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Sample records for rocket exhaust plumes

  1. Infrared Imagery of Solid Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Moran, Robert P.; Houston, Janice D.

    2011-01-01

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

  2. Implementation of microwave transmissions for rocket exhaust plume diagnostics

    NASA Astrophysics Data System (ADS)

    Coutu, Nicholas George

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

  3. Bipropellant rocket exhaust plume analysis on the Galileo spacecraft

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

    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.

  4. Environmental Effects of Space Shuttle Solid Rocket Motor Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Hwang, B.; Pergament, H. S.

    1976-01-01

    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.

  5. Zone radiometer measurements on a model rocket exhaust plume

    NASA Technical Reports Server (NTRS)

    1972-01-01

    Radiometer for analytical prediction of rocket plume-to-booster thermal radiation and convective heating is described. Applications for engine combustion analysis, incineration, and pollution control by high temperature processing are discussed. Illustrations of equipment are included.

  6. Impact of rocket exhaust plumes on atmospheric composition and climate ― an overview

    NASA Astrophysics Data System (ADS)

    Voigt, Ch.; Schumann, U.; Graf, K.; Gottschaldt, K.-D.

    2013-03-01

    Rockets are the only direct anthropogenic emission sources into the upper atmosphere. Gaseous rocket emissions include CO, N2, H2, H2O, and CO2, while solid rocket motors (SRM) additionally inject significant amounts of aluminum oxide (Al2O3) particles and gaseous chlorine species into the atmosphere. These emissions strongly perturb local atmospheric trace gas and aerosol distributions. Here, previous aircraft measurements in various rocket exhaust plumes including several large space shuttle launch vehicles are compiled. The observed changes of the lower stratospheric composition in the near field are summarized. The injection of chlorine species and particles into the stratosphere can lead to ozone loss in rocket exhaust plumes. Local observations are compared with global model simulations of the effects of rocket emissions on stratospheric ozone concentrations. Large uncertainties remain concerning individual ozone loss reaction rates and the impact of small-scale plume effects on global chemistry. Further, remote sensing data from satellite indicate that rocket exhaust plumes regionally increase iron and water vapor concentrations in the mesosphere potentially leading to the formation of mesospheric clouds at 80- to 90-kilometer altitude. These satellite observations are summarized and the rocket emission inventory is compared with other natural and anthropogenic sources to the stratosphere such as volcanism, meteoritic material, and aviation.

  7. Hydrazine engine plume contamination mapping. [measuring instruments for rocket exhaust from liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Chirivella, J. E.

    1975-01-01

    Instrumentation for the measurement of plume exhaust specie deposition rates were developed and demonstrated. The instruments, two sets of quartz crystal microbalances, were designed for low temperature operation in the back flow and variable temperature operation in the core flow regions of an exhaust plume. These quartz crystal microbalances performed nominally, and measurements of exhaust specie deposition rates for 8400 number of pulses for a 0.1-lb monopropellant thruster are reported.

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

    NASA Technical Reports Server (NTRS)

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

    2007-01-01

    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

  9. Species separation in rocket exhaust plumes and analytic plume flow models

    NASA Astrophysics Data System (ADS)

    Koppenwallner, G.

