar 1988-Doppler (WSR-88D). Radar reflectivity measurements of the resulting cloud provided good estimates of the location and dimensions of the cloud over a four-hour period after the explosion. The results indicated that RAMS and HYPACT models performed reasonably well. Future upgrades to ERDAS are recommended.

Evans, Randolph J.

2000-01-01

154

ICOARE: Impacts on Climate and Ozone from Aircraft and Rocket Emissions  

NASA Astrophysics Data System (ADS)

This presentation will provide an overview of an Earth Venture proposal for a series of in situ measurements in the exhaust plumes of aircraft and rockets with the following objectives: to obtain information that is critical for reducing the uncertainties in assessments (e.g., WMO and IPCC) of the impacts of aviation and aerospace activities on regional and global climate; to assess the viability of a climate engineering scheme that employs injection of reflective particles into the lower stratosphere; and to initiate the development of an operational modeling tool that can be used by the aviation and aerospace industries to guide design of new transporation systems that minimize the impact on Earth’s climate. The ICOARE mission will deploy instruments to measure water vapor, ice water content, tracers, reactive species, particles, and radiation fields on a high-altitude aircraft to characterize the variability of water vapor in aircraft and rocket contrails, determine accurate emission indices for initialization of plume-wake and regional scale models, investigate the microphysical properties of cirrus particles in and out of aircraft corridors, and examine the light scattering properties of contrail ice crystals and small alumina particles. Focused campaigns will be timed to occur around the launch schedules of a variety of rocket types in order to characterize the range of emissions from the current launch suite. There will be special emphasis on characterizing the emissions from rockets employing new propellants, in particular those that may produce soot and nitrogen oxides. Observations in aircraft exhaust, and examinations of cirrus cloud properties and persistent contrails, will occur on flights that are not dedicated to studies of rockets (e.g., test, transit, and rocket-scrub flights). ICOARE will offer a unique opportunities for training students and postdoctorates, especially those from underrepresented groups, in areas of project management, logistics, instrument development, data acquisition and analysis, and education and outreach.

Toohey, D. W.; Ross, M.

2009-12-01

155

Position paper on the potential of inadvertent weather modification of the Florida Peninsula resulting from neutralization of space shuttle solid rocket booster exhaust clouds  

NASA Technical Reports Server (NTRS)

A concept of injecting compounds into the exhaust cloud was proposed to neutralize the acidic nature of the low-level stabilized ground cloud (SGC) was studied. The potential Inadvertent Weather Modification caused by exhaust cloud characteristics from three hours to seven days after launch was studied. Possible effects of the neutralized SGC in warm and cloud precipitation processes were discussed. Based on a detailed climatology of the Florida Peninsula, the risk for weather modification under a variety of weather situations was assessed.

Bollay, E.; Bosart, L.; Droessler, E.; Jiusto, J.; Lala, G. G.; Mohnen, V.; Schaefer, V.; Squires, P.

1979-01-01

156

Particle Rotation Effects in Rarefied Two-Phase Plume Flows  

NASA Astrophysics Data System (ADS)

We evaluate the effects of solid particle rotation in high-altitude solid rocket exhaust plume flows, through the development and application of methods for the simulation of two phase flows involving small rotating particles and a nonequilibrium gas. Green's functions are derived for the force, moment, and heat transfer rate to a rotating solid sphere within a locally free-molecular gas, and integration over a Maxwellian gas velocity distribution is used to determine the influence of particle rotation on the heat transfer rate at the equilibrium limit. The use of these Green's functions for the determination of particle phase properties through the Direct Simulation Monte Carlo method is discussed, and a procedure is outlined for the stochastic modeling of interphase collisions. As a test case, we consider the nearfield plume flow for a Star-27 solid rocket motor exhausting into a vacuum, and vary particle angular velocities at the nozzle exit plane in order to evaluate the influence of particle rotation on various flow properties. Simulation results show that rotation may lead to slightly higher particle temperatures near the central axis, but for the case considered the effects of particle rotation are generally found to be negligible.

Burt, Jonathan M.; Boyd, Iain D.

2005-05-01

157

Overview of NASA GRCs Green Propellant Infusion Mission Thruster Testing and Plume Diagnostics  

NASA Technical Reports Server (NTRS)

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The models describe the pressure, temperature, density, Mach number, and species concentration of the AF-M315E thruster exhaust plumes. The models are being used to assess the impingement effects of the AF-M315E thrusters on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters will be tested in a small rocket, altitude facility at NASA GRC. The GRC thruster testing will be conducted at duty cycles representatives of the planned GPIM maneuvers. A suite of laser-based diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, Schlieren imaging, and physical probes will be used to acquire plume measurements of AFM315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

Deans, Matthew C.; Reed, Brian D.; Yim, John T.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; McLean, Christopher H.

2014-01-01

158

Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix B: Liquid rocket booster acoustic and thermal environments  

NASA Technical Reports Server (NTRS)

The ascent thermal environment and propulsion acoustic sources for the Martin-Marietta Corporation designed Liquid Rocket Boosters (LRB) to be used with the Space Shuttle Orbiter and External Tank are described. Two designs were proposed: one using a pump-fed propulsion system and the other using a pressure-fed propulsion system. Both designs use LOX/RP-1 propellants, but differences in performance of the two propulsion systems produce significant differences in the proposed stage geometries, exhaust plumes, and resulting environments. The general characteristics of the two designs which are significant for environmental predictions are described. The methods of analysis and predictions for environments in acoustics, aerodynamic heating, and base heating (from exhaust plume effects) are also described. The acoustic section will compare the proposed exhaust plumes with the current SRB from the standpoint of acoustics and ignition overpressure. The sections on thermal environments will provide details of the LRB heating rates and indications of possible changes in the Orbiter and ET environments as a result of the change from SRBs to LRBs.

1989-01-01

159

Ionospheric hole made by the 2012 North Korean rocket observed with a dense GNSS array in Japan  

NASA Astrophysics Data System (ADS)

dense array of Global Navigation Satellite System (GNSS) receivers is useful to study ionospheric disturbances. Here we report observations by a Japanese GNSS array of an ionospheric hole, i.e., localized electron depletion, made by water vapor molecules in the exhaust plume of the second-stage engine of the Unha-3 rocket launched from North Korea, on 12 December 2012. The Russian GNSS was used for the first time to observe such an ionospheric hole. The hole emerged ~6 min after the launch above the middle of the Yellow Sea, and its size and depth suggest that the Unha-3 is slightly less powerful than the 2009 Taepodong-2 missile, also from North Korea. Smaller-scale electron depletion signatures appeared ~10 min after the launch above the southern East China Sea, which is possibly caused by the exhaust plume of the third-stage engine.

Nakashima, Yuki; Heki, Kosuke

2014-07-01

160

Safe testing nuclear rockets economically  

SciTech Connect

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

Howe, S. D. (Steven D.); Travis, B. J. (Bryan J.); Zerkle, D. K. (David K.)

2002-01-01

161

Pop Rockets  

NSDL National Science Digital Library

Students design and build paper rockets around film canisters, which serve as engines. An antacid tablet and water are put into each canister, reacting to form carbon dioxide gas, and acting as the pop rocket's propellant. With the lid snapped on, the continuous creation of gas causes pressure to build up until the lid pops off, sending the rocket into the air. The pop rockets demonstrate Newton's third law of motion: for every action, there is an equal and opposite reaction.

Integrated Teaching And Learning Program

162

Plume Impingement on a Dusty Lunar Surface  

NASA Astrophysics Data System (ADS)

A loosely coupled continuum-DSMC solver is used to simulate the interaction between the exhaust from a rocket engine with the lunar surface. This problem is of particular interest because the high velocity dust spray can damage nearby structures. The flow field is challenging to simulate because continuum assumptions are no longer valid in the far field, while in the near field DSMC becomes impractical because of the high collision rate. In the current work the high density core of the rocket plume is modeled with NASA's continuum flow solver, DPLR [1]. Since the two solvers are loosely coupled, i.e. one-way coupling from the DPLR to the DSMC regimes, the interface between the two solvers is placed in the supersonic region above the surface shock. At the lunar surface, a boundary layer develops and the shear stress causes dust grains to slide and eventually enter the flow field. Robert's theory of dust entrainment [2,3] is used to predict how much dust is lofted into the flow field by the near surface flow conditions. In Robert's original theory the interaction between entrained dust grains and the gas was neglected and the particles were assumed to follow ballistic trajectories. In our current model, the dust grains are coupled with the DSMC gas model. Both the dust trajectories and the flow fields are computed for various hovering altitudes and dust grain sizes. Comparisons are made to Robert's original predictions and Apollo photogrammetry [4].

Morris, A. B.; Goldstein, D. B.; Varghese, P. L.; Trafton, L. M.

2011-05-01

163

Supersonic Rocket Thruster Flow Predicted by Numerical Simulation  

NASA Technical Reports Server (NTRS)

Despite efforts in the search for alternative means of energy, combustion still remains the key source. Most propulsion systems primarily use combustion for their needed thrust. Associated with these propulsion systems are the high-velocity hot exhaust gases produced as the byproducts of combustion. These exhaust products often apply uneven high temperature and pressure over the surfaces of the appended structures exposed to them. If the applied pressure and temperature exceed the design criteria of the surfaces of these structures, they will not be able to protect the underlying structures, resulting in the failure of the vehicle mission. An understanding of the flow field associated with hot exhaust jets and the interactions of these jets with the structures in their path is critical not only from the design point of view but for the validation of the materials and manufacturing processes involved in constructing the materials from which the structures in the path of these jets are made. The hot exhaust gases often flow at supersonic speeds, and as a result, various incident and reflected shock features are present. These shock structures induce abrupt changes in the pressure and temperature distribution that need to be considered. In addition, the jet flow creates a gaseous plume that can easily be traced from large distances. To study the flow field associated with the supersonic gases induced by a rocket engine, its interaction with the surrounding surfaces, and its effects on the strength and durability of the materials exposed to it, NASA Glenn Research Center s Combustion Branch teamed with the Ceramics Branch to provide testing and analytical support. The experimental work included the full range of heat flux environments that the rocket engine can produce over a flat specimen. Chamber pressures were varied from 130 to 500 psia and oxidizer-to-fuel ratios (o/f) were varied from 1.3 to 7.5.

Davoudzadeh, Farhad

2004-01-01

164

Prediction of Acoustic Environments from Horizontal Rocket Firings  

NASA Technical Reports Server (NTRS)

In recent years, advances in research and engineering have led to more powerful launch vehicles which can reach areas of space not yet explored. These more powerful vehicles yield acoustic environments potentially destructive to the vehicle or surrounding structures. Therefore, it has become increasingly important to be able to predict the acoustic environments created by these vehicles in order to avoid structural and/or competent failure. The current industry standard technique for predicting launch-induced acoustic environments was developed by Eldred in the early 1970's and is published in NASA SP-80721. Recent work2 has shown Eldred's technique to be inaccurate for current state-of-the-art launch vehicles. Due to the high cost of full-scale and even sub-scale rocket experiments, very little rocket noise data is available. Furthermore, much of the work thought to be applicable to rocket noise has been done with heated jets. Tam3,4 has done an extensive amount of research on jets of different nozzle exit shape, diameter, velocity, and temperature. Though the values of these parameters, especially exit velocity and temperature, are often very low compared to these values in rockets, a lot can be learned about rocket noise from jet noise literature. The turbulent nature of jet and rocket exhausts is quite similar. Both exhausts contain turbulent structures of varying scale-termed the fine and large scale turbulence by Tam. The finescale turbulence is due to small eddies from the jet plume interacting with the ambient atmosphere. According to Tam et al., the noise radiated by this envelope of small-scale turbulence is statistically isotropic. Hence, one would expect the noise from the small scale turbulence of the jet to be nearly omni-directional. The coherent nature of the large-scale turbulence results in interference of the noise radiated from different spatial locations within the jet. This interference-whether it is constructive or destructive-results in highly directional noise radiation. Tam3 has proposed a model to predict the acoustic environment due to jets and while it works extremely well for jets, it was found to be inappropriate for rockets8. A model to predict the acoustic environment due to a launch vehicle in the far-field which incorporates concepts from both Eldred and Tam was created. This was done using five sets of horizontally fired rocket data, obtained between 2008 and 2012. Three of these rockets use solid propellant and two use liquid propellant. Through scaling analysis, it is shown that liquid and solid rocket motors exhibit similar spectra at similar amplitudes. This model is accurate for these five data sets within 5 dB of the measured data for receiver angles of 30deg to 160deg (with respect to the downstream exhaust centerline). The model uses the following vehicle parameters: nozzle exit diameter and velocity, radial distance from source to receiver, receiver angle, mass flow rate, and acoustic efficiency.

Giacomoni, Clothilde

2014-01-01

165

Flight Investigation to Determine the Effect of Jet Exhaust on Drag, Trim Characteristics, and Afterbody Pressures of a 0.125-Scale Rocket Model of the Mcdonnell F-101A Airplane  

NASA Technical Reports Server (NTRS)

A flight investigation was conducted to determine the effect of jet exhaust on the drag, trim characteristics, and afterbody pressures on a 0.125-scale rocket model of the McDonnell F-101A airplance. Power-off data were obtained over a Mach number range of 1.04 to 1.9 and power-on data were obtained at a Mach number of about 1.5. The data indicated that with power-on the change in external drag coefficient was within the data accuracy and there was a decrease in trim angle of attack of 1.27 degrees with a corresponding decrease of 0.07 in lift coefficient. Correspondingly, pressure coefficients on the side and bottom of the fuselage indicated a positive increment near the jet exit. As the distance downstream of the jet exit increased, the increment on the bottom of the fuselage increased, whereas the increments on the side decreased to a negative peak.

Kennedy, Thomas L.

1956-01-01

166

JOURNAL OF SPACECRAFT AND ROCKETS Vol. 42, No. 4, JulyAugust 2005  

E-print Network

JOURNAL OF SPACECRAFT AND ROCKETS Vol. 42, No. 4, July­August 2005 Lidar Backscatter Properties of Al2O3 Rocket Exhaust Particles Robert A. Reed and Mitchel K. Nolen Jacobs--Sverdrup Corporation sections of aluminum-oxide rocket exhaust particles are sensitive to slight deviations from their nominal

167

Ionospheric modification by rocket effluents. Final report  

SciTech Connect

This report describes experimental and theoretical studies related to ionospheric disturbances produced by rocket exhaust vapors. The purpose of our research was to estimate the ionospheric effects of the rocket launches which will be required to place the Satellite Power System (SPS) in operation. During the past year, we have developed computational tools for numerical simulation of ionospheric changes produced by the injection of rocket exhaust vapors. The theoretical work has dealt with (1) the limitations imposed by condensation phenomena in rocket exhaust; (2) complete modeling of the ionospheric depletion process including neutral gas dynamics, plasma physics, chemistry and thermal processes; and (3) the influence of the modified ionosphere on radio wave propagation. We are also reporting on electron content measurements made during the launch of HEAO-C on Sept. 20, 1979. We conclude by suggesting future experiments and areas for future research.

Bernhardt, P.A.; Price, K.M.; da Rosa, A.V.

1980-06-01

168

Rocket Principles  

NSDL National Science Digital Library

On this site from the NASA Glenn Research Center Learning Technologies Project, the science and history of rocketry is explained. Visitors will find out how rocket principles illustrate Newton's Laws of Motion. There is a second page of this site, Practical Rocketry, which discusses the workings of rockets, including propellants, engine thrust control, stability and control systems, and mass.

2008-07-29

169

Foam Rocket  

NSDL National Science Digital Library

In this activity, learners work in teams build and launch rubberband-powered foam rockets. Through a controlled investigation, learners will explore rocket stability and the trajectory relationship between launch angle and range. This lesson plan includes background information, diagrams, and handouts for learners.

Shearer, Deborah A.; Gregory L. Vogt, Ed D.

2012-06-26

170

Solid rocket booster thermal radiation model, volume 1  

NASA Technical Reports Server (NTRS)

A solid rocket booster (SRB) thermal radiation model, capable of defining the influence of the plume flowfield structure on the magnitude and distribution of thermal radiation leaving the plume, was prepared and documented. Radiant heating rates may be calculated for a single SRB plume or for the dual SRB plumes astride the space shuttle. The plumes may be gimbaled in the yaw and pitch planes. Space shuttle surface geometries are simulated with combinations of quadric surfaces. The effect of surface shading is included. The computer program also has the capability to calculate view factors between the SRB plumes and space shuttle surfaces as well as surface-to-surface view factors.

Watson, G. H.; Lee, A. L.

1976-01-01

171

POD Analysis of Jet-Plume/Afterbody-Wake Interaction  

NASA Astrophysics Data System (ADS)

The understanding of the flow physics in the base region of a powered rocket is one of the keys to designing the next generation of reusable launchers. The base flow features affect the aerodynamics and the heat loading at the base of the vehicle. Recent efforts at the National Center for Physical Acoustics at the University of Mississippi have refurbished two models for studying jet-plume/afterbody-wake interactions in the NCPA's 1-foot Tri-Sonic Wind Tunnel Facility. Both models have a 2.5 inch outer diameter with a nominally 0.5 inch diameter centered exhaust nozzle. One of the models is capable of being powered with gaseous H2 and O2 to study the base flow in a fully combusting senario. The second model uses hi-pressure air to drive the exhaust providing an unheated representative flow field. This unheated model was used to acquire PIV data of the base flow. Subsequently, a POD analysis was performed to provide a first look at the large-scale structures present for the interaction between an axisymmetric jet and an axisymmetric afterbody wake. PIV and Schlieren data are presented for a single jet-exhaust to free-stream flow velocity along with the POD analysis of the base flow field.

Murray, Nathan E.; Seiner, John M.; Jansen, Bernard J.; Gui, Lichuan; Sockwell, Shuan; Joachim, Matthew

2009-11-01

172

Engineering Issues of Iridium Coated Rhenium Rockets  

Microsoft Academic Search

The key to the performance and lifetime of radiation-cooled rockets is the chamber temperature capability. Temperature limitations (1370°C) of current state-of-art chamber materials force the use of film cooling, which degrades rocket performance and imposes plume contamination from unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium (Ir) coating has demonstrated operation, for long lifetimes

Brian D. Reed; James A. Biaglow; Steven J. Schneider

1998-01-01

173

Rocket Engines  

NSDL National Science Digital Library

This video from SpaceTEC National Aerospace Technical Education Center explains the theory of rocket engines using Newton's third law of motion. This five minute video is one of the aerospace certification readiness courses.

2011-07-27

174

Stomp Rocket  

NSDL National Science Digital Library

In this activity, learners build rockets and shoot them into the air by stomping on the plastic bottle launchers. Use this activity to demonstrate air pressure, projectile motion, Newton's Laws of Motion, and vectors.

Workshop, Fresno C.

2012-01-01

175

Rocket Launchers  

NSDL National Science Digital Library

In this activity, learners work with an adult to build a rocket and launcher out of a plastic 2-liter bottle, flexible plastic hose, plastic tubing, toilet paper tube, and duct tape. Use this stomp rocket activity to demonstrate that air is something, comprised of molecules that, when acted upon, have the power to move things. This activity guide includes an extension activity and related activity for younger learners.

Museum, Chicago C.

2010-01-01

176

Methylhydrazinium nitrate. [rocket plume deposit chemistry  

NASA Technical Reports Server (NTRS)

Methylhydrazinium nitrate was synthesized by the reaction of dilute nitric acid with methylhydrazine in water and in methanol. The white needles formed are extremely hygroscopic and melt at 37.5-40.5 C. The IR spectrum differs from that reported elsewhere. The mass spectrum exhibited no parent peak at 109 m/z, and thermogravimetric analysis indicated that the compound decomposed slowly at 63-103 C to give ammonium and methylammonium nitrate. The density is near 1.55 g/cu cm.

Lawton, E. A.; Moran, C. M.

1983-01-01

177

JOURNAL OF SPACECRAFT AND ROCKETS Vol. 41, No. 4, JulyAugust 2004  

E-print Network

JOURNAL OF SPACECRAFT AND ROCKETS Vol. 41, No. 4, July­August 2004 Modeling of Chemically Reacting) thruster positioned on the side of a small rocket, with the rarefied atmosphere between altitudes of 80 inside the nozzle as well as the rocket speed (5­8 km/s) and flight altitude (80­160 km) on the plume

Alexeenko, Alina

178

Air-Powered Rockets.  

ERIC Educational Resources Information Center

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)

Rodriguez, Charley; Raynovic, Jim

179

Space shuttle plume simulation application. Results and math model. [Ames unitary plan wind tunnel test  

NASA Technical Reports Server (NTRS)

Pressure and gauge wind tunnel data from a transonic test of a 0.02 scale model of the space shuttle launch vehicle was analyzed to define the aerodynamic influence of the main propulsion system and solid rocket booster plumes during the transonic portion of ascent flight. Air was used as a simulant gas to develop the model exhaust plumes. A math model of the plume induced aerodynamic characteristics was developed for a range of Mach numbers to match the forebody aerodynamic math model. The base aerodynamic characteristics are presented in terms of forces and moments versus attitude. Total vehicle base and forebody aerodynamic characteristics are presented in terms of aerodynamic coefficients for Mach number from 0.6 to 1.4 Element and component base and forebody aerodynamic characteristics are presented for Mach numbers of 0.6, 1.05, 1.1, 1.25 and 1.4. The forebody data is available at Mach 1.55. Tolerances for all plume induced aerodynamic characteristics are developed in terms of a math model.

Boyle, W.; Conine, B.

1978-01-01

180

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines  

NASA Technical Reports Server (NTRS)

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 effects in the nozzle flowfield. Additionally, comparisons of the model results to performance data from CalTech, as well as experimental flowfield measurements from Stanford University, are also reported.

Morris, Christopher I.

2005-01-01

181

Measuring Fluctuating Pressure Levels and Vibration Response in a Jet Plume  

NASA Technical Reports Server (NTRS)

The characterization of loads due to solid rocket motor plume impingement allows for moreaccurate analyses of components subjected to such an environment. Typically, test verification of predicted loads due to these conditions is widely overlooked or unsuccessful. ATA Engineering, Inc., performed testing during a solid rocket motor firing to obtain acceleration and pressure responses in the hydrodynamic field surrounding the jet plume. The test environment necessitated a robust design to facilitate measurements being made in close proximity to the jet plume. This paper presents the process of designing a test fixture and an instrumentation package that could withstand the solid rocket plume environment and protect the required instrumentation.

Osterholt, Douglas J.; Knox, Douglas M.

2011-01-01

182

Vertical Landing Aerodynamics of Reusable Rocket Vehicle  

NASA Astrophysics Data System (ADS)

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

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

183

Rockets Away  

NSDL National Science Digital Library

In this activity, learners build a simple "rocket" with ordinary household materials to demonstrate the basic principles behind rocketry and the principle of reaction. This activity can be completed indoors or experimented with outdoors using a much longer piece of string.

Space Sciences Laboratory, Uc B.

2001-01-01

184

Rubberband Rockets  

NSDL National Science Digital Library

This fun and simple activity is a rubberband rocket design challenge! Learners will explore how tail fins can help to stabilize a flying object, while also exploring potential and kinetic energy. The activity, which can be used alone or as part of a visit to COSI, is located on page 10 of COSI's Force and Motion Teacher Guide.

Cosi

2009-01-01

185

Balloon Rocket  

NSDL National Science Digital Library

Experiment with force and pressure by building a balloon rocket. When launched, the balloon will run a track wherever you place the string. All you need is a balloon, clothespin, a straw, some tape, and some string, then get ready for take off!

Minnesota, Science M.

1995-01-01

186

Relativistic rocket and space flight  

NASA Astrophysics Data System (ADS)

We introduce the new concept of "proper speed". It turns out that relativistic space flight can be sized and described more adequately by this concept. In particular, it is shown that the physics and equations of relativistic and classical space flight have a one-to-one relationship. The rocket equations are even identical. In addition, we generalize the expressions for the rocket exhaust speed to systems with mass and/or photon propulsion. With this new insight the basic relativistic spaceflight are easily derived and summarized and it is shown that they can be readily transformed into those of classical space flight. We also provide the space-time transformation equations for a cruising rocket, for a rocket with constant acceleration, and for a rocket with constant thrust (with and without inflight thrust reversal to slow down to zero speed at the target). For each of these cases a representative flight through the milky way with a ultimate matter-antimatter annihilation drive is calculated.

Walter, Ulrich

2006-09-01

187

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

NASA Technical Reports Server (NTRS)

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

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

1993-01-01

188

Analysis of a Nuclear Enhanced Airbreathing Rocket for Earth to Orbit Applications  

NASA Technical Reports Server (NTRS)

The proposed engine concept is the Nuclear Enhanced Airbreathing Rocket (NEAR). The NEAR concept uses a fission reactor to thermally heat a propellant in a rocket plenum. The rocket is shrouded, thus the exhaust mixes with ingested air to provide additional thermal energy through combustion. The combusted flow is then expanded through a nozzle to provide thrust.

Adams, Robert B.; Landrum, D. Brian; Brown, Norman (Technical Monitor)

2001-01-01

189

Supplemental final environmental impact statement for advanced solid rocket motor testing at Stennis Space Center  

NASA Technical Reports Server (NTRS)

Since the Final Environmental Impact Statement (FEIS) and Record of Decision on the FEIS describing the potential impacts to human health and the environment associated with the program, three factors have caused NASA to initiate additional studies regarding these issues. These factors are: (1) The U.S. Army Corps of Engineers and the Environmental Protection Agency (EPA) agreed to use the same comprehensive procedures to identify and delineate wetlands; (2) EPA has given NASA further guidance on how best to simulate the exhaust plume from the Advanced Solid Rocket Motor (ASRM) testing through computer modeling, enabling more realistic analysis of emission impacts; and (3) public concerns have been raised concerning short and long term impacts on human health and the environment from ASRM testing.

1990-01-01

190

Rocket Pinwheel  

NSDL National Science Digital Library

This is an activity about motion, power, air and Newtonâs Third Law of Motion, which states that for every action there is an equal and opposite reaction. Learners will harness the power of thrust forces to build a rocket pinwheel. They will do this by making a pinwheel with a balloon, straw and pin. Thrust causes the balloon to spin around in a circular motion.

Center, Reuben H.

1999-01-01

191

Reusable Rockets  

NSDL National Science Digital Library

This activity (located on page 3 of the PDF) is a full inquiry investigation into design optimization using recycled materials. Groups of learners will design a study that seeks out the relationship between the amount of water carried by rockets built from used soda bottles and their total flight times, ultimately graphing their results to determine a point of diminishing return. Relates to linked video, DragonflyTV GPS: Garbology.

Twin Cities Public Television, Inc.

2007-01-01

192

Atomic hydrogen rocket engine  

NASA Technical Reports Server (NTRS)

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

Etters, R. D.; Flurchick, K.

1981-01-01

193

Performance prediction of a ducted rocket combustor  

Microsoft Academic Search

The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on Computational Fluid Dynamics (CFD) is increasingly popular, but their application to reacting flow requires specific attention and validation. Through a

Robert Stowe

2001-01-01

194

Maximum terminal velocity of relativistic rocket  

Microsoft Academic Search

The maximum terminal velocity problem of the classical propulsion is extended to a relativistic rocket assumed broken down into active mass, inert mass and gross payload. A fraction of the active mass is converted into energy shared between inert mass and active mass residual. Significant effects are considered. State and co-state equations are carried out to find the exhaust speed

G. Vulpetti

1985-01-01

195

Plasma Diagnostics Development for Advanced Rocket Engines  

NASA Astrophysics Data System (ADS)

The VASIMR (Variable Specific Impulse Magnetoplasma Rocket) engine is a next-generation rocket engine under development at the Johnson Space Center's Advanced Space Propulsion Laboratory. With an exhaust velocity up to 50 times that of chemical rocket engines such as the Space Shuttle Main Engine, the VASIMR concept promises fast, efficient interplanetary flight. Rice University has participated in VASIMR research since 1996 and at present is developing two new diagnostic probes: a retarding potential analyzer to measure the velocity of ions in the rocket's exhaust, and a moveable optical probe to examine the spectrum of the rocket's helicon plasma source. In support of the probe development, a test facility is under construction at Rice, consisting of a small electric rocket engine firing into a 2-m vacuum chamber. This engine, the MPD (magnetoplasmadynamic) thruster, dates from the 1960's and provides a well-characterized source plasma for testing of the probes under development. We present details of the ion energy analyzer and the facility under construction at Rice.

Glover, Timothy; Kittrell, Carter; Chan, Anthony; Chang-Diaz, Franklin

2000-10-01

196

Space shuttle exhaust cloud properties  

NASA Technical Reports Server (NTRS)

A data base describing the properties of the exhaust cloud produced by the launch of the Space Transportation System and the acidic fallout observed after each of the first four launches was assembled from a series of ground and aircraft based measurements made during the launches of STS 2, 3, and 4. Additional data were obtained from ground-based measurements during firings of the 6.4 percent model of the Solid Rocket Booster at the Marshall Center. Analysis indicates that the acidic fallout is produced by atomization of the deluge water spray by the rocket exhaust on the pad followed by rapid scavening of hydrogen chloride gas aluminum oxide particles from the Solid Rocket Boosters. The atomized spray is carried aloft by updrafts created by the hot exhaust and deposited down wind. Aircraft measurements in the STS-3 ground cloud showed an insignificant number of ice nuclei. Although no measurements were made in the column cloud, the possibility of inadvertent weather modification caused by the interaction of ice nuclei with natural clouds appears remote.

Anderson, B. J.; Keller, V. W.

1983-01-01

197

Bottle Rockets Mechanical Engineering  

E-print Network

Bottle Rockets Mechanical Engineering Objective This lesson introduces students to forces and how they affect the maximum height of a rocket. In this activity, students will build a rocket, launch it can be used to determine the height of object 2. That the more pressure applied to a rocket will make

Provancher, William

198

Dispersion of turbojet engine exhaust in flight  

NASA Technical Reports Server (NTRS)

The dispersion of the exhaust of turbojet engines into the atmosphere is estimated by using a model developed for the mixing of a round jet with a parallel flow. The analysis is appropriate for determining the spread and dilution of the jet exhaust from the engine exit until it is entrained in the aircraft trailing vortices. Chemical reactions are not expected to be important and are not included in the flow model. Calculations of the dispersion of the exhaust plumes of three aircraft turbojet engines with and without afterburning at typical flight conditions are presented. Calculated average concentrations for the exhaust plume from a single engine jet fighter are shown to be in good agreement with measurements made in the aircraft wake during flight.

Holdeman, J. D.

1973-01-01

199

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

NASA Astrophysics Data System (ADS)

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

Lehman, Matthew Kurt

200

Plume detachment from a magnetic nozzle  

SciTech Connect

High-powered electric propulsion thrusters utilizing a magnetized plasma require that plasma exhaust detach from the applied magnetic field in order to produce thrust. This paper presents experimental results demonstrating that a sufficiently energetic and flowing plasma can indeed detach from a magnetic nozzle. Microwave interferometer and probe measurements provide plume density, electron temperature, and ion flux measurements in the nozzle region. Measurements of ion flux show a low-beta plasma plume which follows applied magnetic field lines until the plasma kinetic pressure reaches the magnetic pressure and a high-beta plume expanding ballistically afterward. Several magnetic configurations were tested including a reversed field nozzle configuration. Despite the dramatic change in magnetic field profile, the reversed field configuration yielded little measurable change in plume trajectory, demonstrating the plume is detached. Numerical simulations yield density profiles in agreement with the experimental results.

Deline, Christopher A. [University of Michigan, Ann Arbor, Michigan 48109 (United States); Bengtson, Roger D.; Breizman, Boris N.; Tushentsov, Mikhail R. [Institute for Fusion Studies, University of Texas at Austin, Austin, Texas 78712 (United States); Jones, Jonathan E.; Chavers, D. Greg; Dobson, Chris C. [Marshall Space Flight Center, Huntsville, Alabama 35805 (United States); Schuettpelz, Branwen M. [University of Alabama at Huntsville, Huntsville, Alabama 35899 (United States)

2009-03-15

201

The plume impingement test program at AEDC utilizing the S-2 ullage motors (November 1973), section 1  

NASA Technical Reports Server (NTRS)

Proposed experiments for analyzing rocket plumes are reported. Two groups of experiments were studied: (1) those that would help define some of the parameters that characterize the plume and (2) those that would enable evaluation of some of the contamination effects of the plume environment on various items of interest. The items investigated, the purpose of the investigation, are given in tabular form.

1976-01-01

202

Gas Emission Measurements from the RD 180 Rocket Engine  

NASA Technical Reports Server (NTRS)

The Science Laboratory operated by GB Tech was tasked by the Environmental Office at the NASA Marshall Space Flight Center (MSFC) to collect rocket plume samples and to measure gaseous components and airborne particulates from the hot test firings of the Atlas III/RD 180 test article at MSFC. This data will be used to validate plume prediction codes and to assess environmental air quality issues.

Ross, H. R.

2001-01-01

203

Chemical conversion of subsonic aircraft emissions in the dispersing plume: Calculation of effective emission indices  

Microsoft Academic Search

A box model representative for a mesoscale volume and three different plume models are used to estimate the chemical conversion of exhaust species of a subsonic aircraft at cruise altitude. Clearly deviating results have been obtained for instantaneous mixing of the exhaust in a box and gradual dispersion of a plume. The effect of varying daytime of release as well

H. Petry; J. Hendricks; M. Möllhoff; E. Lippert; A. Meier; A. Ebel; R. Sausen

1998-01-01

204

Plume Busters  

NSDL National Science Digital Library

Environmental and earth science students seldom have an opportunity to apply what they learn in class to the solution of real-world problems. With NSF support we have developed the prototype Plume Busters software, in which students take on the role of an environmental consultant. Following a pipeline spill, the environmental consultant is hired by the pipeline owner to locate the resulting plume created by the spill and remediate the contaminated aquifer at minimum monetary and time cost. The contamination must be removed from the aquifer before it reaches the river and eventually a downstream public water supply. The software consists of an interactive Java application and accompanying HTML linked pages. The application simulates movement of a plume from a pipeline break through a shallow alluvial aquifer towards the river. The accompanying web pages establish the simulated contamination scenario and provide students with background material on ground-water flow and transport principles. To make the role-play more realistic, the student must consider cost and time when making decisions about siting observation wells and wells for the pump-and-treat remediation system.

Macfarlane, Allen

205

Crater Formation Due to Lunar Plume Impingement  

NASA Technical Reports Server (NTRS)

Thruster plume impingement on a surface comprised of small, loose particles may cause blast ejecta to be spread over a large area and possibly cause damage to the vehicle. For this reason it is important to study the effects of plume impingement and crater formation on surfaces like those found on the moon. Lunar soil, also known as regolith, is made up of fine granular particles on the order of 100 microns.i Whenever a vehicle lifts-off from such a surface, the exhaust plume from the main engine will cause the formation of a crater. This crater formation may cause laterally ejected mass to be deflected and possibly damage the vehicle. This study is a first attempt at analyzing the dynamics of crater formation due to thruster exhaust plume impingement during liftoff from the moon. Though soil erosion on the lunar surface is not considered, this study aims at examining the evolution of the shear stress along the lunar surface as the engine fires. The location of the regions of high shear stress will determine where the crater begins to form and will lend insight into how big the crater will be. This information will help determine the probability that something will strike the vehicle. The final sections of this report discuss a novel method for studying this problem that uses a volume of fluid (VOF)ii method to track the movement of both the exhaust plume and the eroding surface.

Marsell, Brandon

2011-01-01

206

Liquid rocket engine nozzles  

NASA Technical Reports Server (NTRS)

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

1976-01-01

207

Brassicaceae (Mustard family) Yellow rocket  

E-print Network

Brassicaceae (Mustard family) Yellow rocket Barbarea vulgaris R. Br. Life cycle Erect winter annual, gradually becoming smaller toward the top. Yellow rocket seedling. Yellow rocket flowers. Back in cross-section. Reproduction Seeds. Yellow rocket lower leaf. Yellow rocket rosette. Yellow rocket

208

Solar Thermal Rocket Propulsion  

NASA Technical Reports Server (NTRS)

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

Sercel, J. C.

1986-01-01

209

Rocket and Space Technology  

NSDL National Science Digital Library

This site, created by author Robert Braeuning, features material on orbital mechanics, propulsion, rocket hardware, space centers and missions. It includes definitions of important terms and black-and-white diagrams. The page also features information on rocket propellants, rocket propulsion, orbital mechanics, spacecraft systems, vehicle specifications, launch vehicles, manned space flights, planetary spacecraft, and lunar spacecraft. A glossary and discussion forum are also provided. This is a nice resource for a overview of all things involving rockets or other space technologies.

Braeuning, Robert

2009-05-04

210

Building Bottle Rockets  

NSDL National Science Digital Library

You will be investigating the physics behind the launching of a bottle rocket that you will design and build. Go to Air resistance definition and answer the following questions: 1. What is air resistance? 2. How will you design your rocket to reduce the effect of the air resistance? Go to Aerodynamic Forces and list the 4 forces that act on a rocket in motion. Which ones propel the rocket upward and which ...

Benenati, Mr.

2008-03-23

211

Altitude compensating ablative stiffening band for rocket motor nozzles  

NASA Astrophysics Data System (ADS)

A rocket motor nozzle with an ablative internal structural member which provides rigidity to the rocket motor both prior to and during motor operation is described. The rocket motor nozzle of the present invention includes an outer shell which converges to a throat and then diverges to form an entrance. An ablative stiffening band, secured to the inner surface of the nozzle and extending from the aft portion of the nozzle to the throat, is contoured to allow the exhaust gases to flow smoothly through a central longitudinal opening in the band.

Brown, J. Lynn; McIntire, Vaughn W., Jr.; Clontz, Leslie A.; Peckham, Richard J.; Dixon, Alan B. C.; West, James C.

1993-03-01

212

Pop Rocket Variables  

NSDL National Science Digital Library

This is a lesson about the concept of variables in relation to launching pop rockets. Learners will work in teams to test a single variable involved in launching a rocket and learn the variables involved with constructing and launching a water rocket. This is activity 1 of 7 in Dynamic Design: Launch and Propulsion.

213

Magnetic Detachment and Plume Control in Escaping Magnetized Plasma  

SciTech Connect

The model of two-fluid, axisymmetric, ambipolar magnetized plasma detachment from thruster guide fields is extended to include plasmas with non-zero injection angular velocity profiles. Certain plasma injection angular velocity profiles are shown to narrow the plasma plume, thereby increasing exhaust efficiency. As an example, we consider a magnetic guide field arising from a simple current ring and demonstrate plasma injection schemes that more than double the fraction of useful exhaust aperture area, more than halve the exhaust plume angle, and enhance magnetized plasma detachment.

P. F. Schmit and N. J. Fisch

2008-11-05

214

Plume Busters  

NSDL National Science Digital Library

This is an interactive simulator in which students take on the role of an environmental consultant to solve a contamination problem (genrally in the Buffalo River valley alluvial aquifer). Students apply ground-water principles to solve a simulated contamination problem. They calculate the average ground-water velocity from the aquifer porosity and the specific discharge, which in turn is calculated from the aquifer hydraulic conductivity and the hydraulic gradient using Darcy's law. The distances traveled away from the spill site by the edges of the plume are calculated from the average ground-water velocity and time since contaminants first and last entered the aquifer. Students use either production wells or a production/injection well couplet placed appropriately with respect to the moving plume. They design the wellfield and need only a qualitative understanding of well hydraulics including the fundamental concepts of cone of depression, cone of impression, capture zone, and zone of influence. Grade 11-12, undergraduate non-hydrogeology major, and undergraduate hydrogeology major versions of the software are currently available.

Macfarlane, P.; Bohling, Geoffrey

215

Pulse Detonation Rocket MHD Power Experiment  

NASA Technical Reports Server (NTRS)

A pulse detonation research engine (MSFC (Marshall Space Flight Center) Model PDRE (Pulse Detonation Rocket Engine) G-2) has been developed for the purpose of examining integrated propulsion and magnetohydrodynamic power generation applications. The engine is based on a rectangular cross-section tube coupled to a converging-diverging nozzle, which is in turn attached to a segmented Faraday channel. As part of the shakedown testing activity, the pressure wave was interrogated along the length of the engine while running on hydrogen/oxygen propellants. Rapid transition to detonation wave propagation was insured through the use of a short Schelkin spiral near the head of the engine. The measured detonation wave velocities were in excess of 2500 m/s in agreement with the theoretical C-J velocity. The engine was first tested in a straight tube configuration without a nozzle, and the time resolved thrust was measured simultaneously with the head-end pressure. Similar measurements were made with the converging-diverging nozzle attached. The time correlation of the thrust and head-end pressure data was found to be excellent. The major purpose of the converging-diverging nozzle was to configure the engine for driving an MHD generator for the direct production of electrical power. Additional tests were therefore necessary in which seed (cesium-hydroxide dissolved in methanol) was directly injected into the engine as a spray. The exhaust plume was then interrogated with a microwave interferometer in an attempt to characterize the plasma conditions, and emission spectroscopy measurements were also acquired. Data reduction efforts indicate that the plasma exhaust is very highly ionized, although there is some uncertainty at this time as to the relative abundance of negative OH ions. The emission spectroscopy data provided some indication of the species in the exhaust as well as a measurement of temperature. A 24-electrode-pair segmented Faraday channel and 0.6 Tesla permanent magnet assembly were then installed on Marshall Space Flight Center's (MSFC's) rectangular channel pulse detonation research engine. Magnetohydrodynamic (MHD) electrical power extraction experiments were carried out for a range of load impedances in which cesium hydroxide seed (dissolved in methanol) was sprayed into the gaseous oxygen/hydrogen propellants. Positive power extraction was obtained, but preliminary analysis of the data indicated that the plasma electrical conductivity is lower than anticipated and the near-electrode voltage drop is not negligible. It is believed that the electrical conductivity is reduced due to a large population of negative OH ions. This occurs because OH has a strong affinity for capturing free electrons. The effect of near-electrode voltage drop is associated with the high surface-to-volume ratio of the channel (1-inch by 1-inch cross-section) where surface effects play a dominant role. As usual for MHD devices, higher performance will require larger scale devices. Overall, the gathered data is extremely valuable from the standpoint of understanding plasma behavior and for developing empirical scaling laws.