    2001-08-01

    Species separation in the exhaust plume of control thrusters of satellites is of main importance for the contamination analysis. Contamination concerns mainly scientific instruments or sensitive surfaces.. In continuum fluid dynamics a multi- component gas mixture can be treated as mixture with mean properties and with a flow field independent composition. This basic feature of continuum flow ceases to be valid in the rarefied flow regimes. In this regime there are two main mechanism which cause a separation of species in the flow field. a. Strong velocity gradients or streamline curvature. Strong stream line curvatures with large centrifugal forces exist close to the nozzle throat of sonic free jets [Sherman] or at the nozzle lip. Heavy gas constituents will not be able to follow these strong stream line curvatures. b. Different thermal velocity or thermal diffusivity of heavy and light gas constituents The transition from continuum to free molecular plume expansion can approximately be described by the sudden freeze model of Bird. At the freezing point molecular collisions suddenly cease and the further expansion is given by the velocity vector of the individual molecules at this freezing point. As light molecules have a larger thermal speed c than the heavy ones their spreading potential is also higher. This mechanism will also produce an enrichment of the plume boundary with light molecules. The approaches to model species separation in exhaust plumes as result of the above mechanism will be reviewed. To gain more insight into the separation the following cases are analyzed in detail: [B ]The free molecular supersonic expansion from a freezing plane. □ The various analytic plume flow models and their capability to predict the lateral spreading at the plume boundary (e.g. Simmons, Boynton, Brook, DLR) □ DSMC test case calculations of single and two-species plumes with mass separation. (M. Ivanov, ITAM) Based on this analysis a new 3 region model for species

  10. Stratospheric plume dispersion: Measurements from STS and Titan solid rocket motor exhaust. Technical report

    SciTech Connect

    Beiting, E.J.

    1999-04-20

    Plume expansion was measured from nine Space Shuttle and Titan IV vehicles at altitudes of 18, 24, and 30 km in the stratosphere. The plume diameters were inferred from electronic images of polarized, near-infrared solar radiation scattered from the exhaust particles, and these diameters were found to increase linearly with time. The expansion rate was measured for as long as 50 min after the vehicle reached altitude. Measurements made simultaneously at multiple altitudes showed that the expansion rate increased with increasing altitude for six measurements made at Cape Canaveral but decreased between 24 and 30 km for the one measurement made at Vandenberg AFB. The average expansion rates for all measurements are 4.3 {+-} 1.0 m/s at 18 km, 6.8 {+-} 1.9 m/s at 24 km, and 8.7 {+-} 2.5 m/s at 30 km. Expansion rates varied from launch to launch by as much as a factor of 1.6 at 18 km, 2.2 at 24 km, and 2.7 at 30 km. No correlation between the expansion rate and wind speed or shear was evident. These data are compared to several models for diffusivity and are used to update a comprehensive particle model of solid rocket motor exhaust in the stratosphere. The expansion rates are required by models to calculate the spatial extent and temporal persistence of the local stratospheric ozone depletion cause by solid rocket exhaust.

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

    SciTech Connect

    Morris, N.; Buttsworth, D.; Jones, T.; Brescianini, C. |

    1995-09-01

    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.

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

    NASA Technical Reports Server (NTRS)

    Barkhoudarian, Sarkis; Kittinger, Scott

    2006-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    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.

  14. Probe samples components of rocket engine exhaust

    NASA Technical Reports Server (NTRS)

    Schumacher, P. E.

    1965-01-01

    Water-cooled, cantilevered probe samples the exhaust plume of rocket engines to recover particles for examination. The probe withstands the stresses of a rocket exhaust plume environment for a sufficient period to obtain a useful sample of the exhaust components.

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

    NASA Technical Reports Server (NTRS)

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

    2008-01-01

    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.

  16. Using Lunar Module Shadows To Scale the Effects of Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    2008-01-01