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

2002-01-01

216

Rockets for spin recovery  

NASA Technical Reports Server (NTRS)

The potential effectiveness of rockets as an auxiliary means for an aircraft to effect recovery from spins was investigated. The advances in rocket technology produced by the space effort suggested that currently available systems might obviate many of the problems encountered in earlier rocket systems. A modern fighter configuration known to exhibit a flat spin mode was selected. An analytical study was made of the thrust requirements for a rocket spin recovery system for the subject configuration. These results were then applied to a preliminary systems study of rocket components appropriate to the problem. Subsequent spin tunnel tests were run to evaluate the analytical results.

Whipple, R. D.

1980-01-01

217

Sounding rockets in Antarctica  

NASA Technical Reports Server (NTRS)

Sounding rockets are versatile tools for scientists studying the atmospheric region which is located above balloon altitudes but below orbital satellite altitudes. Three NASA Nike-Tomahawk sounding rockets were launched from Siple Station in Antarctica in an upper atmosphere physics experiment in the austral summer of 1980-81. The 110 kg payloads were carried to 200 km apogee altitudes in a coordinated project with Arcas rocket payloads and instrumented balloons. This Siple Station Expedition demonstrated the feasibility of launching large, near 1,000 kg, rocket systems from research stations in Antarctica. The remoteness of research stations in Antarctica and the severe environment are major considerations in planning rocket launching expeditions.

Alford, G. C.; Cooper, G. W.; Peterson, N. E.

1982-01-01

218

Monitoring thermal impact resulting from solid rocket motor test operations  

NASA Technical Reports Server (NTRS)

The use of remote sensing is discussed with respect to determining the thermal conditions and the immediate environmental effects of large-scale rocket propulsion tests. Data acquired during a test firing of a solid rocket motor are presented including thermal data and surface temperatures from before, during, and after the firing. Thermal impact directly behind the nozzle is assessed, temperature values within the plume are determined, and data are generated for use in an environmental monitoring system which can analyze and forecast impact. The airborne multispectral scanner and thermocouples behind the solid rocket motor discerned that radiant temperatures are higher than predictions indicate and that the testing affects 34 acres of ground. The results are of use in determining the design and area required for developing testing facilities for rocket motors.

Davis, Bruce A.; Thurman, Charles; Carr, Hugh V.

1990-01-01

219

ASSESSMENT OF PLUME DIVING  

EPA Science Inventory

This presentation presents an assessment of plume diving. Observations included: vertical plume delineation at East Patchogue, NY showed BTEX and MTBE plumes sinking on either side of a gravel pit; Lake Druid TCE plume sank beneath unlined drainage ditch; and aquifer recharge/dis...

220

NIMROD Code Simulation of Plasma Exhaust Expansion in the VASIMR Magnetic Nozzle  

Microsoft Academic Search

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR, [1]) engine is an advanced propulsion concept that uses radio frequency waves to accelerate a propellant (typically a Hydrogen or Helium plasma) at much higher speeds than can be reached by any conventional chemical rocket. The high exhaust speed results in a very efficient spacecraft design, as much less propellant mass is required

Alfonso G. Tarditi

2001-01-01

221

ASTRID rocket flight test  

Microsoft Academic Search

On February 4, 1994, we successfully flight tested the ASTRID rocket from Vandenberg Air Force Base. The technology for this rocket originated in the Brilliant Pebbles program and represents a five-year development effort. This rocket demonstrated how our new pumped-propulsion technology-which reduced the total effective engine mass by more than one half and cut the tank mass to one fifth

J. C. Whitehead; L. C. Pittenger; N. J. Colella

1994-01-01

222

Rocket Wind Tunnel  

NSDL National Science Digital Library

In this activity, learners evaluate the potential performance of air rockets placed inside a wind tunnel. Learners measure the rocket's resistance to the flow of air in the tunnel and use the data to construct better rockets. The wind tunnel is prepared by the educator before the activity, but can be built by learners with adult supervision. This lesson plan includes instructions on how to build and use a wind tunnel, extensions, and sample data sheets.

Nasa

2012-05-15

223

Mississippi Plumes  

NSDL National Science Digital Library

The MODIS satellite image above, taken on March 5, shows sediment plumes moving into the Gulf of Mexico from the main branch of the Mississippi River as well as through the bayous in its delta region. It's easy to understand how our nation's longest river is often referred to as 'The Big Muddy'. From the end of the last ice age until the mid 1900's, the Mississippi River created more area each year, but the river has been confined in it's levees since a major flood in 1927. The benefits of controlling the Mississippi River extend throughout the watershed because such control reduces the cost of exporting grain from the midwest and importing petroleum from around the world. Such benefits have come at a tremendous ecological cost that are concentrated in coastal Louisiana. Wetland loss there averaged an acre every 20 minutes throughout the 1950's, 1960's and 1970's. The most recent estimates are about an acre every 40 minutes. Before the mid 1900's, natural wetland loss processes were slower than natural wetland building processes, but human activities have accelerated wetland loss processes and virtually eliminated wetland creation processes.

Center, Nasa G.; Day, Earth S.

224

Soda Straw Rockets  

NSDL National Science Digital Library

This activity is about rocket shape and performance. Learners will test a rocket model and predict its motion. They will launch their rocket multiple times, make observations and record the distance it traveled. They will have the opportunity to answer a research question by collecting and analyzing data related to finding out the best nose cone length and predicting the motion of their model rockets. The lesson models the engineering design process using the 5E instructional model and includes teacher notes, vocabulary, student journal and reading.

225

The effects of solid rocket motor effluents on selected surfaces and solid particle size, distribution, and composition for simulated shuttle booster separation motors  

NASA Technical Reports Server (NTRS)

A series of three tests was conducted using solid rocket propellants to determine the effects a solid rocket plume would have on thermal protective surfaces (TPS). The surfaces tested were those which are baselined for the shuttle vehicle. The propellants used were to simulate the separation solid rocket motors (SSRM) that separate the solid rocket boosters (SRB) from the shuttle launch vehicle. Data cover: (1) the optical effects of the plume environment on spacecraft related surfaces, and (2) the solid particle size, distribution, and composition at TPS sample locations.

Jex, D. W.; Linton, R. C.; Russell, W. M.; Trenkle, J. J.; Wilkes, D. R.

1976-01-01

226

Rockets -- Part II.  

ERIC Educational Resources Information Center

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

Leitner, Alfred

1982-01-01

227

Rockets in Astronomy  

NASA Astrophysics Data System (ADS)

Prior to the launch of the first satellite, Sputnik 1, in 1957, the only way to place scientific instruments above the atmosphere was to launch them on board rockets. In the following four decades, suborbital sounding rockets have continued to play a major role in improving our knowledge of the Earth, geospace and the universe as a whole....

Bond, P.; Murdin, P.

2000-11-01

228

The Rocket Project.  

ERIC Educational Resources Information Center

Describes an extra credit science project in which students compete to see who can build the most efficient water rocket out of a two-liter pop bottle. Provides instructions on how to build a demonstration rocket and launching pad. (MDH)

Winemiller, Jake; And Others

1991-01-01

229

An Ejector Air Intake Design Method for a Novel Rocket-Based Combined-Cycle Rocket Nozzle  

NASA Astrophysics Data System (ADS)

Rocket-based combined-cycle (RBCC) vehicles have the potential to reduce launch costs through the use of several different air breathing engine cycles, which reduce fuel consumption. The rocket-ejector cycle, in which air is entrained into an ejector section by the rocket exhaust, is used at flight speeds below Mach 2. This thesis develops a design method for an air intake geometry around a novel RBCC rocket nozzle design for the rocket-ejector engine cycle. This design method consists of a geometry creation step in which a three-dimensional intake geometry is generated, and a simple flow analysis step which predicts the air intake mass flow rate. The air intake geometry is created using the rocket nozzle geometry and eight primary input parameters. The input parameters are selected to give the user significant control over the air intake shape. The flow analysis step uses an inviscid panel method and an integral boundary layer method to estimate the air mass flow rate through the intake geometry. Intake mass flow rate is used as a performance metric since it directly affects the amount of thrust a rocket-ejector can produce. The design method results for the air intake operating at several different points along the subsonic portion of the Ariane 4 flight profile are found to under predict mass flow rate by up to 8.6% when compared to three-dimensional computational fluid dynamics simulations for the same air intake.

Waung, Timothy S.

230

Imaging Fourier transform spectrometry of chemical plumes  

NASA Astrophysics Data System (ADS)

A midwave infrared (MWIR) imaging Fourier transform spectrometer (FTS), the Telops FIRST-MWE (Field-portable Imaging Radiometric Spectrometer Technology - Midwave Extended) has been utilized for the standoff detection and characterization of chemical plumes. Successful collection and analysis of MWIR hyperspectral imagery of jet engine exhaust has allowed us to produce spatial profiles of both temperature and chemical constituent concentrations of exhaust plumes. Successful characterization of this high temperature combustion event has led to the collection and analysis of hyperspectral imagery of lower temperature emissions from industrial smokestacks. This paper presents MWIR data from remote collection of hyperspectral imagery of methyl salicilate (MeS), a chemical warfare agent simulant, during the Chemical Biological Distributed Early Warning System (CBDEWS) test at Dugway Proving Grounds, UT in 2008. The data did not contain spectral lines associated with emission of MeS. However, a few broad spectral features were present in the background-subtracted plume spectra. Further analysis will be required to assign these features, and determine the utility of MWIR hyperspectral imagery for analysis of chemical warfare agent plumes.

Bradley, Kenneth C.; Gross, Kevin C.; Perram, Glen P.

2009-05-01

231

Solid Rocket Motor Backflow Analysis For CONTOUR Mishap Investigation  

NASA Astrophysics Data System (ADS)

A procedure developed for free molecule modeling of plume backflow from a STAR™ 30BP solid rocket motor is presented for work performed in support of the Comet Nucleus Tour spacecraft mishap investigation. Good general agreement is established with DSMC flowfield results, with interesting deviations developing as the plume backflow approaches the spacecraft surfaces closely, providing insights regarding characteristics of the surface Knudsen layer. Also, investigation of related free expansion results indicate significant discrepancies exist between the rarefied techniques and the continuum results from which their starting surfaces were created. The nature of these differences suggests that convective fluxes to CONTOUR may have been much higher than the rarefied analyses indicated.

Woronowicz, Michael

2005-05-01

232

Modification of the Simons model for calculation of nonradial expansion plumes  

NASA Technical Reports Server (NTRS)

The Simons model is a simple model for calculating the expansion plumes of rockets and thrusters and is a widely used engineering tool for the determination of spacecraft impingement effects. The model assumes that the density of the plume decreases radially from the nozzle exit. Although a high degree of success has been achieved in modeling plumes with moderate Mach numbers, the accuracy obtained under certain conditions is unsatisfactory. A modification made to the model that allows effective description of nonradial behavior in plumes is presented, and the conditions under which its use is preferred are prescribed.

Boyd, I. D.; Stark, J. P. W.

1989-01-01

233

HYDROGEN-OXYGEN ROCKETS  

NSDL National Science Digital Library

During this activity students build a plastic pipette rocket. The first concept will to learn how igniting varying mixtures of hydrogen and oxygen will affect how far the rocket will fly. Second students will observe and manipulate variables to better understand the fundamental chemistry concepts: principles of combustion reactions, kinetics, stoichiometry, gas mixtures, rocketry, and different types of chemical reactions. Finally, students will assess their own understanding of these chemistry concepts by investigating how NASA scientists launch real rockets into space. One follow-up activity would be to investigate and collect data on a launching a heavier object at the school football field.

Reierson, David

234

Andoya Rocket Range  

NSDL National Science Digital Library

The National Aeronautic and Space Administration (NASA) has sponsored the Cleft Accelerated Plasma Experimental Rocket, CAPER, campaign. The objective of this mission is to "probe a fountain of ions that is always blowing into space." Scientists have launched this project just after a solar storm tore apart a part of the Earth's upper atmosphere. The CAPER Rocket launch will take place at the Andoya Rocket Range in January, 1999. This Website offers more information about the CAPER project as well as the launch site.

235

Exhaust gas recirculation system  

Microsoft Academic Search

In an exhaust gas recirculation system for internal combustion engines having a detachable gasket member between an exhaust gas recirculation valve and an exhaust pipe of the engine, the exhaust gas recirculation rate is controlled by a flow control orifice formed in the detachable gasket member. The recirculation valve can be applied to various types of engines requiring various recirculation

K. Numata; Y. Muramatu

1977-01-01

236

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for use in an internal combustion engine includes an intake manifold having a riser portion serving as a heating source for an intake mixture charge, an exhaust gas recirculation passage running from an exhaust manifold to an intake system for introducing part of exhaust gases from the former to the latter, and a temperature-responsive valve

N. Kawai; H. Yamamoto

1980-01-01

237

Exhaust gas recirculator  

Microsoft Academic Search

An exhaust gas recirculator for an internal combustion engine having an exhaust pipe, an intake manifold and a carburetor throttle valve. The exhaust gas recirculator comprises an egr passage which makes the exhaust pipe communicate with the intake manifold, an egr controlling valve and an egr valve respectively arranged in the upper and lower portions of the egr passage. The

Suda

1983-01-01

238

Characterization of rocket propellant combustion products  

SciTech Connect

The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.

1991-12-09

239

Environment effects from SRB exhaust effluents: Technique development and preliminary assessment  

NASA Technical Reports Server (NTRS)

Techniques to determine the environmental effects from the space shuttle SRB (Solid Rocket Booster) exhaust effluents are used to perform a preliminary climatological assessment. The exhaust effluent chemistry study was performed and the exhaust effluent species were determined. A reasonable exhaust particle size distribution is constructed for use in nozzle analyses and for the deposition model. The preliminary assessment is used to identify problems that are associated with the full-scale assessment; therefore, these preliminary air quality results are used with caution in drawing conclusion regarding the environmental effects of the space shuttle exhaust effluents.

Goldford, A. I.; Adelfang, S. I.; Hickey, J. S.; Smith, S. R.; Welty, R. P.; White, G. L.

1977-01-01

240

Investigation of solid plume simulation criteria to produce flight plume effects on multibody configuration in wind tunnel tests  

NASA Technical Reports Server (NTRS)

An investigation to determine the sensitivity of the space shuttle base and forebody aerodynamics to the size and shape of various solid plume simulators was conducted. Families of cones of varying angle and base diameter, at various axial positions behind a Space Shuttle launch vehicle model, were wind tunnel tested. This parametric evaluation yielded base pressure and force coefficient data which indicated that solid plume simulators are an inexpensive, quick method of approximating the effect of engine exhaust plumes on the base and forebody aerodynamics of future, complex multibody launch vehicles.

Frost, A. L.; Dill, C. C.

1986-01-01

241

Antares Rocket Lifts Off!  

NASA Video Gallery

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

242

Microgravity Rockets in Sweden.  

National Technical Information Service (NTIS)

The Swedish sounding rocket program MASER provides periodically recurring flight opportunities for experiments (approximately 1 per year) under microgravity conditions. Conducted by the Swedish Space Corporation MASER is a cost effective, easily accessibl...

L. Bjorn, J. Zaar

1988-01-01

243

NASA: Rocket Activities  

NSDL National Science Digital Library

There are many things in this world that are described as not being as difficult as rocket science. Then, of course, there is the actual science behind rockets. Understandably, this can be difficult for budding space scientists to grasp. Fortunately, NASA has created these fun and interactive activities which relate both to the science and math of rocketry. These particular activities are taken from the "Rocket Educators Guide", and they include activities related to altitude tracking, the world of pinwheels, balloon staging, and of course the construction of an actual paper rocket. Each activity comes complete with instructions, diagrams, and information on the necessary materials. Taken as a whole, these activities could be equally fun whether outside on a brisk fall day as in a classroom setting.

2008-03-21

244

Robust Rocket Engine Concept  

NASA Technical Reports Server (NTRS)

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

Lorenzo, Carl F.

1995-01-01

245

Rocket Launch Probability  

NSDL National Science Digital Library

This applet is designed to teach an application of probability. This Java applet works by simulating a situation where a three stage rocket is about to be launched. In order for a successful launch to occur all three stages of the rocket must successfully pass their pre-takeoff tests. By default, each stage has a 50% chance of success, however, this can be altered by dragging the bar next to each stage.

Exner, Nicholas

2009-01-13

246

Action-Reaction Rocket!  

NSDL National Science Digital Library

Learners construct a rocket from a balloon propelled along a guide string. They use this model to learn about Newton's three laws of motion, examining the effect of different forces on the motion of the rocket. This activity can be combined with other activities to create a larger lesson. Resource contains vocabulary definitions and suggestions for assessment, extensions, and scaling for different levels of learners.

Duren, Sabre; Heavner, Ben; Zarske, Malinda S.; Carlson, Denise

2004-01-01

247

Rocket ballistics and navigation  

NASA Astrophysics Data System (ADS)

The ballistics and navigation of rockets are presented with emphasis on methods for navigating unmanned flight vehicles and possible implementations of these methods. Topics discussed include the flight conditions determined by the geophysical fields and earth atmosphere, mathematical principles of the ballistic support of controlled flight, and principal methods for the guidance of rockets toward moving targets. Attention is also given to the inertial navigation and guidance of ballistic missiles.

Dmitrievskii, A. A.; Ivanov, N. M.; Lysenko, L. N.; Bogodistov, S. S.

248

Monitoring Engine Vibrations And Spectrum Of Exhaust  

NASA Technical Reports Server (NTRS)

Real-time computation of intensities of peaks in visible-light emission spectrum of exhaust combined with real-time spectrum analysis of vibrations into developmental monitoring technique providing up-to-the-second information on conditions of critical bearings in engine. Conceived to monitor conditions of bearings in turbopump suppling oxygen to Space Shuttle main engine, based on observations that both vibrations in bearings and intensities of visible light emitted at specific wavelengths by exhaust plume of engine indicate wear and incipient failure of bearings. Applicable to monitoring "health" of other machinery via spectra of vibrations and electromagnetic emissions from exhausts. Concept related to one described in "Monitoring Bearing Vibrations For Signs Of Damage", (MFS-29734).

Martinez, Carol L.; Randall, Michael R.; Reinert, John W.

1991-01-01

249

Investigation of the Rocket Induced Flow Field in a Rectangular Duct  

NASA Technical Reports Server (NTRS)

Rocket-Based Combined Cycle (RBCC) concepts attempt to improve the performance of launch vehicles at all points in the launch trajectory and make highly reusable launch vehicles a reality. The Aerojet Strutjet RBCC concept consists of a variable geometry duct with internal, vertical struts that functions in ducted rocket, ramjet, scramjet, and pure rocket modes. These struts have rocket and turbine exhaust nozzles imbedded within them. The rocket flows create an ejector effect with the ingested air at subsonic flight velocities. In ramjet and scramjet modes, the fuel rich nozzle flows react with the ingested air producing an afterburner effect. Under a NASA Marshall Space Flight Center contract, the UAH Propulsion Research Center (PRC) has designed and built a Strutjet simulation facility. A scale model of a single strut has been built and is undergoing cold-flow testing to investigate the mixing of the rocket and turbine exhausts with the ingested air. A complementary experimental program is also underway to examine the induced flow-field generated by rocket nozzles confined in a rectangular duct. Characterizing the induced flow behavior is critical to understanding and optimizing the performance of future Strutjet-based RBCC propulsion systems. The proposed paper will present results from the rocket induced flow investigation.

Landrum, D. Brian; Lambert, James; Thames, Mignon; Hawk, Clark

1999-01-01

250

Highlights of Transient Plume Impingement Model Validation and Applications  

NASA Technical Reports Server (NTRS)

This paper describes highlights of an ongoing validation effort conducted to assess the viability of applying a set of analytic point source transient free molecule equations to model behavior ranging from molecular effusion to rocket plumes. The validation effort includes encouraging comparisons to both steady and transient studies involving experimental data and direct simulation Monte Carlo results. Finally, this model is applied to describe features of two exotic transient scenarios involving NASA Goddard Space Flight Center satellite programs.

Woronowicz, Michael

2011-01-01

251

Balloon Rockets in 1D  

NSDL National Science Digital Library

In this structured inquiry activity students will work in groups/ teams to build a balloon rocket of their own design. The rocket will race in one dimension and require that they apply their knowledge of position, time, and velocity.

252

The Optimal Bottle Rocket Lauch  

NSDL National Science Digital Library

This is a computer and outdoor lab based activity in which students design two bottle rockets that are designed to reach maximum height. Students will calculate maximum height and terminal velocity for each rocket launched.

Menzies, Margaret

253

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system is disclosed which includes a control valve inserted in an exhaust gas recirculation passageway for controlling the flow rate of the exhaust gases to be recirculated, a constant pressure chamber defined in the recirculation passageway upstream of the control valve, and a modulator valve with a diaphragm chamber in communication with the constant pressure chamber

S. Nakamura; H. Nohira; H. Tokuda

1980-01-01

254

Exhaust gas recirculation system  

Microsoft Academic Search

In a diesel engine having a turbocharger for feeding supercharged air to the engine, an exhaust gas recirculation passage communicates between the exhaust passage from the engine and the intake passage to a compressor of the turbocharger. A first control valve closes the exhaust gas recirculation passage when the output pressure of the air leading from the compressor is lower

Yoshiba

1982-01-01

255

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for an internal combustion engine. The internal combustion engine including at least one combustion chamber; an intake mechanism for delivering a combustible fluid mixture to the combustion chamber; an ignition system for igniting the combustible mixture; and an exhaust system for carrying exhaust fluid produced by the combustion of the combustible fluid mixture away from

Freesh

1982-01-01

256

Exhaust gas recirculation system  

Microsoft Academic Search

In response to the operation of an internal combustion engine, an input signal pressure is developed which differs from the atmospheric pressure by more than a first predetermined amount when exhaust gas recirculation is desirable. A recirculation valve then opens to permit the recirculation of exhaust gases from the exhaust passage to the intake passage of the engine. In response

1975-01-01

257

Volcanic Plume Measurements with UAV (Invited)  

NASA Astrophysics Data System (ADS)

Volatiles in magmas are the driving force of volcanic eruptions and quantification of volcanic gas flux and composition is important for the volcano monitoring. Recently we developed a portable gas sensor system (Multi-GAS) to quantify the volcanic gas composition by measuring volcanic plumes and obtained volcanic gas compositions of actively degassing volcanoes. As the Multi-GAS measures variation of volcanic gas component concentrations in the pumped air (volcanic plume), we need to bring the apparatus into the volcanic plume. Commonly the observer brings the apparatus to the summit crater by himself but such measurements are not possible under conditions of high risk of volcanic eruption or difficulty to approach the summit due to topography etc. In order to overcome these difficulties, volcanic plume measurements were performed by using manned and unmanned aerial vehicles. The volcanic plume measurements by manned aerial vehicles, however, are also not possible under high risk of eruption. The strict regulation against the modification of the aircraft, such as installing sampling pipes, also causes difficulty due to the high cost. Application of the UAVs for the volcanic plume measurements has a big advantage to avoid these problems. The Multi-GAS consists of IR-CO2 and H2O gas analyzer, SO2-H2O chemical sensors and H2 semiconductor sensor and the total weight ranges 3-6 kg including batteries. The necessary conditions of the UAV for the volcanic plumes measurements with the Multi-GAS are the payloads larger than 3 kg, maximum altitude larger than the plume height and installation of the sampling pipe without contamination of the exhaust gases, as the exhaust gases contain high concentrations of H2, SO2 and CO2. Up to now, three different types of UAVs were applied for the measurements; Kite-plane (Sky Remote) at Miyakejima operated by JMA, Unmanned airplane (Air Photo Service) at Shinomoedake, Kirishima volcano, and Unmanned helicopter (Yamaha) at Sakurajima volcano operated by ERI, Tokyo University. In all cases, we could estimated volcanic gas compositions, such as CO2/SO2 ratios, but also found out that it is necessary to improve the techniques to avoid the contamination of the exhaust gases and to approach more concentrated part of the plume. It was also revealed that the aerial measurements have an advantage of the stable background. The error of the volcanic gas composition estimates are largely due to the large fluctuation of the atmospheric H2O and CO2 concentrations near the ground. The stable atmospheric background obtained by the UAV measurements enables accurate estimate of the volcanic gas compositions. One of the most successful measurements was that on May 18, 2011 at Shinomoedake, Kirishima volcano during repeating Vulcanian eruption stage. The major component composition was obtained as H2O=97, CO2=1.5, SO2=0.2, H2S=0.24, H2=0.006 mol%; the high CO2 contents suggests relatively deep source of the magma degassing and the apparent equilibrium temperature obtained as 400°C indicates that the gas was cooled during ascent to the surface. The volcanic plume measurement with UAV will become an important tool for the volcano monitoring that provides important information to understand eruption processes.

Shinohara, H.; Kaneko, T.; Ohminato, T.

2013-12-01

258

Rocket Me into Space  

NSDL National Science Digital Library

One of the exciting challenges for engineers is the idea of exploration. This lesson looks more closely at Spaceman Rohan, Spacewoman Tess, their daughter Maya, and their challenges with getting to space, setting up satellites, and exploring uncharted waters via a canoe. This lesson reinforces rockets as a vehicle that helps us explore outside the Earth's atmosphere (that is, to move without air) by using the principles of Newton's third law of motion. Also, the ideas of thrust, control and weight â all principles that engineers deal with when building a rocket â are introduced.

Integrated Teaching And Learning Program

259

Advanced liquid rockets  

NASA Technical Reports Server (NTRS)

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

Schneider, Steven J.

1992-01-01

260

Laser Rayleigh and Raman Diagnostics for Small Hydrogen/oxygen Rockets  

NASA Technical Reports Server (NTRS)

Localized velocity, temperature, and species concentration measurements in rocket flow fields are needed to evaluate predictive computational fluid dynamics (CFD) codes and identify causes of poor rocket performance. Velocity, temperature, and total number density information have been successfully extracted from spectrally resolved Rayleigh scattering in the plume of small hydrogen/oxygen rockets. Light from a narrow band laser is scattered from the moving molecules with a Doppler shifted frequency. Two components of the velocity can be extracted by observing the scattered light from two directions. Thermal broadening of the scattered light provides a measure of the temperature, while the integrated scattering intensity is proportional to the number density. Spontaneous Raman scattering has been used to measure temperature and species concentration in similar plumes. Light from a dye laser is scattered by molecules in the rocket plume. Raman spectra scattered from major species are resolved by observing the inelastically scattered light with linear array mounted to a spectrometer. Temperature and oxygen concentrations have been extracted by fitting a model function to the measured Raman spectrum. Results of measurements on small rockets mounted inside a high altitude chamber using both diagnostic techniques are reported.

Degroot, Wilhelmus A.; Zupanc, Frank J.

1993-01-01

261

Overview of rocket engine control  

NASA Technical Reports Server (NTRS)

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

Lorenzo, Carl F.; Musgrave, Jeffrey L.

1991-01-01

262

Jet exhaust noise suppressor  

NASA Technical Reports Server (NTRS)

Noise suppression for a jet engine exhaust is provided by an annular divergent body attached to an exhaust nozzle. The smallest diameter of the divergent body is larger than the diameter of the exhaust nozzle exit to form an annular step which produces a shock wave in the exhaust as it passes the step. An annular shroud is disposed around the divergent body and causes outside air to pass through voids in the divergent body to mix with the jet exhaust gas. The divergent body includes a plurality of channels with separators between the channels.

Huff, R. G. (inventor)

1974-01-01

263

Lightweight polymeric exhaust components  

US Patent & Trademark Office Database

Disclosed is a muffler assembly including: a) polymeric housing having an interior surface and at least one opening for at least one inlet and one outlet exhaust pipe; b) at least one metal inlet exhaust pipe and at least one metal outlet exhaust pipe positioned within the openings to provide housing-exhaust pipe interfaces; c) a thermal insulating material coating the interior surface of the polymeric housing and extending through the housing-exhaust pipe interfaces; wherein the thermal insulating material seals the muffler assembly at the housing-exhaust pipe interfaces; and wherein the muffler assembly has a leak rate of 105 Liters/minute or less at 4.5 psig pressure. An optional muffler assembly has body mounting adapters attached to the inlet and outlet exhaust pipes and positioned within the openings to provide housing-body mounting adapter interfaces. Also disclosed are processes for manufacturing the muffler assemblies.

2013-08-13

264

Liquid Rocket Engine Testing  

NASA Technical Reports Server (NTRS)

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

Rahman, Shamim

2005-01-01

265

Liquid rocket valve components  

NASA Technical Reports Server (NTRS)

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

1973-01-01

266

Solid Rocket Motor Test  

NASA Technical Reports Server (NTRS)

Shown is a test of the TEM-13 solid rocket motor at the ATK test facility in Utah in support of the Ares/CLV first stage. This image is extracted from high definition video and is the highest resolution available.

2008-01-01

267

Solid Rocket Motor Test  

NASA Technical Reports Server (NTRS)

Shown is a test of the TEM-13 Solid Rocket Motor in support of the Ares/CLV first stage at ATK, Utah . Constellation/Ares project. This image is extracted from a high definition video file and is the highest resolution available.

2008-01-01

268

Solid Rocket Motor Test  

NASA Technical Reports Server (NTRS)

Shown is a test of the TEM-13 Solid Rocket Motor in support of the Ares/CLV first stage at ATK, Utah . Constellaton/Ares project. This image is extracted from a high definition video file and is the highest resolution available.

2008-01-01

269

Dr. Goddard Transports Rocket  

NASA Technical Reports Server (NTRS)

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

1974-01-01

270

Water Rocket Launch  

NSDL National Science Digital Library

In this activity, learners explore rocketry and the principals of space flight. Learners work in teams with adult supervision and construct and launch a rocket from a soda bottle and everyday materials powered by an air pump. Learners observe their own achievements and challenges, as well as those of other teams, complete a reflection sheet, and present their experiences to the class.

Ieee

2014-06-18

271

Rockets using Liquid Oxygen  

NASA Technical Reports Server (NTRS)

It is my task to discuss rocket propulsion using liquid oxygen and my treatment must be highly condensed for the ideas and experiments pertaining to this classic type of rocket are so numerous that one could occupy a whole morning with a detailed presentation. First, with regard to oxygen itself as compared with competing oxygen carriers, it is known that the liquid state of oxygen, in spite of the low boiling point, is more advantageous than the gaseous form of oxygen in pressure tanks, therefore only liquid oxygen need be compared with the oxygen carriers. The advantages of liquid oxygen are absolute purity and unlimited availability at relatively small cost in energy. The disadvantages are those arising from the impossibility of absolute isolation from heat; consequently, allowance must always be made for a certain degree of vaporization and only vented vessels can be used for storage and transportation. This necessity alone eliminates many fields of application, for example, at the front lines. In addition, liquid oxygen has a lower specific weight than other oxygen carriers, therefore many accessories become relatively larger and heavier in the case of an oxygen rocket, for example, the supply tanks and the pumps. The advantages thus become effective only in those cases where definitely scheduled operation and a large ground organization are possible and when the flight requires a great concentration of energy relative to weight. With the aim of brevity, a diagram of an oxygen rocket will be presented and the problem of various component parts that receive particularly thorough investigation in this classic case but which are also often applicable to other rocket types will be referred to.

Busemann, Adolf

1947-01-01

272

Exhaust system combustor  

SciTech Connect

This patent describes a combustor for use in an exhaust gas system which combustor is tolerant to thermal gradients so as not to degrade the atomization of fuel therein. It comprises an exhaust duct for conveying exhaust gas therethrough having a side wall, an inlet end through which exhaust gas enters the exhaust duct, and an outlet end through which the exhaust gas exits the exhaust duct; a combustion chamber having an atomizer end and a combustion end fixedly mounted in the exhaust duct facing the outlet end of the exhaust duct; an atomizer mounted in the atomizer end of the combustion chamber for spraying atomized fuel into the combustion chamber; an air duct for conveying combustion air to the combustion chamber and extending through the side wall of the exhaust duct to the atomizer end of the combustion chamber; and a fuel conduit fixedly joined to the atomizer for conveying fuel to the atomizer, the fuel conduit having at least a portion thereof extending in the air duct, the portion also including a longitudinal compliance portion for allowing thermal expansion and contraction of the air duct and the combustion chamber relative to the fuel conduit while maintaining a constant position and alignment of the atomizer with respect to the combustion chamber.

Simmons, H.C.; Jones, R.V.

1992-04-21

273

Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors  

NASA Technical Reports Server (NTRS)

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.

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

2014-01-01

274

Rapid Mars transits with exhaust-modulated plasma propulsion  

NASA Technical Reports Server (NTRS)

The operational characteristics of the Exhaust-Modulated Plasma Rocket are described. Four basic human and robotic mission scenarios to Mars are analyzed using numerical optimization techniques at variable specific impulse and constant power. The device is well suited for 'split-sprint' missions, allowing fast, one-way low-payload human transits of 90 to 104 days, as well as slower, 180-day, high-payload robotic precursor flights. Abort capabilities, essential for human missions, are also explored.

Chang-Diaz, Franklin R.; Braden, Ellen; Johnson, Ivan; Hsu, Michael M.; Yang, Tien Fang

1995-01-01

275

Liquid rocket engine injectors  

NASA Technical Reports Server (NTRS)

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

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

1976-01-01

276

Laser rocket system analysis  

NASA Technical Reports Server (NTRS)

The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

Jones, W. S.; Forsyth, J. B.; Skratt, J. P.

1979-01-01

277

Advanced rocket propulsion  

NASA Astrophysics Data System (ADS)

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

Obrien, Charles J.

1993-02-01

278

Microfabricated Liquid Rocket Motors  

NASA Technical Reports Server (NTRS)

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

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

2003-01-01

279

Newton Rocket Car  

NSDL National Science Digital Library

The purpose of this activity is to demonstrate Newton's third law of motion â which states that every action has an equal and opposite reaction â through a small wooden car. The Newton cars show how action/reaction works and how the mass of a moving object affects the acceleration and force of the system. Subsequently, the Newton cars provide students with an excellent analogy for how rockets actually work.

Integrated Teaching And Learning Program

280

Two-Dimensional Motions of Rockets  

ERIC Educational Resources Information Center

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…

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

281

CHLORINATED SOLVENT PLUME CONTROL  

EPA Science Inventory

This lecture will cover recent success in controlling and assessing the treatment of shallow ground water plumes of chlorinated solvents, other halogenated organic compounds, and methyl tert-butyl ether (MTBE)....

282

Sulfur plumes off Namibia  

NASA Technical Reports Server (NTRS)

Sulfur plumes rising up from the bottom of the ocean floor produce colorful swirls in the waters off the coast of Namibia in southern Africa. The plumes come from the breakdown of marine plant matter by anaerobic bacteria that do not need oxygen to live. This image was acquired by the Moderate Resolution Imaging Spectroradiometer (MODIS) on the Terra satellite on April 24, 2002 Credit: Jacques Descloitres, MODIS Land Rapid Response Team, NASA/GSFC

2002-01-01

283

Validation of scramjet exhaust simulation technique at Mach 6  

NASA Technical Reports Server (NTRS)

Current design philosophy for hydrogen-fueled, scramjet-powered hypersonic aircraft results in configurations with strong couplings between the engine plume and vehicle aerodynamics. The experimental verification of the scramjet exhaust simulation is described. The scramjet exhaust was reproduced for the Mach 6 flight condition by the detonation tube simulator. The exhaust flow pressure profiles, and to a large extent the heat transfer rate profiles, were then duplicated by cool gas mixtures of Argon and Freon 13B1 or Freon 12. The results of these experiments indicate that a cool gas simulation of the hot scramjet exhaust is a viable simulation technique except for phenomena which are dependent on the wall temperature relative to flow temperature.

Hopkins, H. B.; Konopka, W.; Leng, J.

1979-01-01

284

Liquid Rocket Engine Testing Overview  

NASA Technical Reports Server (NTRS)

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.

Rahman, Shamim

2005-01-01

285

Exhaust gas recirculation system  

Microsoft Academic Search

An auxiliary flow control valve is disposed in an exhaust gas recirculation passageway at a location upstream of a main flow control valve controlling the flow rate of exhaust gases recirculated into an intake system. Two conduits communicating with an intake manifold communicate with a vacuum chamber of a servo motor controlling the main flow control valve. One of the

Y. Fujikawa; Y. Nakajima; Y. Hayashi; K. Sugihara; Y. Hase

1976-01-01

286

Exhaust gas recirculation system  

Microsoft Academic Search

Apparatus for controlling exhaust gas recirculation in an internal combustion engine employs a first control valve in an exhaust gas recirculation passageway, a second control valve in an air conduit connecting the intake passage to atmosphere through selective restriction means, and a regulating valve responsive to differential vacuum intensities for actuating the control valves. The restriction means comprise a plurality

Y. Itoh; A. Totsune; H. Yamabe

1982-01-01

287

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system comprises a valve assembly for controlling the recirculation of exhaust gases in such a manner as to maintain a difference between a first and second pressure at a predetermined value. The first pressure is a pressure in a zone in an air induction passage between a throttle valve therein and a flow restrictor disposed therein

Aoyama

1980-01-01

288

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for an internal combustion engine is disclosed in which an exhaust gas control valve and a pressure regulating valve is constructed as one body to facilitate the mounting of the system on the engine and to improve the response characteristic of the system. Also, the chamber of the pressure regulating valve opposite to the chamber

S. Hayashi; N. Shibata; Y. Takahara; S. Yamada

1981-01-01

289

EXHAUST GAS RECIRCULATION  

E-print Network

EXHAUST GAS RECIRCULATION (EGR) COOLER TESTING Southwest Research Institute® #12;overnment) and oxides of nitrogen (NOx). The use of exhaust gas recirculation (EGR) coolers is considered research, offers complete facilities for testing diesel engines and their emissions control systems

Chapman, Clark R.

290

Exhaust gas recirculating system  

Microsoft Academic Search

This patent describes an exhaust gas recirculating system for an internal combustion engine having an intake passage, a main throttle valve located in the intake passage, and an exhaust passage. The engine is installed in an automotive vehicle provided with a traction control system including means for sensing slippage of a drive wheel of the vehicle during acceleration, a second

H. Takahashi; T. Naganawa

1988-01-01

291

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for cleaning exhaust gas from an internal combustion engine is provided in which a variable constriction is provided between an intake pipe and a pressure control valve in operative connection to a throttle valve in the carburetor and the pressure differential across said variable constriction is maintained constant to keep off any influence of the

M. Minoura; K. Yorioka

1980-01-01

292

Diesel engine exhaust oxidizer  

SciTech Connect

This patent describes a diesel engine exhaust oxidizing device. It comprises: an enclosure having an inlet for receiving diesel engine exhaust, a main flow path through the enclosure to an outlet of the enclosure, a by-ass through the enclosure, and a microprocessor control means.

Kammel, R.A.

1992-06-16

293

Exhaust gas recirculating system  

Microsoft Academic Search

An internal combustion engine equipped with an exhaust gas recirculating system is described. The exhaust gas recirculating system comprises an EGR valve normally closing an EGR duct to prevent recirculation and movable by a signal vacuum applied thereto to an open position and control means operable to provide a signal vacuum. The control means includes a vacuum line connecting a

Harada

1978-01-01

294

Duplex tab exhaust nozzle  

NASA Technical Reports Server (NTRS)

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

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

2012-01-01

295

Electric exhaust gas recirculation valve  

Microsoft Academic Search

An electrically actuated EGR valve is described for controlling EGR gases in response to electric signals from a computer, the EGR valve comprising: a valve housing having an exhaust gas inlet port for passage of exhaust gases; an exhaust gas outlet port; an exhaust gas passage extending between; poppet valve means for selectively opening and closing exhaust gas passage, the

Akagi

1987-01-01

296

ISS Update: VASIMR Plasma Rocket  

NASA Video Gallery

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

297

Dual-fuel, dual-mode rocket engine  

NASA Technical Reports Server (NTRS)

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

Martin, James A. (inventor)

1989-01-01

298

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

NASA Technical Reports Server (NTRS)

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

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

2011-01-01

299

Particulate exhaust emissions from an experimental combustor. [gas turbine engine  

NASA Technical Reports Server (NTRS)

The concentration of dry particulates (carbon) in the exhaust of an experimental gas turbine combustor was measured at simulated takeoff operating conditions and correlated with the standard smoke-number measurement. Carbon was determined quantitatively from a sample collected on a fiberglass filter by converting the carbon in the smoke sample to carbon dioxide and then measuring the volume of carbon dioxide formed by gas chromatography. At a smoke of 25 (threshold of visibility of the smoke plume for large turbojets) the carbon concentration was 2.8 mg carbon/cu m exhaust gas, which is equivalent to an emission index of 0.17 g carbon/kg fuel.

Norgren, C. T.; Ingebo, R. D.

1975-01-01

300

Catalytic Microtube Rocket Igniter  

NASA Technical Reports Server (NTRS)

Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Schneider, Steven J.; Deans, Matthew C.

2011-01-01

301

Rocket + Science = Dialogue  

NASA Technical Reports Server (NTRS)

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.

Morris,Bruce; Sullivan, Greg; Burkey, Martin

2010-01-01

302

Exhaust-gas recirculation system  

Microsoft Academic Search

An exhaust-gas recirculation (EGR) system for an internal-combustion engine comprises an EGR valve for controlling the amount of exhaust-gas recirculation installed midway in an EGR passage that establishes communication between the intake and exhaust pipes of the engine, and an exhaust-gas transducer valve for opening and closing by the exhaust pressure of the exhaust gas an atmospheric-releasing orifice formed midway

K. Yamada; C. Niida; T. Takayama

1979-01-01

303

An Injector for the Variable Specific Impulse Magnetoplasma Rocket  

NASA Astrophysics Data System (ADS)

We present a summary of progress on the development of a plasma injector for NASA's VASIMR (Variable Specific Impulse Magnetoplasma Rocket) engine. The plasma rocket constrains a flowing plasma in an asymmetric magnetic bottle and exhausts it through a magnetic nozzle to produce thrust. The injector is a plasma source located on the axis of symmetry, forward of the series of coils forming the constraining magnetic field. The injector is intended to produce a well-collimated jet of highly ionized plasma which will enter the central cell of the machine through its forward mirror. The prototype design is based on that of a Lorentz Force Accelerator developed as a thruster by the electric propulsion research group at Princeton. Our investigation focuses on the effects of the rocket's magnetic field on the operation of the injector, the effect of a local magnetic field on the discharge behavior, and the effectiveness of discharge initiation by glow discharge versus initiation by ECRH. We evaluate the performance of this prototype injector by comparing the characteristics of the plasma it inserts into the central cell of the engine with the characteristics called for in the design of the plasma rocket.

Glover, T. W.; Chang-Diaz, F. R.; Squire, J. P.; Chan, A. A.

1997-11-01

304

Coal-Fired Rocket Engine  

NASA Technical Reports Server (NTRS)

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

Anderson, Floyd A.

1987-01-01

305

Otrag rocket experiments in Africa  

NASA Technical Reports Server (NTRS)

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

1978-01-01

306

What fuel for a rocket?  

E-print Network

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

E. N. Miranda

2012-08-13

307

What fuel for a rocket?  

E-print Network

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

Miranda, E N

2012-01-01

308

Electric rockets get a boost  

Microsoft Academic Search

This article reports that xenon-ion thrusters are expected to replace conventional chemical rockets in many nonlaunch propulsion tasks, such as controlling satellite orbits and sending space probes on long exploratory missions. The space age dawned some four decades ago with the arrival of powerful chemical rockets that could propel vehicles fast enough to escape the grasp of earth`s gravity. Today,

Ashley

1995-01-01

309

Relativistic rocket and space flight  

Microsoft Academic Search

We introduce the new concept of “proper speed”. It turns out that relativistic space flight can be sized and described more adequately by this concept. In particular, it is shown that the physics and equations of relativistic and classical space flight have a one-to-one relationship. The rocket equations are even identical. In addition, we generalize the expressions for the rocket

Ulrich Walter

2006-01-01

310

Handheld Water Bottle Rocket & Launcher  

NSDL National Science Digital Library

In this activity, learners build handheld rockets and launchers out of PVC pipes and plastic bottles. Use this activity to demonstrate acceleration, air pressure, and Newton's Laws of Motion. Note: a PVC cutter, side cutters, PVC cement glue and other tools are required to build this project. Safety note: These rockets should only be launched in large, open, outdoor areas.

Workshop, Fresno C.

2012-01-01

311

Soda-Bottle Water Rockets.  

ERIC Educational Resources Information Center

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

Kagan, David; And Others

1995-01-01

312

Density and optical properties of SPARCS plumes  

NASA Technical Reports Server (NTRS)

Propellant gases emitted by attitude control systems such as SPARCS (Solar Pointing Aerobee Rocket Control System) and possible interference with experiments aboard the payloads are discussed. The optical properties of seven actual and potential gases emitted by propellant systems (CF4, N2H4, NH3, N2, CO2, Ar, and He) are presented. A compilation of absorption coefficients from 1 Angstrom to 50 microns and a summary of fluorescent spectra and efficiencies are provided. Since Freon-14 (CF4) is of primary importance to SPARCS, an experimental search for the fluorescent spectrum of CF4 was performed by exciting the gas with 920 Angstrom UV photons. The result was compared with an electron impact induced spectrum of CF4, and conclusions drawn about the nature of the radiating species. A detailed study of the CF4 flow fields and plume densities for typical SPARCS controlled payloads was made using gas dynamic codes which included the effects of vehicle shading and condensation. The importance of the optical properties of CF4 plumes was investigated and it is concluded that absorption is negligible but fluoresence may be significant in some cases.

Brown, W. A.; Kumer, J. B.; Cooper, C. E., Jr.

1972-01-01

313

Contamination control and plume assessment of low-energy thrusters  

NASA Technical Reports Server (NTRS)

Potential contamination of a spacecraft cryogenic surface by a xenon (Xe) ion generator was evaluated. The analysis involves the description of the plume exhausted from the generator with its relative component fluxes on the spacecraft surfaces, and verification of the conditions for condensation, adsorption, and sputtering at those locations. The data describing the plume fluxes and their effects on surfaces were obtained from two sources: the tests carried out with the Xe generator in a small vacuum chamber to indicate deposits and sputter on monitor slides; and the extensive tests with a mercury (Hg) ion thruster in a large vacuum chamber. The Hg thruster tests provided data on the neutrals, on low-energy ion fluxes, on high-energy ion fluxes, and on sputtered materials at several locations within the plume.

Scialdone, John J.

1993-01-01

314

Mars Rocket Propulsion System  

NASA Technical Reports Server (NTRS)

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

Zubrin, Robert; Harber, Dan; Nabors, Sammy

2008-01-01

315

Rhenium Rocket Manufacturing Technology  

NASA Technical Reports Server (NTRS)

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

1997-01-01

316

Near-field vector intensity measurements of a small solid rocket motor.  

PubMed

Near-field vector intensity measurements have been made of a 12.7-cm diameter nozzle solid rocket motor. The measurements utilized a test rig comprised of four probes each with four low-sensitivity 6.35-mm pressure microphones in a tetrahedral arrangement. Measurements were made with the rig at nine positions (36 probe locations) within six nozzle diameters of the plume shear layer. Overall levels at these locations range from 135 to 157 dB re 20 microPa. Vector intensity maps reveal that, as frequency increases, the dominant source region contracts and moves upstream with peak directivity at greater angles from the plume axis. PMID:20707417

Gee, Kent L; Giraud, Jarom H; Blotter, Jonathan D; Sommerfeldt, Scott D

2010-08-01

317

Development of Jovian Impactor Plumes  

Microsoft Academic Search

We have simulated the entry of cometary nuclei into Jupiter (Korycansky et al. 2006 ApJ 646. 642-652) and now extend our 3D hydrodynamic models to plume formation and blowout. We investigate the physics that control plume behavior, including why and where plumes \\

Csaba J. Palotai; D. G. Korycansky; J. Harrington; T. Gabriel

2010-01-01

318

Heat sources for mantle plumes  

Microsoft Academic Search

Melting anomalies in the Earth's upper mantle have often been attributed to the presence of mantle plumes that may originate in the lower mantle, possibly from the core-mantle boundary. Globally, mantle plumes exhibit a large range in buoyancy flux that is proportional to their temperature and volume. Plumes with higher buoyancy fluxes should have higher temperatures and experience higher degrees

C. Beier; T. Rushmer; S. P. Turner

2008-01-01

319

Buoyant plume calculations  

SciTech Connect

Smoke from raging fires produced in the aftermath of a major nuclear exchange has been predicted to cause large decreases in surface temperatures. However, the extent of the decrease and even the sign of the temperature change, depend on how the smoke is distributed with altitude. We present a model capable of evaluating the initial distribution of lofted smoke above a massive fire. Calculations are shown for a two-dimensional slab version of the model and a full three-dimensional version. The model has been evaluated by simulating smoke heights for the Hamburg firestorm of 1943 and a smaller scale oil fire which occurred in Long Beach in 1958. Our plume heights for these fires are compared to those predicted by the classical Morton-Taylor-Turner theory for weakly buoyant plumes. We consider the effect of the added buoyancy caused by condensation of water-laden ground level air being carried to high altitude with the convection column as well as the effects of background wind on the calculated smoke plume heights for several fire intensities. We find that the rise height of the plume depends on the assumed background atmospheric conditions as well as the fire intensity. Little smoke is injected into the stratosphere unless the fire is unusually intense, or atmospheric conditions are more unstable than we have assumed. For intense fires significant amounts of water vapor are condensed raising the possibility of early scavenging of smoke particles by precipitation. 26 references, 11 figures.

Penner, J.E.; Haselman, L.C.; Edwards, L.L.

1985-01-01

320

Evaluation of Visible Plumes.  

ERIC Educational Resources Information Center

Developed for presentation at the 12th Conference on Methods in Air Pollution and Industrial Hygiene Studies, University of Southern California, April, 1971, this outline discusses plumes with contaminants that are visible to the naked eye. Information covers: (1) history of air pollution control regulations, (2) need for methods of evaluating…

Brennan, Thomas

321

Enceladus' Water Vapour Plumes  

NASA Technical Reports Server (NTRS)

A viewgraph presentation on the discovery of Enceladus water vapor plumes is shown. Conservative modeling of this water vapor is also presented and also shows that Enceladus is the source of most of the water required to supply the neutrals in Saturn's system and resupply the E-ring against losses.

Hansen, Candice J.; Esposito, L.; Colwell, J.; Hendrix, A.; Matson, Dennis; Parkinson, C.; Pryor, W.; Shemansky, D.; Stewart, I.; Tew, J.; Yung, Y.

2006-01-01

322

Improving operational plume forecasts  

NASA Astrophysics Data System (ADS)

Forecasting how plumes of particles, such as radioactive particles from a nuclear disaster, will be transported and dispersed in the atmosphere is an important but computationally challenging task. During the Fukushima nuclear disaster in Japan, operational plume forecasts were produced each day, but as the emissions continued, previous emissions were not included in the simulations used for forecasts because it became impractical to rerun the simulations each day from the beginning of the accident. Draxler and Rolph examine whether it is possible to improve plume simulation speed and flexibility as conditions and input data change. The authors use a method known as a transfer coefficient matrix approach that allows them to simulate many radionuclides using only a few generic species for the computation. Their simulations work faster by dividing the computation into separate independent segments in such a way that the most computationally time consuming pieces of the calculation need to be done only once. This makes it possible to provide real-time operational plume forecasts by continuously updating the previous simulations as new data become available. They tested their method using data from the Fukushima incident to show that it performed well. (Journal of Geophysical Research-Atmospheres, doi:10.1029/2011JD017205, 2012)

Balcerak, Ernie

2012-04-01

323

British used Congreve Rockets to Attack Napoleon  

NASA Technical Reports Server (NTRS)

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

2004-01-01

324

Facility for cold flow testing of solid rocket motor models  

NASA Astrophysics Data System (ADS)

A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solid propellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The performance of the self-pumping exhaust diffuser was verified down to exhaust pressures of 1.379 x 10 exp 4 Pa. The facility was successfully operated over the entire range of design pressures and flowrates and is available for national use by industry and government agencies requiring facilities capable of testing SRM cold flow models to support development programs or resolve problems arising on operational flight systems.

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

1992-02-01

325

Rocketing into Adaptive Inquiry  

NSDL National Science Digital Library

To ensure that each student achieves success, teachers can tailor activities with students' strengths and weaknesses in mind using the process of adaptive inquiry. Adaptive inquiry is the product of the synergistic relationship between what a student brings to the classroom and the teacher's ability to shape a lesson in response to the needs of the student. The following is an example of an adaptive inquiry activity that uses Launch System Compressor (LCS) Rockets (paper tubes launched by squeezing a plastic bag filled with air). Many divergent outcomes are possible with this activity, but each one can be used to reach the ultimate objective of this lesson--teaching Newton's third law of motion.

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

2002-01-01

326

Rocket/launcher structural dynamics  

NASA Technical Reports Server (NTRS)

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

Ferragut, N. J.

1976-01-01

327

Computational study of variable area ejector rocket flowfields  

NASA Astrophysics Data System (ADS)

Access to space has always been a scientific priority for countries which can afford the prohibitive costs associated with launch. However, the large scale exploitation of space by the business community will require the cost of placing payloads into orbit be dramatically reduced for space to become a truly profitable commodity. To this end, this work focuses on a next generation propulsive technology called the Rocket Based Combined Cycle (RBCC) engine in which rocket, ejector, ramjet, and scramjet cycles operate within the same engine environment. Using an in house numerical code solving the axisymmetric version of the Favre averaged Navier Stokes equations (including the Wilcox ko turbulence model with dilatational dissipation) a systematic study of various ejector designs within an RBCC engine is undertaken. It is shown that by using a central rocket placed along the axisymmetric axis in combination with an annular rocket placed along the outer wall of the ejector, one can obtain compression ratios of approximately 2.5 for the case where both the entrained air and rocket exhaust mass flows are equal. Further, it is shown that constricting the exit area, and the manner in which this constriction is performed, has a significant positive impact on the compression ratio. For a decrease in area of 25% a purely conical ejector can increase the compression ratio by an additional 23% compared to an equal length unconstricted ejector. The use of a more sharply angled conical section followed by a cylindrical section to maintain equivalent ejector lengths can further increase the compression ratio by 5--7% for a total increase of approximately 30%.

Etele, Jason

328

Atmospheric diffusion predictions for the exhaust effluents from the launch of a Titan 3C, December 13, 1973  

NASA Technical Reports Server (NTRS)

Results for the predictions with the NASA/MSFC Multilayer Diffusion Model for the dispersive transport of the Titan 3C rocket exhaust effluents for the 1857 EST launch on December 13, 1973, from the Eastern Test Range at Cape Canaveral Air Force Station are presented. An atmospheric assessment is made in support of the joint Marshall Space Flight Center, Langley Research Center, and Kennedy Space Center rocket exhaust prediction and measurement program. The predictions are primarily intended to define a monitoring grid and for a postflight assessment of the field measurements in order to improve diffusion prediction techniques.

Stephens, J. B. (editor)

1974-01-01

329

Sandia Laboratories rocket program - A review  

Microsoft Academic Search

A historical review of Sandia Laboratories rocket programs is presented. From the 60 rocket systems developed at Sandia since 1957, 1225 rockets have been launched at 19 sites, worldwide. Typical rockets developed for the nuclear readiness test program are the Terrier-Sandhawk sounding rocket (boosts a 91-kg, 33-cm-diam payload to an altitude of 427 km) and the Strypi II warhead carrier

G. A. Fowler; R. C. Maydew; W. R. Barton

1976-01-01

330

Exhaust gas recirculator  

SciTech Connect

An exhaust gas recirculator for an internal combustion engine having an exhaust pipe, an intake manifold and a carburetor throttle valve. The exhaust gas recirculator comprises an egr passage which makes the exhaust pipe communicate with the intake manifold, an egr controlling valve and an egr valve respectively arranged in the upper and lower portions of the egr passage. The egr valve operates in association with the carburetor throttle valve for metering the flow of egr gas. The egr controlling valve is separated by a diaphragm into an egr gas chamber communicating with the egr passage between the egr controlling valve and the egr valve and a negative pressure chamber communicating with the intake manifold. The negative pressure chamber contains a compression spring, and the diaphragm is connected with a valve member through a rod upon which is disposed a stopper to serve as a different seal in place of the valve member to close off the exhaust gas passage, which valve member and stopper are constructed to be opened and closed by pressure difference between the egr gas chamber and the negative pressure chamber and by elastic force of the compression spring. The egr controlling valve functions to control the pressure difference around the egr valve to be constant.

Suda, K.

1983-01-04

331

Behavior of Mercury Emissions from a Commercial Coal-Fired Utility Boiler: TheRelationship Between Stack Speciation and Near-Field Plume Measurements  

EPA Science Inventory

The reduction of divalent gaseous mercury (HgII) to elemental gaseous mercury (Hg0) in a commercial coal-fired power plant (CFPP)exhaust plume was investigated by simultaneous measurement in-stack and in-plume as part of a collaborative study among the U.S....

332

World Data Center A (rockets and satellites) catalogue of data. Volume 1, part A: Sounding rockets  

NASA Technical Reports Server (NTRS)

A cumulative listing of all scientifically successful rockets that have been identified from various sources is presented. The listing starts with the V-2 rocket launched on 7 March 1947 and contains all rockets identified up to 31 December 1971.

1972-01-01

333

Navigating the Rockets Educator Guide  

NASA Video Gallery

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

334

Small Solid Rocket Motor Test  

NASA Video Gallery

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

335

Easier Analysis With Rocket Science  

NASA Technical Reports Server (NTRS)

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

2003-01-01

336

Impact of aircraft plume dynamics on airport local air quality  

NASA Astrophysics Data System (ADS)

Air quality degradation in the locality of airports poses a public health hazard. The ability to quantitatively predict the air quality impacts of airport operations is of importance for assessing the air quality and public health impacts of airports today, of future developments, and for evaluating approaches for mitigating these impacts. However, studies such as the Project for the Sustainable Development of Heathrow have highlighted shortcomings in understanding of aircraft plume dispersion. Further, if national or international aviation environmental policies are to be assessed, a computationally efficient method of modeling aircraft plume dispersion is needed. To address these needs, we describe the formulation and validation of a three-dimensional integral plume model appropriate for modeling aircraft exhaust plumes at airports. We also develop a simplified concentration correction factor approach to efficiently account for dispersion processes particular to aircraft plumes. The model is used to explain monitoring station results in the London Heathrow area showing that pollutant concentrations are approximately constant over wind speeds of 3-12 m s-1, and is applied to reproduce empirically derived relationships between engine types and peak NOx concentrations at Heathrow. We calculated that not accounting for aircraft plume dynamics would result in a factor of 1.36-2.3 over-prediction of the mean NOx concentration (depending on location), consistent with empirical evidence of a factor of 1.7 over-prediction. Concentration correction factors are also calculated for aircraft takeoff, landing and taxi emissions, providing an efficient way to account for aircraft plume effects in atmospheric dispersion models.

Barrett, Steven R. H.; Britter, Rex E.; Waitz, Ian A.

2013-08-01

337

Exhaust gas recirculation system  

SciTech Connect

An engine exhaust gas recirculation (EGR) system is provided in which a sonic flow EGR valve is moved to open positions to establish a different constant rate of flow at each open position of the EGR valve in response to air pressure acting on a servo means secured to the valve, the air pressure force being controlled by changes in a control vacuum opposing the air pressure force and modified by an air bleed device as a function of changes in engine exhaust gas backpressure levels, to provide an EGR valve movement that varies essentially in proportion to changes in engine air flow.

Rachedi, S.H.

1983-08-30

338

Hyperventilation and exhaustion syndrome.  

PubMed

Chronic stress is among the most common diagnoses in Sweden, most commonly in the form of exhaustion syndrome (ICD-10 classification - F43.8). The majority of patients with this syndrome also have disturbed breathing (hyperventilation). The aim of this study was to investigate the association between hyperventilation and exhaustion syndrome. Thirty patients with exhaustion syndrome and 14 healthy subjects were evaluated with the Nijmegen Symptom Questionnaire (NQ). The participants completed questionnaires about exhaustion, mental state, sleep disturbance, pain and quality of life. The evaluation was repeated 4 weeks later, after half of the patients and healthy subjects had engaged in a therapy method called 'Grounding', a physical exercise inspired by African dance. The patients reported significantly higher levels of hyperventilation as compared to the healthy subjects. All patients' average score on NQ was 26.57 ± 10.98, while that of the healthy subjects was 15.14 ± 7.89 (t = -3.48, df = 42, p < 0.001). The NQ scores correlated strongly with two measures of exhaustion (Karolinska Exhaustion Scale KES r = 0.772, p < 0.01; Shirom Melamed Burnout Measure SMBM r = 0.565, p < 0.01), mental status [Hospital Anxiety and Depression Score (HADS) depression r = 0.414, p < 0.01; HADS anxiety r = 0.627, p < 0.01], sleep disturbances (r = -0.514, p < 0.01), pain (r = -.370, p < 0.05) and poor well-being (Medical Outcomes Survey Short Form 36 questionnaire- SR Health r = -0.529, p < 0.05). In the logistic regression analysis, the variance in the scores from NQ were explained to a high degree (R(2)  = 0.752) by scores in KES and HADS. The brief Grounding training contributed to a near significant reduction in hyperventilation (F = 2.521, p < 0.124) and to significant reductions in exhaustion scores and scores of depression and anxiety. The conclusion is that hyperventilation is common in exhaustion syndrome patients and that it can be reduced by systematic physical therapy such as Grounding. PMID:24134551

Ristiniemi, Heli; Perski, Aleksander; Lyskov, Eugene; Emtner, Margareta

2014-12-01

339

Hotspots: Mantle Thermal Plumes  

NSDL National Science Digital Library

This article discusses the idea of 'hot spot' volcanoes, those not associated with plate tectonic boundaries, but rather with relatively stationary sources of heat energy (thermal plumes) in the mantle. Topics include the development of the theory by Canadian geophysicist J. Tuzo Wilson; the mechanics of volcanism over a hot spot as seen in the Hawaiian Islands; ancient Hawaiian observations of the ages of their islands; and the distribution of other hot spots around the world.

340

Contrail formation and impacts on aerosol properties in aircraft plumes: Effects of fuel sulfur content  

Microsoft Academic Search

The formation and evolution of fine particles and ice contrails in an aircraft exhaust plume containing varying amounts of fuel sulfur have been simulated using an advanced aerosol microphysics model. The ``core'' sulfate and soot particles are tracked during the contrail formation and dissipation phases. When ion electrostatic effects are incorporated into the microphysics, sulfuric acid vapor emitted by high-sulfur-content

Fangqun Yu; Richard P. Turco

1998-01-01

341

Designing stainless exhaust systems  

SciTech Connect

With the ever-increasing price of automobiles, durability and reduced operating costs have become major concerns in North America, Europe, and Japan. In the US, the exhaust system was once thought of as disposable every 3--4 years, but it is now considered a nonreplaceable item for at least 5--7 years, the average time an initial owner keeps a vehicle. Through the mid-1980s, the only stainless steel on most US car exhausts was the downpipe and catalytic converter, and these were due to government warranty mandates. Today, most US passenger car exhaust systems are almost entirely stainless steel, and with the 1996 model year switch of GM light trucks, the average use of stainless alloys in US vehicles will exceed 23 kg per vehicle. The US experience with stainless has shown that certain design considerations can further increase system life and reduce manufacturing problems. Such considerations may also benefit the European situation, which has seen an increase in the use of stainless alloys in exhaust components since tighter pollution laws began taking effect in 1990.

Douthett, J.A.

1995-11-01

342

Exhaust gas reflux apparatus  

SciTech Connect

An exhaust gas reflux apparatus is described comprising: EGR regulating valve means for controlling a reflux amount of an exhaust gas which is refluxed into an intake system of an internal combustion engine; having at least two operating states; sensing means for sensing the operating state of the engine; first control means responsive to the sensing means for detecting the start of a predetermined one of the operating states of the engine, in which the temperature of the exhaust gas becomes high; timer means responsive to detection by the first control means of the start of the predetermined operating state for measuring a time delay, and time delay beginning immediately after and only in response to the detection of the start of the predetermined operating state and expiring a predetermined time interval thereafter; and second control means responsive to the timer means for controlling the EGR regulating valve means to stop the reflux of the exhaust gas immediately after expiration of the predetermined time interval.

Osada, A.

1988-01-19

343

Exhaust gas recirculation system  

Microsoft Academic Search

An engine exhaust gas recirculation (EGR) system is provided in which a sonic flow EGR valve is moved to open positions to establish a different constant rate of flow at each open position of the EGR valve in response to air pressure acting on a servo means secured to the valve, the air pressure force being controlled by changes in

Rachedi

1983-01-01

344

Exhaust gas recirculation system  

Microsoft Academic Search

The EGR control system comprises EGR control valve means including a first fluid chamber the vacuum in which increases and decreases in accordance with operating conditions of the engine whereby EGR control valve means control the recirculated amount of exhaust gases back to the engine and a second fluid chamber receiving therein a suction vacuum to thereby cause multiplication of

Aoyama

1979-01-01

345

Exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system is described for an engine having a throttle valve in an intake passage, comprising: the intake passage having an EGR port provided adjacent to the upstream side of the throttle valve at the closed position. A control port provides a position at downstream of the throttle valve at the closed position. A leak port provides

Sugiura

1987-01-01

346

Exhaust gas recirculation system  

Microsoft Academic Search

Passage means are described for communicating an EGR passage between first and second EGR control valves. A vacuum chamber of an actuator of the second EGR control valve is prevented from being clogged by solids of the engine exhaust gases by the provision of an air pump for feeding air to fill the passage means between an orifice therein and

Aoyama

1977-01-01

347

Understanding Exhaustive Pattern Learning  

Microsoft Academic Search

Pattern learning in an important problem in Natural Language Processing (NLP). Some exhaustive pattern learning (EPL) methods (Bod, 1992) were proved to be flawed (Johnson, 2002), while similar algorithms (Och and Ney, 2004) showed great advantages on other tasks, such as machine translation. In this article, we first formalize EPL, and then show that the probability given by an EPL

Libin Shen

2011-01-01

348

EUVS Sounding Rocket Payload  

NASA Technical Reports Server (NTRS)

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

Stern, Alan S.

1996-01-01

349

Deuterium microbomb rocket propulsion  

E-print Network

Large scale manned space flight within the solar system is still confronted with the solution of two problems: 1. A propulsion system to transport large payloads with short transit times between different planetary orbits. 2. A cost effective lifting of large payloads into earth orbit. For the solution of the first problem a deuterium fusion bomb propulsion system is proposed where a thermonuclear detonation wave is ignited in a small cylindrical assembly of deuterium with a gigavolt-multimegampere proton beam, drawn from the magnetically insulated spacecraft acting in the ultrahigh vacuum of space as a gigavolt capacitor. For the solution of the second problem, the ignition is done by argon ion lasers driven by high explosives, with the lasers destroyed in the fusion explosion and becoming part of the exhaust.

Friedwardt Winterberg

2008-12-02

350

Deuterium microbomb rocket propulsion  

E-print Network

Large scale manned space flight within the solar system is still confronted with the solution of two problems: 1. A propulsion system to transport large payloads with short transit times between different planetary orbits. 2. A cost effective lifting of large payloads into earth orbit. For the solution of the first problem a deuterium fusion bomb propulsion system is proposed where a thermonuclear detonation wave is ignited in a small cylindrical assembly of deuterium with a gigavolt-multimegampere proton beam, drawn from the magnetically insulated spacecraft acting in the ultrahigh vacuum of space as a gigavolt capacitor. For the solution of the second problem, the ignition is done by argon ion lasers driven by high explosives, with the lasers destroyed in the fusion explosion and becoming part of the exhaust.

Winterberg, Friedwardt

2008-01-01

351

Rocket Science 101 Interactive Educational Program  

NASA Technical Reports Server (NTRS)

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.

Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

2007-01-01

352

Aircraft emissions, plume chemistry, and alternative fuels: results from the APEX, AAFEX, and MDW-2009 campaigns  

NASA Astrophysics Data System (ADS)

We describe observations of aircraft emissions from the APEX, JETS-APEX2, APEX3, MDW-2009 and AAFEX campaigns. Direct emissions of HOx precursors are important for understanding exhaust plume chemistry due to their role in determining HOx concentrations. Nitrous acid (HONO) and formaldehyde are crucial HOx precursors and thus drivers of plume chemistry. At idle power, aircraft engine exhaust is unique among fossil fuel combustion sources due to the speciation of both NOx and VOCs. The impacts of emissions of HOx precursors on plume chemistry at low power are demonstrated with empirical observations of rapid NO to NO2 conversion, indicative of rapid HOx chemistry. The impacts of alternative fuels (derived from biomass, coal, and natural gas) on emissions of NOx, CO, and speciated VOCs are discussed.

Wood, E. C.; Herndon, S. C.; Timko, M.; Yu, Z.; Miake-Lye, R. C.; Lee, B. H.; Santoni, G.; Munger, J. W.; Wofsy, S.; Anderson, B.; Knighton, W. B.

2009-12-01

353

Thrust Performance Improvement for a Water/Liquid Nitrogen Rocket Engine  

NASA Astrophysics Data System (ADS)

We propose a water/liquid nitrogen rocket engine as a new non-combustion type rocket engine. Liquid nitrogen is mixed with heated water and specific volume of nitrogen is increased by evaporation. Thrust force is obtained by exhaust of nitrogen gas through a nozzle with water particles. Results of previous experiments indicated a specific impulse is 60 % of the theoretically estimated value. By evaluating the characteristic exhaust velocity and other thrust characteristics, we found that the lower-than-expected specific impulse is due to insufficient propellant mixing and heat transfer between heated water and liquid nitrogen in the mixing chamber. We also performed high-speed imaging experiments to visualize impinging and mixing of propellants. Results indicate that in the original injection setup, heat conveyed by heated water is not adequately transferred to the liquid nitrogen. An alternative injection pattern was tested, which resulted in a 10% increase in the characteristic exhaust velocity. In addition, we tested a new type of injector designed for more efficient mixing and heat transfer that exhibited 30 % increase in characteristic exhaust velocity. Furthermore, we modified the theoretical expression for the characteristic exhaust velocity based on multi-phased flow theory so that it agrees well with the experimental results.

Watanabe, Rikio; Mikami, Ryo

354

Satellite Observations of Space Shuttle Main Engine Exhaust: Vertical Diffusion and Meridional Transport  

NASA Astrophysics Data System (ADS)

The Sounding of the Atmosphere using Broadband Emission Radiometry (SABER) experiment on NASA’s Thermosphere Ionosphere Mesosphere Energetics and Dynamics (TIMED) satellite has observed water vapor radiances near 6.6 microns on the Earth’s limb since the TIMED launch in December, 2001. Following a space shuttle launch, SABER typically observes enhanced water vapor emission between 90-110 km altitude near the east coast of the United States, where the shuttle injects about 300 metric tons of water vapor exhaust from its three main engines. SABER has observed plumes from 20 space shuttle launches since 2002, all within 25 hours of injection. The database of observations now consists of over 80 separate plume scans, each of which is identified with a peak altitude, a peak brightness and a plume thickness. We compare these SABER shuttle plume observations with a two-dimensional diffusion model that includes photodissociation to determine whether the time evolution of the plume altitude and thickness can be reproduced. Some observations indicate that the shuttle plume is subject to rapid meridional transport. We compare the inferred meridional motion of the plumes with a satellite-derived wind climatology. We include the effects of tidal variability on the shuttle plume and determine whether there is a time of year during which the wind climatology better explains the observed meridional transport.

Stevens, M. H.; Meier, R. R.; Plane, J. M.; Emmert, J. T.; Russell, J.

2010-12-01

355

Exhaust gas recirculation control system  

Microsoft Academic Search

An EGR control valve is closably disposed in an EGR passageway connecting an intake passageway and an exhaust gas passageway which leads to an internal combustion engine. The EGR control valve is operated to control recirculated exhaust gas flow by varying the exhaust gas pressure in a chamber between a restriction disposed in the EGR passageway and the EGR valve,

Aoyama

1979-01-01

356

Multistage exhaust gas recirculation system  

Microsoft Academic Search

An automotive type exhaust gas recirculation (EGR) system has two modes of operation, a first one that regulates EGR flow at a constant percentage rate as a function of throttle valve position independently of exhaust gas backpressure changes, and a second one that provides a variable percentage rate of flow of EGR gases in response to changes in exhaust gas

D. C. Ahrns; S. H. Rachedi

1983-01-01

357

Exhaust gas recirculation control system  

Microsoft Academic Search

An EGR control valve is closely disposed in an EGR passageway connecting an intake passageway and an exhaust gas passageway which leads to an internal combustion engine. The EGR control valve is operated to control recirculated exhaust gas flow by varying the exhaust gas pressure in a chamber between a restriction disposed in the EGR passageway and the EGR valve,

Aoyama

1978-01-01

358

Nuclear thermal rockets using indigenous extraterrestrial propellants  

NASA Technical Reports Server (NTRS)

A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS.

Zubrin, Robert M.

1990-01-01

359

A catalogue of deep mantle plumes: New results from finite-frequency tomography  

NASA Astrophysics Data System (ADS)

New finite-frequency tomographic images of S-wave velocity confirm the existence of deep mantle plumes below a large number of known hot spots. We compare S-anomaly images with an updated P-anomaly model. Deep mantle plumes are present beneath Ascension, Azores, Canary, Cape Verde, Cook Island, Crozet, Easter, Kerguelen, Hawaii, Samoa, and Tahiti. Afar, Atlantic Ridge, Bouvet(Shona), Cocos/Keeling, Louisville, and Reunion are shown to originate at least below the upper mantle if not much deeper. Plumes that reach only to midmantle are present beneath Bowie, Hainan, Eastern Australia, and Juan Fernandez; these plumes may have tails too thin to observe in the lowermost mantle, but the images are also consistent with an interpretation as "dying plumes" that have exhausted their source region. In the tomographic images, only the Eifel and Seychelles plumes are unambiguously confined to the upper mantle. Starting plumes are visible in the lowermost mantle beneath South of Java, East of Solomon, and in the Coral Sea. All imaged plumes are wide and fail to show plumeheads, suggesting a very weakly temperature-dependent viscosity for lower mantle minerals, and/or compositional variations. The S-wave velocity images show several minor differences with respect to the earlier P-wave results, including plume conduits that extend down to the core-mantle boundary beneath Cape Verde, Cook Island, and Kerguelen. A more substantial disagreement between P-wave and S-wave images reopens the question on the depth extent of the Iceland plume. We suggest that a pulsating behavior of the plume may explain the shape of the conduit beneath Iceland.

Montelli, R.; Nolet, G.; Dahlen, F. A.; Masters, G.

2006-11-01

360

An Acoustic Impedance Model for Evaluating Ground Effect of Static-Firing Tests on Rocket Motors  

NASA Astrophysics Data System (ADS)

Evaluation of ground effect is important for acoustic measurement in static-firing tests on rocket motors. The effectiveness of existing acoustic impedance models is examined by comparing with some experimental results. Through the comparison and evaluation of effect of meteorological condition, it is confirmed that the existing impedance models are unsatisfactory for the evaluation of long distance propagation over a hard surface, which corresponds to the far field condition in the present static-firing tests of rocket motors. In this study, a new acoustic impedance model is proposed. From the comparison with the existing acoustic measurement data, it is shown that the new model is effective for both of near and far field propagation. The proposed model is applied to acoustic data measured in the static-firing tests of solid rocket motors, assuming distributed acoustic sources along the exhaust jet axis.

Fukuda, Kota; Tsutsumi, Seiji; Ui, Kyoichi; Ishii, Tatsuya; Takaki, Ryoji; Fujii, Kozo

361

Development of 90 kgf Class CAMUI Hybrid Rocket for a CanSat Experiment  

NASA Astrophysics Data System (ADS)

A newly designed CAMUI hybrid rocket motor of 900 N (90 kgf) thrust class, CAMUI-90, was developed. It uses a combination of polyethylene and liquid oxygen as propellants. CAMUI hybrid rocket is an explosive-flee small rocket motor to realize a small launch system with low cost and flexibility. The motor produces a thrust of 900 N for four seconds, keeping the optimal characteristic exhaust velocity of the fuel-oxidizer combination (exceeding 1800 m/s). A main application of the CAMUI-90 motor is for a CanSat experiment. A launch vehicle employing CAMUI-90 motor, 120 mm in diameter and 3.05 m in length, accelerates a payload of 500 g to 140 m/s in four seconds and reaches to an altitude of about 1 km. The first launch of this vehicle was on December 2006.

Nagata, Harunori; Uematsu, Tsutomu; Ito, Mitsunori; Kakikura, Akihito; Kaneko, Yudai; Mori, Kazuhiro; Murai, Norikazu; Sato, Tatsuhiro; Mitsuhashi, Ryuichi; Totani, Tsuyoshi

362

Small-Scale Rocket Motor Test  

NASA Video Gallery

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

363

Black Carbon Emissions by Rocket Engines Types of rocket engines Emissions  

E-print Network

Black Carbon Emissions by Rocket Engines Types of rocket engines Emissions Liquid Hydrogen and Oxygen Mainly 2 , and some . Aluminum/Ammonium Perchlorate and 23 Rockets that use hydrazine (24) and tetroxide (24) Large amounts of nitrogen oxides. Kerosene Rockets 2 and black carbon (soot). Focus: New

Toohey, Darin W.

364

Automated Rocket Propulsion Test Management  

NASA Technical Reports Server (NTRS)

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

Walters, Ian; Nelson, Cheryl; Jones, Helene

2007-01-01

365

Saving Lives With Rocket Power  

NASA Technical Reports Server (NTRS)

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

2000-01-01

366

Low-thrust rocket trajectories  

NASA Astrophysics Data System (ADS)

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

Keaton, P. W.

1986-01-01

367

Small rocket research and technology  

NASA Technical Reports Server (NTRS)

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

Schneider, Steven; Biaglow, James

1993-01-01

368

14 CFR 29.1123 - Exhaust piping.  

...CATEGORY ROTORCRAFT Powerplant Exhaust System § 29.1123 Exhaust piping. (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Exhaust...

2014-01-01

369

14 CFR 29.1123 - Exhaust piping.  

Code of Federal Regulations, 2011 CFR

...Exhaust System § 29.1123 Exhaust piping. (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Exhaust piping must be supported to...

2011-01-01

370

14 CFR 29.1123 - Exhaust piping.  

Code of Federal Regulations, 2010 CFR

...Exhaust System § 29.1123 Exhaust piping. (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Exhaust piping must be supported to...

2010-01-01

371

14 CFR 27.1123 - Exhaust piping.  

Code of Federal Regulations, 2011 CFR

...Exhaust System § 27.1123 Exhaust piping. (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Exhaust piping must be supported to...

2011-01-01

372

14 CFR 27.1123 - Exhaust piping.  

Code of Federal Regulations, 2010 CFR

...Exhaust System § 27.1123 Exhaust piping. (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Exhaust piping must be supported to...

2010-01-01

373

Measuring Model Rocket Engine Thrust Curves  

ERIC Educational Resources Information Center

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…

Penn, Kim; Slaton, William V.

2010-01-01

374

Alka-Seltzer Rockets Film Canisters & Caps  

E-print Network

Alka-Seltzer Rockets Materials · Film Canisters & Caps · 8" Index Cards · Alka-Seltzer tablets · Markers · Water · Measuring spoons · Scissors · Tape · Cups Part 1: Preparing the Rocket 1. Have each student color one side of an index card (this is the body of the rocket) 2. Roll the index card around

Benitez-Nelson, Claudia

375

PERFORMANCE ENHANCEMENTS ON A PULSED DETONATION ROCKET  

E-print Network

PERFORMANCE ENHANCEMENTS ON A PULSED DETONATION ROCKET The members of the Committee approve #12;To Grandma and Grandpa #12;PERFORMANCE ENHANCEMENTS ON A PULSED DETONATION ROCKET by JASON MATTHEW DETONATION ROCKET Publication No. Jason Matthew Meyers, M.S. The University of Texas at Arlington, 2002

Texas at Arlington, University of

376

Hybrid Rocket Propulsion for Future Space Launch  

E-print Network

Hybrid Rocket Propulsion for Future Space Launch May 09, 2008 Aero/Astro 50th Year Anniversary Arif and Astronautics, Stanford University #12;Aero/Astro 50th Year Anniversary Hybrid Rocket Configuration Most Hybrids Reverse Hybrids: Oxidizer: Solid Fuel: Liquid #12;Aero/Astro 50th Year Anniversary Hybrid Rocket System

Stanford University

377

Apollo Video Photogrammetry Estimation Of Plume Impingement Effects  

NASA Technical Reports Server (NTRS)

The Constellation Project's planned return to the moon requires numerous landings at the same site. Since the top few centimeters are loosely packed regolith, plume impingement from the Lander ejects the granular material at high velocities. Much work is needed to understand the physics of plume impingement during landing in order to protect hardware surrounding the landing sites. While mostly qualitative in nature, the Apollo Lunar Module landing videos can provide a wealth of quantitative information using modem photogrammetry techniques. The authors have used the digitized videos to quantify plume impingement effects of the landing exhaust on the lunar surface. The dust ejection angle from the plume is estimated at 1-3 degrees. The lofted particle density is estimated at 10(exp 8)- 10(exp 13) particles per cubic meter. Additionally, evidence for ejection of large 10-15 cm sized objects and a dependence of ejection angle on thrust are presented. Further work is ongoing to continue quantitative analysis of the landing videos.

Immer, Christopher; Lane, John; Metzger, Philip T.; Clements, Sandra

2008-01-01

378

Simple models of tropical plumes  

E-print Network

Tropical plumes are upper and mid-level cloud bands at least 2000 km long that cross 15' latitude. The simplest conditions that lead to tropical plume development are sought in a barotropic model simulating winter 200 mb flow. The features sought...

Carrie, Gordon David, d 1960-

2012-06-07

379

Commercial Development Suborbital Rocket Program  

NASA Technical Reports Server (NTRS)

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

1993-01-01

380

Robert Goddard and His Rockets  

NSDL National Science Digital Library

This page, part of "From Stargazers to Starships," is devoted to the story of Robert H. Goddard and his contributions. Biographical details of Goddard's life and career are provided. This page also features an explanation of how Goddard derived the efficiency of his rockets, links to further exploration of Goddard, and a glossary of terms from the web page.

Stern, David

2007-01-03

381

Launch Excitement with Water Rockets  

ERIC Educational Resources Information Center

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

Sanchez, Juan Carlos; Penick, John

2007-01-01

382

Laser-heated rocket studies  

NASA Technical Reports Server (NTRS)

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

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

1976-01-01

383

Photon rockets and gravitational radiation  

E-print Network

The absence of gravitational radiation in Kinnersley's ``photon rocket'' solution of Einstein's equations is clarified by studying the mathematically well-defined problem of point-like photon rockets in Minkowski space (i.e. massive particles emitting null fluid anisotro\\-pically and accelerating because of the recoil). We explicitly compute the (uniquely defined) {\\it linearized} retarded gravitational waves emitted by such objects, which are the coherent superposition of the gravitational waves generated by the motion of the massive point-like rocket and of those generated by the energy-momentum distribution of the photon fluid. In the special case (corresponding to Kinnersley's solution) where the anisotropy of the photon emission is purely dipolar we find that the gravitational wave amplitude generated by the energy-momentum of the photons exactly cancels the usual $1/r$ gravitational wave amplitude generated by the accelerated motion of the rocket. More general photon anisotropies would, however, generate genuine gravitational radiation at infinity. Our explicit calculations show the compatibility between the non-radiative character of Kinnersley's solution and the currently used gravitational wave generation formalisms based on post-Minkowskian perturbation theory.

T. Damour

1994-12-21

384

A comparison of plume images in P- and S-wave tomography  

NASA Astrophysics Data System (ADS)

We present the results of a global, finite-frequency, tomography experiment using long period S,SS and ScS data. The images of S-wave velocity confirm the existence of deep mantle plumes imaged earlier using P-wave tomography (Montelli et al., Science 2004). We have updated the P wave model using improved crustal corrections and compare the two models. Deep mantle plumes are present beneath Ascension, Azores, Canary, Cape Verde, Cook Island, Crozet, Easter, Kerguelen, Hawaii, Samoa and Tahiti, while Afar, Atlantic Ridge, Bouvet(Shona), Cocos/Keeling, Louisville and Reunion are shown to originate at least below the upper mantle if not much deeper. Plumes that reach only to mid-mantle are present beneath Bowie, Hainan, Eastern Australia and Juan Fernandez; these plumes may have tails too thin to observe in the lowermost mantle, but in either case the images are consistent with an interpretation as dying plumes that have exhausted their source region. In the tomographic images only Eifel, and Seychelles are confined to the upper mantle. Starting plumes are visible in the lowermost mantle beneath South of Java, East of Solomon and in the Coral Sea. All imaged plumes are wide an fail to show plumeheads, suggesting a very weak temperature dependence viscosity for lower mantle minerals, and/or compositional variations. The S-wave velocity images show several minor differences with respect to the earlier P-wave results including plume conduits that extend down to the core-mantle boundary beneath Cape Verde, Cook Island and Kerguelen. A more substantial disagreement between P- and S-wave images reopens the question on the depth extent of the Iceland plume. We suggest that a pulsating behaviour of the plume may explain the variability of the images.

Montelli, R.; Nolet, G.; Dahlen, T.; Masters, G.

2005-12-01

385

Electric exhaust gas recirculation valve  

SciTech Connect

An electrically actuated EGR valve is described for controlling EGR gases in response to electric signals from a computer, the EGR valve comprising: a valve housing having an exhaust gas inlet port for passage of exhaust gases; an exhaust gas outlet port; an exhaust gas passage extending between; poppet valve means for selectively opening and closing exhaust gas passage, the poppet valve means including a valve rod slidably supported for movement in a linear direction in the housing; an electric step motor responsive to electric pulse signals and including output means for directly contacting and moving the valve rod; and a motor casing secured to the valve housing with an insulating element. The step motor is isolated from heat from the exhaust gases and motor casing includes apertures formed where air may be circulated into the motor casing to further thermally isolate the step motor.

Akagi, M.

1987-06-23

386

Exhaust gas recirculation system  

SciTech Connect

An exhaust gas recirculation system is described for an engine having a throttle valve in an intake passage, comprising: the intake passage having an EGR port provided adjacent to the upstream side of the throttle valve at the closed position. A control port provides a position at downstream of the throttle valve at the closed position. A leak port provides at upstream of the throttle valve; an EGR valve recirculates exhaust gases to the intake passage. The EGR valve has a diaphragm defining a first chamber applied with the pressure at the EGR port and a second chamber is applied with the pressure at the leak port. A valve body connects to the diaphragm for controlling the amount of recirculated gases; a control valve has a diaphragm defining a first control chamber and a second control chamber, and valve means on the diaphragm. The first control chamber is pressured at the EGR port. The second control chamber is applied with the pressure at the control port through a first conduit having an end port and with the pressure at the leak port through a second conduit. The valve means is arranged to open the end port of the first conduit when the difference between pressures in the first and second control chambers exceeds a predetermined value. Pressure regulating means renders the pressure in the second control chamber lower than the pressure in the second chamber of the EGR valve.

Sugiura, K.

1987-08-04

387

COMPARING AND LINKING PLUMES ACROSS MODELING APPROACHES  

EPA Science Inventory

River plumes carry many pollutants, including microorganisms, into lakes and the coastal ocean. The physical scales of many stream and river plumes often lie between the scales for mixing zone plume models, such as the EPA Visual Plumes model, and larger-sized grid scales for re...

388

A multidisciplinary optimization methodology for rocket vehicle systems  

NASA Astrophysics Data System (ADS)

Rocket vehicles have traditionally been designed in an iterative fashion, beginning with system requirements before proceeding sequentially through requisite analytical disciplines until resources are exhausted. A sequentially designed system, while adequate, is not an optimum due to the approximations and loss of fidelity inherent in separating analytical disciplines which are, in fact, coupled. Recently, increased computational power and advances in algorithms have allowed multidisciplinary optimization (MDO) to emerge as a system-level design tool accessible to industry. To date, MDO has primarily been applied to some facets of aircraft systems and, to a lesser extent, rocket vehicles in literature but has not yet met with widespread industry use. To this end, four obstacles have been identified: (1) MDO efforts to date have focused on system-level parameters rather than physical dimensions and hence have not yielded a preliminary design which includes manufacturing, cost, and other constraints, (2) Prohibitive computational performance requirements associated with high-fidelity analyses such as computational fluid mechanics (CFD) and finite element analysis (FEA), (3) Lack of an integrated design environment which incorporates computational tools already widely used in industry while remaining accessible to individual users without high-level expertise in the individual tools, and (4) The widely-varying and tightly-coupled environments to which rocket vehicles are typically exposed, including analyses not required for aircraft applications. Here, an MDO method for rocket systems has been formulated which simultaneously overcomes the challenges listed above. First, a response surface-based approach to modeling computationally expensive analyses with arbitrary dimensionality and general constraints was developed. This method focused on an evenly-distributed representation of the entire feasible region at any fidelity level, including combinations of discrete and continuous variables. The analytical disciplines required in the design of a general rocket vehicle were then developed, focusing on computational cost and multi-fidelity methods were applicable. Finally, this integrated framework was applied to three diverse case studies. Where possible, the results obtained were compared to traditional design methods demonstrating considerable performance gains while maintaining manageable computational cost. Within this framework, many opportunities for improvement and future directions were noted, both in the analytical disciplines and optimization architecture as a whole.

Colonno, Michael Richard

389

Finiteness of steady state plumes  

NASA Astrophysics Data System (ADS)

The finite maximum length of a steady state contaminant plume is determined by developing and employing a new analytical solution which overcomes two drawbacks associated with existing approaches. First, we account for a sharp front caused by the complete consumption of the pollutant ("electron donor") and some electron acceptor in an instantaneous binary reaction occurring at the front. This approach is not based on purely conservative or first-order degradation models which lead to theoretically infinite plumes and, in addition, depend on a concentration threshold. Second, a vertical aquifer cross section with finite thickness is selected as a model in order to better represent the supply of electron acceptors mostly entering the aquifer from the top. This type of setting allows investigation of the impact of aquifer thickness on plume length. An implicit representation of the donor-acceptor front in a finite vertical domain previously required numerical solutions of the underlying advection-dispersion-reaction equation; we provide for the first time an analytical solution of this two-dimensional transport problem. The length of the plume is found to be given by the point of intersection of the donor-acceptor front and the aquifer bottom. Furthermore, a rather simple and highly accurate approximation is derived to compute the steady state plume length. A comprehensive sensitivity analysis reveals that results are most strongly influenced by aquifer thickness, followed by vertical transverse dispersivity and, to a somewhat lesser extent, by chemical reaction parameters. Longitudinal dispersivity has practically no effect on plume length, and furthermore, there is zero impact of linear velocity. With regard to groundwater risk assessment at the field scale it is also important to note that the present approach is meant to provide an upper bound on the actual plume length. Further research activities may be directed to refine the transport model by considering, for instance, degradation inside the plume and the limited vertical extent of the contaminant source.

Liedl, Rudolf; Valocchi, Albert J.; Dietrich, Peter; Grathwohl, Peter

2005-12-01

390

Diesel Engine Exhaust Emissions  

Microsoft Academic Search

\\u000a The direct release of exhaust gas components from combustion processes into the environment, i.e. \\u000a \\u000a \\u000a \\u000a emission, is the primary and most important process in the chain of emission, \\u000a \\u000a \\u000a \\u000a transmission, \\u000a \\u000a \\u000a \\u000a pollutant input and \\u000a \\u000a \\u000a \\u000a impact. Naturally, a basic distinction is made between emissions from vegetation, oceans, volcanic activity or biomass decomposition\\u000a for instance and anthropogenic emissions, i.e. emissions caused or influenced by humans,

Helmut Tschoeke; Andreas Graf; Jürgen Stein; Michael Krüger; Johannes Schaller; Norbert Breuer; Kurt Engeljehringer; Wolfgang Schindler

391

Low thrust rocket test facility  

NASA Technical Reports Server (NTRS)

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

Arrington, Lynn A.; Schneider, Steven J.

1990-01-01

392

Unique nuclear thermal rocket engine  

SciTech Connect

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

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

1993-01-15

393

Unique nuclear thermal rocket engine  

SciTech Connect

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

Culver, D.W.; Rochow, R.

1993-06-01

394

Plasma core nuclear rocket technology  

Microsoft Academic Search

The nuclear lightbulb (NLB) rocket propulsion concept furnishes specific impulse above 2000 sec in conjunction with the greater-than-50,000 lb thrust levels required for rapid transit-time round-trip Mars missions requiring low initial mass in earth orbit. The NLB transfers energy from the gaseous nuclear fuel region to a hydrogen propellant via thermal radiation, thereby precluding material temperature constraints. An evaluation is

Thomas S. Latham; Ward C. Roman; Bruce V. Johnson

1993-01-01

395

Rocket Engine Altitude Simulation Technologies  

NASA Technical Reports Server (NTRS)

John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

Woods, Jody L.; Lansaw, John

2010-01-01

396

Engine exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for an internal combustion engine employs a recirculation control valve in a passageway connecting the engine exhaust passage to the engine intake passage downstream from the throttle valve. An air conduit having an air control valve therein furnishes atmospheric air to the intake passage downstream from the throttle valve. Vacuum responsive actuators are provided for

H. Nishimura; T. Shioya; T. Umemoto

1981-01-01

397

Engine exhaust gas recirculation system  

Microsoft Academic Search

An exhaust gas recirculation system for an internal combustion engine employs a recirculation control valve in a passageway connecting the engine exhaust passage to the engine intake passage downstream from the throttle valve. An air conduit having an air control valve therein draws atmospheric air into the intake passage downstream from the throttle valve. Vacuum responsive actuators are provided for

K. Ishii; H. Nishimura; K. Osawa

1981-01-01

398

Power Exhaust in Fusion Plasmas  

NASA Astrophysics Data System (ADS)

Preface; 1. Introduction; 2. Magnetized plasma physics; 3. Magnetized plasma equilibrium; 4. Magnetized plasma stability; 5. Collisional transport in magnetized plasmas; 6. Turbulent transport in magnetized plasmas; 7. Tokamak plasma boundary and power exhaust; 8. Outlook: power exhaust in fusion reactors; Appendix A. Maxwellian distribution; Appendix B. Curvilinear co-ordinates; References; Index.

Fundamenski, Wojciech

2014-07-01

399

Exhaust gas recirculaton control system  

Microsoft Academic Search

An exhaust gas recirculation (EGR) control system for an internal combustion engine is presented in which an EGR control valve including a vacuum-operated actuator governs the flow of recirculated exhaust gas according to the admission of air into the engine. To accomplish the control with high precision, the control system includes a flow sensor which provides an electrical signal representing

K. Maruyama; Y. Hata; A. Ohnishi

1979-01-01

400

Exhaust gas recirculation control system  

Microsoft Academic Search

A multiple spark plug ignition internal combustion engine is equipped with an EGR control system which consists of an EGR control valve disposed in an EGR passageway connecting an exhaust gas passageway and an intake passageway. The EGR control system is arranged to control EGR rate in accordance with venturi vacuum and in accordance with the exhaust gas pressure in

Aoyama

1979-01-01

401

Automotive Fuel and Exhaust Systems.  

ERIC Educational Resources Information Center

Materials are provided for a 14-hour course designed to introduce the automotive mechanic to the basic operations of automotive fuel and exhaust systems incorporated on military vehicles. The four study units cover characteristics of fuels, gasoline fuel system, diesel fuel systems, and exhaust system. Each study unit begins with a general…

Irby, James F.; And Others

402

Distributed Exhaust Nozzles for Jet Noise Reduction  

NASA Technical Reports Server (NTRS)

The main objective of this study is to validate the jet noise reduction potential of a concept associated with distributed exhaust nozzles. Under this concept the propulsive thrust is generated by a larger number of discrete plumes issuing from an array of small or mini-nozzles. The potential of noise reduction of this concept stems from the fact that a large number of small jets will produce very high frequency noise and also, if spaced suitably, they will coalesce at a smaller velocity to produce low amplitude, low frequency noise. This is accomplished through detailed acoustic and fluid measurements along with a Computational Fluidic Dynamic (CFD) solution of the mean (DE) Distributed Exhaust nozzle flowfield performed by Northrop-Grumman. The acoustic performance is quantified in an anechoic chamber. Farfield acoustic data is acquired for a DE nozzle as well as a round nozzle of the same area. Both these types of nozzles are assessed numerically using Computational Fluid Dynamic (CFD) techniques. The CFD analysis ensures that both nozzles issued the same amount of airflow for a given nozzle pressure ratio. Data at a variety of nozzle pressure ratios are acquired at a range of polar and azimuthal angles. Flow visualization of the DE nozzle is used to assess the fluid dynamics of the small jet interactions. Results show that at high subsonic jet velocities, the DE nozzle shifts its frequency of peak amplitude to a higher frequency relative to a round nozzle of equivalent area (from a S(sub tD) = 0.24 to 1. 3). Furthermore, the DE nozzle shows reduced sound pressure levels (as much as 4 - 8 dB) in the low frequency part of the spectrum (less than S(sub tD) = 0.24 ) compared to the round nozzle. At supersonic jet velocities, the DE nozzle does not exhibit the jet screech and the shock-associated broadband noise is reduced by as much as 12 dB.

Ahuja, K. K.; Gaeta, R. J.; Hellman, B.; Schein, D. B.; Solomon, W. D., Jr.; Huff, Dennis (Technical Monitor)

2001-01-01

403

Multicomponent remote sensing of vehicle exhaust emissions by dispersive IR and UV spectroscopy  

NASA Astrophysics Data System (ADS)

Direct remote sensing of vehicle exhaust emissions under real-world driving conditions is desirable for a number of reasons, including: identifying high emitters, investigating the chemical composition of the exhaust, and probing fast reactions in the plume. A remote sensor, incorporating IR and UV spectrometers, was developed. The IR spectrometer consists of a grating system mounted on a synchronous motor, optically interfaced to a room temperature PbSe detector. UV-vis measurements are made with a CCD array spectrometer. Eight optical passes through the exhaust plume allow rapid and sensitive monitoring of the exhaust stream emitted by moving vehicles on a car-by-car basis. The combination of these two techniques resulted in unprecedented, direct measurement capability of over 25 pollutants in the exhaust plume. Emissions from a fleet of vehicles powered by a range of fuels (gasoline, diesel, natural gas, and methanol) were tested. The exhaust from hot gasoline- and methanol-powered cars contained high levels of NH3, up to 1500 ppm. These emissions were up to 14 times higher than the corresponding NOx emissions. Unlike most previous work, NOx was measured as the sum of NO and NO2; N2O was also measured. Field testing at a southern California freeway on-ramp was conducted over a one week period, totaling >4,500 measurements. It was found that 66.4% of the emitted NH3 was produced by 10% of the fleet, following the (gamma) - distribution that has been reported for criteria pollutants. Mean NH3 emission rates were calculated at 138 mg km-1, nearly twice as high was previous estimates.

Baum, Marc M.; Kiyomiya, Eileen S.; Kumar, Sasi; Lappas, Anastasios M.; Lord, Harry C., III

2000-12-01

404

Active Volcanic Plumes on Io  

NASA Technical Reports Server (NTRS)

This color image, acquired during Galileo's ninth orbit around Jupiter, shows two volcanic plumes on Io. One plume was captured on the bright limb or edge of the moon (see inset at upper right), erupting over a caldera (volcanic depression) named Pillan Patera after a South American god of thunder, fire and volcanoes. The plume seen by Galileo is 140 kilometers (86 miles) high and was also detected by the Hubble Space Telescope. The Galileo spacecraft will pass almost directly over Pillan Patera in 1999 at a range of only 600 kilometers (373 miles).

The second plume, seen near the terminator (boundary between day and night), is called Prometheus after the Greek fire god (see inset at lower right). The shadow of the 75-kilometer (45- mile) high airborne plume can be seen extending to the right of the eruption vent. The vent is near the center of the bright and dark rings. Plumes on Io have a blue color, so the plume shadow is reddish. The Prometheus plume can be seen in every Galileo image with the appropriate geometry, as well as every such Voyager image acquired in 1979. It is possible that this plume has been continuously active for more than 18 years. In contrast, a plume has never been seen at Pillan Patera prior to the recent Galileo and Hubble Space Telescope images.

North is toward the top of the picture. The resolution is about 6 kilometers (3.7 miles) per picture element. This composite uses images taken with the green, violet and near infrared filters of the solid state imaging (CCD) system on NASA's Galileo spacecraft. The images were obtained on June 28, 1997, at a range of more than 600,000 kilometers (372,000 miles).

The Jet Propulsion Laboratory, Pasadena, CA manages the Galileo mission for NASA's Office of Space Science, Washington, DC. JPL is an operating division of California Institute of Technology (Caltech).

This image and other images and data received from Galileo are posted on the World Wide Web, on the Galileo mission home page at URL http://galileo.jpl.nasa.gov. Background information and educational context for the images can be found at URL http://www.jpl.nasa.gov/galileo/sepo

1997-01-01

405

Acoustical and Flowfield Characterization of a Scaled Tabletop Rocket  

NASA Technical Reports Server (NTRS)

An analysis of the acoustical and flowfield environment for the scaled 1-pound-force (lbf) thrust tabletop motor was performed. This tabletop motor from NASA Stennis Space Center Is composed of Plexiglas burning In gaseous oxygen with a graphite insert for the nozzle portion. The nozzle has a throat diameter of 0.2 inch and an exit diameter of 0.38 Inch. With a chamber pressure at 55 pounds per square Inch absolute (psia), a normal shock is formed immediately downstream of the nozzle exit plane as the combustion products exhaust into the ambient at atmospheric pressure. The jet characterization Is based on computational fluid dynamics (CFD) in conjunction with Kirchhoff surface integral formulation and compared with correlations developed for measured rocket noise and a pressure fluctuation scaling (PFS) method. Predictions and comparisons are made for the overall sound pressure levels (OASPL's) and spectral dependence of sound pressure level (SPL). The overall objective of this effort is to develop methods for scaling the acoustic and flowfield environment of rockets with a wide range of thrust (1 lbf to 1 million lbf).

Kandula, Max; Margasahayam, Ravi; Norton, Michael; Caimi, Raoul; Steinrock, T. (Technical Monitor); Venegas, Augusto (Technical Monitor)

2001-01-01

406

Thermal stratification potential in rocket engine coolant channels  

NASA Technical Reports Server (NTRS)

The potential for rocket engine coolant channel flow stratification was computationally studied. A conjugate, 3-D, conduction/advection analysis code (SINDA/FLUINT) was used. Core fluid temperatures were predicted to vary by over 360 K across the coolant channel, at the throat section, indicating that the conventional assumption of a fully mixed fluid may be extremely inaccurate. Because of the thermal stratification of the fluid, the walls exposed to the rocket engine exhaust gases will be hotter than an assumption of full mixing would imply. In this analysis, wall temperatures were 160 K hotter in the turbulent mixing case than in the full mixing case. The discrepancy between the full mixing and turbulent mixing analyses increased with increasing heat transfer. Both analysis methods predicted identical channel resistances at the coolant inlet, but in the stratified analysis the thermal resistance was negligible. The implications are significant. Neglect of thermal stratification could lead to underpredictions in nozzle wall temperatures. Even worse, testing at subscale conditions may be inadequate for modeling conditions that would exist in a full scale engine.

Kacynski, Kenneth J.

1992-01-01

407

Effects of plume produced by the Nd:YAG laser and electrocautery on the respiratory system.  

PubMed

Sprague-Dawley rats were exposed to Nd:YAG laser exhaust (contact and noncontact) as well as to electrocautery exhaust passed through smoke evacuation filters. Exposure periods for each group were equal and increasing in time. Histologic analysis revealed alveolar congestion and emphysematous changes in all modes. Controls exhibited similar change but to a milder degree. It appears that any plume produced by lasers or electrosurgical devices produces pathologic change in rat lungs and that effective smoke evacuation will help control these effects. PMID:8464311

Wenig, B L; Stenson, K M; Wenig, B M; Tracey, D

1993-01-01

408

Designing An Adaptor To Connect Rocket Stages  

NASA Technical Reports Server (NTRS)

Report describes design of light-weight truss structure to serve as adaptor to connect two rocket stages. Larger stage is Centaur rocket, on which provided eight attachment points arranged in circle. Smaller stage is Thiokol Star 48b rocket, which includes round attachment flange. For mating with attachment flange, design provides for grooved slats at upper ends of struts. Weight of truss structure less than corresponding stiffened conical shell.

Gann, Lisa L.; Hicks, Michael T.

1995-01-01

409

SHARPI/PICTURE Sounding Rocket Telescope  

NASA Technical Reports Server (NTRS)

The Solar High Angular Resolution Photometric Imager (SHARPI)/Planet Imaging Concept Testbed Using a Rocket Experiment (PICTURE) Sounding Rocket Telescope is described. The topics include: 1) Lightweight precision mirror development; 2) Two sounding rocket concepts sharing a telescope; 3) Optical Telescope Assembly (OTA) overview; 4) PM development program; 5) PM figure testing; 6) Mirror coatings; 7) PM mount and verification; 8) Secondary Mirror (SM); and 9) OTA.

Content, D.; Antonille, S.; Wallace, T.; Rabin, D.; Wake, S.

2006-01-01

410

Plume effects on the flow around a blunted cone at hypersonic speeds  

NASA Technical Reports Server (NTRS)

Tests at M = 8.2 show that a simulated rocket plume at the base of a blunted cone can cause large areas of separated flow, with dramatic effects on the heat transfer rate distribution. The plume was simulated by solid discs of varying sizes or by an annular jet of gas. Flow over the cone without a plume is fully laminar and attached. Using a large disc, the boundary layer is laminar at separation at the test Reynolds number. Transition occurs along the separated shear layer and the boundary layer quickly becomes turbulent. The reduction in heat transfer associated with a laminar separated region is followed by rising values as transition occurs and the heat transfer rates towards the rear of the cone substantially exceed the values obtained without a plume. With the annular jet or a small disc, separation occurs much further aft, so that heat transfer rates at the front of the cone are comparable with those found without a plume. Downstream of separation the shear layer now remains laminar and the heat transfer rates to the surface are significantly lower than the attached flow values.

Atcliffe, P.; Kumar, D.; Stollery, J. L.

1992-01-01

411

NASA Sounding Rockets and Hi-C  

NASA Video Gallery

The Sounding Rockets Program Office (SRPO), located at NASA Goddard Space Flight Center's Wallops Flight Facility, provides suborbital launch vehicles, payload development, and field operations sup...

412

A miniature solid propellant rocket motor  

SciTech Connect

A miniature solid-propellant rocket motor has been developed to impart a specific motion to an object deployed in space. This rocket motor effectively eliminated the need for a cold-gas thruster system or mechanical spin-up system. A low-energy igniter, an XMC4397, employing a semiconductor bridge was used to ignite the rocket motor. The rocket motor was ground-tested in a vacuum tank to verify predicted space performance and successfully flown in a Sandia National Laboratories flight vehicle program.

Grubelich, M.C.; Hagan, M.; Mulligan, E.

1997-08-01

413

Focused Rocket-Ejector RBCC Experiments  

NASA Technical Reports Server (NTRS)

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

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

414

14 CFR 23.1123 - Exhaust system.  

...AIRPLANES Powerplant Exhaust System § 23.1123 Exhaust system. (a) Each exhaust system must be fireproof and corrosion-resistant, and must have means to prevent failure due to expansion by operating temperatures. (b) Each...

2014-01-01

415

77 FR 50584 - Voluntary Licensing of Amateur Rocket Operations  

Federal Register 2010, 2011, 2012, 2013

...Voluntary Licensing of Amateur Rocket Operations AGENCY: Federal...that conduct certain amateur rocket launches an opportunity to...the FAA notes that amateur rocket operators would incur costs...RFA) establishes ``as a principle of regulatory issuance...

2012-08-22

416

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2011 CFR

...FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS 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...

2011-01-01

417

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2010 CFR

...FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS 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...

2010-01-01

418

Shuttle active thermal control system development testing. Volume 6: Water ejector plume tests  

NASA Technical Reports Server (NTRS)

Results are given of vacuum testing of nozzles designed to eject water vapor away from the space shuttle to prevent contamination of the spacecraft surfaces and payload. The water vapor is generated by an active cooling system which evaporates excess fuel cell water to supplement a modular radiator system (MRS). The complete heat rejection system including the MRS, flash evaporator or sublimator and nozzle were first tested to demonstrate the system operational characteristics. The plume tests were performed in two phases and the objectives of this test series were: (1) to determine the effectiveness of a supersonic nozzle and a plugged nozzle in minimizing impingement upon the spacecraft of water vapor exhausted by an active device (flash evaporator or sublimator); and (2) to obtain basic data on the flow fields of exhaust plumes generated by these active devices, both with and without nozzles installed.

Mcginnis, F. K.; Summerhays, R. M.

1973-01-01

419

Ablation Performance of Carbon/Carbon Composite Throat after a Solid Rocket Motor Ground Ignition Test  

NASA Astrophysics Data System (ADS)

The ablation performances of a fine-woven, pierced carbon/carbon (C/C) composite throats for solid rocket motor were investigated by a ground ignition test. The ablation surface morphologies of three regions (entrance, throat and exit) of the throats were examined in detail by scanning electron microscopy. The results show that the C/C composite throats retain smooth inner surface, experiencing ablation rates of 0.142-0.146 mm/s under a pressure of about 6.0 MPa. But ablation morphologies of the three regions are different, due to the continuously changing of temperatures, velocities, and oxidant concentrations of combustion gas plume.

Yin, Jian; Zhang, Hongbo B.; Xiong, Xiang; Zuo, Jinlv L.; Huang, Baiyun Y.

2012-06-01

420

FTIR airborne measurement of aircraft jet engine exhaust gas emissions under cruise conditions  

NASA Astrophysics Data System (ADS)

A flight qualified Fourier transform infrared spectrometer has been built by applying the MIROR principle (Michelson interferometer with rotating retroreflector) where an eccentrically rotating retroreflector generates optical path differences. The unique optical design is especially suiting the rough environment of airborne missions. The purely optical and passive method in no way influences the gases as sample collecting procedures are likely to do. It is able to deliver true space/time resolved spectra of several species in the exhaust plume simultaneously. First measurements aboard a civil jet aircraft have been performed, successfully collecting spectra of the infrared radiation emitted by the hot exhaust gases just behind the engine's nozzle. The spectra were radiometrically calibrated and column densities of trace gases in the plume and in the for- and background as well as gas temperatures were calculated applying inversion algorithms. From these then emission indices (mass pollutant per mass kerosene) of the engine for specific trace gases were determined.

Tank, Volker; Haschberger, Peter; Lindermeir, Erwin; Matthern, K. H.

1995-09-01

421

The asteroid laser rocket engine  

NASA Astrophysics Data System (ADS)

Recently the scientific community has come to comprehension of gravity of a problem of threat to the Earth from action of asteroids and comets There is a real and probable threat of the future collisions of the Earth with a celestial body Therefore the person to guarantee the future in long-term prospect should be prepared for to reflect such danger Time come of not passive but active action on comets the asteroids which are bringing the threat for the Earth The NASA carried out recently experiment on collision with comet Tempel-1 which is representing the most direct interest for the solution of to take away dangerous bodies for the Earth into other harmless orbits The purpose of the report is a statement of idea of use of space-rocket conveyances for transfer of dangerous celestial bodies for the Earth into other orbit through a creation in their body of a laser rocket engine The similar experience is present already when orbit of station Mir was corrected previously that the station has falled in lonely ocean Certainly it was an insignificant body by astronomical measures But it does not mean that these problems should not be developed The first results on studying a structure of comet Tempel-1 have shown that an ice nucleus is under firm crust In the future it allows to create in a body of a comet well-known laser rocket engine of the enormous sizes by means of laser beams when their power and long-range action begins comprehensible for realization of this idea and by melting ice and evaporating water to receive jet force which

Prisniakov, V.

422

Lidar sounding of volcanic plumes  

NASA Astrophysics Data System (ADS)

Accurate knowledge of gas composition in volcanic plumes has high scientific and societal value. On the one hand, it gives information on the geophysical processes taking place inside volcanos; on the other hand, it provides alert on possible eruptions. For this reasons, it has been suggested to monitor volcanic plumes by lidar. In particular, one of the aims of the FP7 ERC project BRIDGE is the measurement of CO2 concentration in volcanic gases by differential absorption lidar. This is a very challenging task due to the harsh environment, the narrowness and weakness of the CO2 absorption lines and the difficulty to procure a suitable laser source. This paper, after a review on remote sensing of volcanic plumes, reports on the current progress of the lidar system.

Fiorani, Luca; Aiuppa, Alessandro; Angelini, Federico; Borelli, Rodolfo; Del Franco, Mario; Murra, Daniele; Pistilli, Marco; Puiu, Adriana; Santoro, Simone

2013-10-01

423

The Advanced Solid Rocket Motor  

NASA Astrophysics Data System (ADS)

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

Mitchell, Royce E.

1992-08-01

424

Uranium droplet core nuclear rocket  

NASA Technical Reports Server (NTRS)

Uranium droplet nuclear rocket is conceptually designed to utilize the broad temperature range ofthe liquid phase of metallic uranium in droplet configuration which maximizes the energy transfer area per unit fuel volume. In a baseline system dissociated hydrogen at 100 bar is heated to 6000 K, providing 2000 second of Isp. Fission fragments and intense radian field enhance the dissociation of molecular hydrogen beyond the equilibrium thermodynamic level. Uranium droplets in the core are confined and separated by an axisymmetric vortex flow generated by high velocity tangential injection of hydrogen in the mid-core regions. Droplet uranium flow to the core is controlled and adjusted by a twin flow nozzle injection system.

Anghaie, Samim

1991-01-01

425

Reusable Rocket Engine Maintenance Study  

NASA Technical Reports Server (NTRS)

Approximately 85,000 liquid rocket engine failure reports, obtained from 30 years of developing and delivering major pump feed engines, were reviewed and screened and reduced to 1771. These were categorized into 16 different failure modes. Failure propagation diagrams were established. The state of the art of engine condition monitoring for in-flight sensors and between flight inspection technology was determined. For the 16 failure modes, the potential measurands and diagnostic requirements were identified, assessed and ranked. Eight areas are identified requiring advanced technology development.

Macgregor, C. A.

1982-01-01

426

Space Shuttle and Launch Pad Lift-Off Debris Transport Analysis: SRB Plume-Driven  

NASA Technical Reports Server (NTRS)

This paper discusses the Space Shuttle Lift-Off model developed for potential Lift-Off Debris transport. A critical Lift-Off portion of the flight is defined from approximately 1.5 sec after SRB Ignition up to 'Tower Clear', where exhaust plume interactions with the Launch Pad occur. A CFD model containing the Space Shuttle and Launch Pad geometry has been constructed and executed. The CFD model works in conjunction with a debris particle transport model and a debris particle impact damage tolerance model. These models have been used to assess the effects of the Space Shuttle plumes, the wind environment, their interactions with the Launch Pad, and their ultimate effect on potential debris during Lift-Off. Emphasis in this paper is on potential debris that might be caught by the SRB plumes.

West, Jeff; Strutzenberg, Louis; Dougherty, Sam; Radke, Jerry; Liever, Peter

2007-01-01

427

Modeling Leaking Gas Plume Migration  

SciTech Connect

In this study, we obtain simple estimates of 1-D plume propagation velocity taking into account the density and viscosity contrast between CO{sub 2} and brine. Application of the Buckley-Leverett model to describe buoyancy-driven countercurrent flow of two immiscible phases leads to a transparent theory predicting the evolution of the plume. We obtain that the plume does not migrate upward like a gas bubble in bulk water. Rather, it stretches upward until it reaches a seal or until the fluids become immobile. A simple formula requiring no complex numerical calculations describes the velocity of plume propagation. This solution is a simplification of a more comprehensive theory of countercurrent plume migration that does not lend itself to a simple analytical solution (Silin et al., 2006). The range of applicability of the simplified solution is assessed and provided. This work is motivated by the growing interest in injecting carbon dioxide into deep geological formations as a means of avoiding its atmospheric emissions and consequent global warming. One of the potential problems associated with the geologic method of sequestration is leakage of CO{sub 2} from the underground storage reservoir into sources of drinking water. Ideally, the injected green-house gases will stay in the injection zone for a geologically long time and eventually will dissolve in the formation brine and remain trapped by mineralization. However, naturally present or inadvertently created conduits in the cap rock may result in a gas leak from primary storage. Even in supercritical state, the carbon dioxide viscosity and density are lower than those of the indigenous formation brine. Therefore, buoyancy will tend to drive the CO{sub 2} upward unless it is trapped beneath a low permeability seal. Theoretical and experimental studies of buoyancy-driven supercritical CO{sub 2} flow, including estimation of time scales associated with plume evolution, are critical for developing technology, monitoring policy, and regulations for carbon dioxide geologic sequestration protecting the sources of potable water.

Silin, Dmitriy; Patzek, Tad; Benson, Sally M.

2007-08-20

428

A dual-cooled hydrogen-oxygen rocket engine heat transfer analysis  

NASA Technical Reports Server (NTRS)

The potential benefits of simultaneously using hydrogen and oxygen as rocket engine coolants are described. A plug-and-spool rocket engine was examined at heat fluxes ranging from 9290 to 163,500 kW/sq m, using a combined 3-D conduction/advection analysis. Both counter flow and parallel flow cooling arrangements were analyzed. The results indicate that a significant amount of heat transfer to the oxygen occurs, reducing both the hot side wall temperature of the rocket engine and also reducing the exit temperature of the hydrogen coolant. In all heat flux and coolant flow rates examined, the total amount of heat transferred to the oxygen was found to be largely independent of the oxygen coolant flow direction. At low heat flux/low coolant flow (throttled) conditions, the oxygen coolant absorbed more than 30 percent of the overall heat transfer from the rocket engine exhaust gasses. Also, hot side wall temperatures were judged to decrease by approximately 120 K in the throat area and up to a 170 K combustion chamber wall temperature reduction is expected if dual cooling is applied. The reduction in combustion chamber wall temperatures at throttled conditions is especially desirable since tha analysis indicates that a double temperature maxima, one at the throat and another in the combustion chamber, occurs with a traditional hydrogen cooled only engine. Conversely, a dual cooled engine essentially eliminates any concern for overheating in the combustion chamber.

Kacynski, Kenneth J.; Kazaroff, John M.; Jankovsky, Robert S.

1991-01-01

429

Plumes, plateaux and congestion in subduction zones  

NASA Astrophysics Data System (ADS)

The geologic record provides numerous examples where buoyant plumes, and their associated plateaux, have disrupted convergent plate margins. These interactions have produced a variety of responses in the overriding plate including transient episodes of arc magmatism, transient episodes of crustal shortening followed by plume-related magmatism in the overriding plate. The latter observation implies the plume must have transitioned from the subducting plate to the overriding plate. We present several 3D numerical models of plume heads of variable dimension and buoyancy interacting with a subducting slab. The models indicate that plume heads impact enormously on trench geometry. Arcuate trenches are created as the trench retreats around the edges of the plume head, whereas trench advance occurs in front of the plume resulting in transient crustal shortening in the overriding plate. Stalling of subduction when the plume head impacts the trench causes slab windowing. The size of the slab window is dependent on the size and buoyancy of the plume. The creation of the slab window provides a potential conduit for plume migration to the overriding plate. Alternatively, the plume head and plateau may be transferred to the overriding plate as subduction is re-established behind the plume. Models with "strong" slabs, characterized by high yield strengths, display different behavior. Plume-heads are entrained in the slab and are subducted without the development of a slab window. We discuss geological evidence for the processes observed in our models.

Moresi, Louis; Betts, Peter; Miller, Meghan; Willis, David

2014-05-01

430

Wind tunnel investigation of the downwash effect of a rooftop structure on plume dispersion  

NASA Astrophysics Data System (ADS)

This paper investigates the downwash effect of a rooftop structure (RTS) representing a typical RTS on plume dispersion. The effect of wind direction, exhaust speed, stack location, stack height, and RTS crosswind width on the severity of the downwash effect on the plume is assessed. Wind tunnel experiments were conducted to obtain plume centerline concentrations on the roof of typical low-rise and high-rise buildings. Measurements were obtained downwind of an RTS with height h = 4 m, along-wind length l = 8 m for 3 crosswind widths w = 10 m, 20 m and 30 m. Flow visualization was also conducted to obtain a qualitative assessment of the flow downwind of the RTS. The downwash produced by the RTS caused a significant increase in roof level concentration depending on building height, stack location, stack height, exhaust speed, wind direction and RTS crosswind width. An attempt is made to provide design guidance for determining stack height required to avoid the downwash effect for an exhaust placed downwind of the RTS.

Gupta, Amit; Stathopoulos, Ted; Saathoff, Patrick

2012-01-01

431

Hybrid rocket engine, theoretical model and experiment  

Microsoft Academic Search

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

Teodor-Viorel Chelaru; Florin Mingireanu

2011-01-01

432

Behavior of Mercury Emissions from a Commercial Coal-Fired Power Plant: The Relationship between Stack Speciation and Near-Field Plume Measurements.  

PubMed

The reduction of divalent gaseous mercury (Hg(II)) to elemental gaseous mercury (Hg(0)) in a commercial coal-fired power plant (CFPP) exhaust plume was investigated by simultaneous measurement in-stack and in-plume as part of a collaborative study among the U.S. EPA, EPRI, EERC, and Southern Company. In-stack continuous emission monitoring data were used to establish the CFPP's real-time mercury speciation and plume dilution tracer species (SO2, NOX) emission rates, and an airship was utilized as an airborne sampling platform to maintain static position with respect to the exhaust plume centerline for semicontinuous measurement of target species. Varying levels of Hg(II) concentration (2.39-3.90 ?g m(-3)) and percent abundance (?87-99%) in flue gas and in-plume reduction were observed. The existence and magnitude of Hg(II) reduction to Hg(0) (0-55%) observed varied with respect to the types and relative amounts of coals combusted, suggesting that exhaust plume reduction occurring downwind of the CFPP is influenced by coal chemical composition and characteristics. PMID:25325168

Landis, Matthew S; Ryan, Jeffrey V; Ter Schure, Arnout F H; Laudal, Dennis

2014-11-18

433

Acoustic Measurements for Small Solid Rocket Motors  

NASA Technical Reports Server (NTRS)

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

Vargas, Magda B.; Kenny, R. Jeremy

2010-01-01

434

Acoustic Measurements of Small Solid Rocket Motor  

NASA Technical Reports Server (NTRS)

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

Vargas, Magda B.; Kenny, R. Jeremy

2010-01-01

435

Subsonic Glideback Rocket Demonstrator Flight Testing  

NASA Technical Reports Server (NTRS)

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.

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

2001-01-01

436

A Flight Demonstration of Plasma Rocket Propulsion  

NASA Technical Reports Server (NTRS)

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.

Petro, Andrew

1999-01-01

437

Opacity meter for monitoring exhaust emissions from non-stationary sources  

DOEpatents

Method and apparatus for determining the opacity of exhaust plumes from moving emissions sources. In operation, a light source is activated at a time prior to the arrival of a diesel locomotive at a measurement point, by means of a track trigger switch or the Automatic Equipment Identification system, such that the opacity measurement is synchronized with the passage of an exhaust plume past the measurement point. A beam of light from the light source passes through the exhaust plume of the locomotive and is detected by a suitable detector, preferably a high-rate photodiode. The light beam is well-collimated and is preferably monochromatic, permitting the use of a narrowband pass filter to discriminate against background light. In order to span a double railroad track and provide a beam which is substantially stronger than background, the light source, preferably a diode laser, must provide a locally intense beam. A high intensity light source is also desirable in order to increase accuracy at the high sampling rates required. Also included is a computer control system useful for data acquisition, manipulation, storage and transmission of opacity data and the identification of the associated diesel engine to a central data collection center.

Dec, John Edward (Livermore, CA)

2000-01-01

438

Turbine engine exhaust gas measurements using in-situ FT-IR emission/transmission spectroscopy  

NASA Astrophysics Data System (ADS)

12 An advanced multiple gas analyzer based on in-situ Fourier transform infrared spectroscopy has been used to successfully measure the exhaust plume composition and temperature of an operating gas turbine engine at a jet engine test stand. The sensor, which was optically coupled to the test cell using novel broadband hollow glass waveguides, performed well in this harsh environment (high acoustical noise and vibration, considerable temperature swings in the ambient with engine operation), providing quantitative gas phase information. Measurements were made through the diameter of the engine's one meter exhaust plume, about 0.7 meters downstream of the engine exit plane. The sensor performed near simultaneous infrared transmission and infrared emission measurements through the centerline of the plume. Automated analysis of the emission and transmission spectra provided the temperature and concentration information needed for engine tuning and control that will ensure optimal engine operation and reduced emissions. As a demonstration of the utility and accuracy of the technique, carbon monoxide, nitric oxide, water, and carbon dioxide were quantified in spite of significant variations in the exhaust gas temperature. At some conditions, unburned fuel, particulates (soot/fuel droplets), methane, ethylene and aldehydes were identified, but not yet quantified.

Marran, David F.; Cosgrove, Joseph E.; Neira, Jorge; Markham, James R.; Rutka, Ronald; Strange, Richard R.

2001-02-01

439

Dynamics of laser ablated colliding plumes  

SciTech Connect

We report the dynamics of single and two collinearly colliding laser ablated plumes of ZnO studied using fast imaging and the spectroscopic measurements. Two dimensional imaging of expanding plume and temporal evolution of various species in interacting zones of plumes are used to calculate plume front velocity, electron temperature, and density of plasma. The two expanding plumes interact with each other at early stage of expansion ({approx}20 ns) resulting in an interaction zone that propagates further leading to the formation of stagnation layer at later times (>150 ns) at the lateral collision front of two plumes. Colliding plumes have larger concentration of higher ionic species, higher temperature, and increased electron density in the stagnation region. A one-to-one correlation between the imaging and optical emission spectroscopic observations in interaction zone of the colliding plumes is reported.

Gupta, Shyam L.; Pandey, Pramod K.; Thareja, Raj K. [Department of Physics, Indian Institute of Technology, Kanpur-208016 (India)

2013-01-15

440

Colorado Hydrogen Imaging Rocket Payload  

NASA Astrophysics Data System (ADS)

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

Burgh, Eric B.; France, K.

2009-01-01