    Excavating granular materials beneath a vertical jet of gas involves several physical mechanisms. These occur, for example, beneath the exhaust plume of a rocket landing on the soil of the Moon or Mars. We performed a series of experiments and simulations (Figure 1) to provide a detailed view of the complex gas-soil interactions. Measurements taken from the Apollo lunar landing videos (Figure 2) and from photographs of the resulting terrain helped demonstrate how the interactions extrapolate into the lunar environment. It is important to understand these processes at a fundamental level to support the ongoing design of higher fidelity numerical simulations and larger-scale experiments. These are needed to enable future lunar exploration wherein multiple hardware assets will be placed on the Moon within short distances of one another. The high-velocity spray of soil from the landing spacecraft must be accurately predicted and controlled or it could erode the surfaces of nearby hardware. This analysis indicated that the lunar dust is ejected at an angle of less than 3 degrees above the surface, the results of which can be mitigated by a modest berm of lunar soil. These results assume that future lunar landers will use a single engine. The analysis would need to be adjusted for a multiengine lander. Figure 3 is a detailed schematic of the Lunar Module camera calibration math model. In this chart, formulas relating the known quantities, such as sun angle and Lunar Module dimensions, to the unknown quantities are depicted. The camera angle PSI is determined by measurement of the imaged aspect ratio of a crater, where the crater is assumed to be circular. The final solution is the determination of the camera calibration factor, alpha. Figure 4 is a detailed schematic of the dust angle math model, which again relates known to unknown parameters. The known parameters now include the camera calibration factor and Lunar Module dimensions. The final computation is the ejected

  17. Numerically Modeling the Erosion of Lunar Soil by Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    2008-01-01

    In preparation for the Apollo program, Leonard Roberts of the NASA Langley Research Center 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 was 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 consumes 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. Roberts calculated the largest and smallest particles that could be eroded based on forces at the particle scale, but the erosion rate equation assumed that only one particle size existed in the soil. He assumed that particle ejection angles were determined entirely by the shape of the terrain, which acts like a ballistic ramp, with the particle aerodynamics being negligible. The predicted erosion rate and the upper limit of particle size 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 limit of particle size and the predictions of ejection angle were not tested. We observed in the Apollo landing videos that the ejection angles of particles streaming out from individual craters were time-varying and correlated to the Lunar Module thrust, thus implying that particle aerodynamics dominate. We modified Roberts theory in two ways. First, we used ad hoc the ejection angles measured in the Apollo landing videos, in lieu of developing a more sophisticated method. Second, we integrated Roberts equations over the lunar-particle size distribution and obtained a compact expression that could be implemented in a numerical code. We also added a material damage model that predicts the number and size of divots which the impinging particles will cause in hardware surrounding the landing

  18. Rocket plume burn hazard.

    PubMed

    Stoll, A M; Piergallini, J R; Chianta, M A

    1980-05-01

    By use of miniature rocket engines, the burn hazard posed by exposure to ejection seat rocket plume flames was determined in the anaesthetized rat. A reference chart is provided for predicting equivalent effects in human skin based on extrapolation of earlier direct measurements of heat input for rat and human burns. The chart is intended to be used in conjunction with thermocouple temperature measurements of the plume environment for design and modification of escape seat system to avoid thermal injury on ejection from multiplace aircraft. PMID:7387571

  19. Two-dimensional calculation of chemical species and electro-magnetic properties in rocket exhaust plume flow fields

    NASA Astrophysics Data System (ADS)

    Zhang, Ping; Cui, Jisong; Liu, Qingyun

    1993-08-01

    A computational modeling technique and prediction method were presented. Additionally, a comprehensive computer code was programmed. The chemical reactions and radar attenuation that occur in rocket plumes can be predicted precisely by using this code. It is suitable to calculating the parameters of rocket plumes under a near complete-expansion condition using a smokeless (or smoke reduced) propellant. The calculation results also indicate that serious errors will occur in the prediction of chemical and electrical properties in the plume flow field if the chemical reactions are not taken into account.

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

    NASA Technical Reports Server (NTRS)

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

    2008-01-01

    Roberts' model of lunar soil erosion beneath a landing rocket has been updated in several ways to predict the effects of future lunar landings. The model predicts, among other things, the number of divots that would result on surrounding hardware due to the impact of high velocity particulates, the amount and depth of surface material removed, the volume of ejected soil, its velocity, and the distance the particles travel on the Moon. The results are compared against measured results from the Apollo program and predictions are made for mitigating the spray around a future lunar outpost.

  1. Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60k-lb Thrust Fastrac Rocket Engine

    NASA Technical Reports Server (NTRS)

    Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

    2000-01-01

    A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location about equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal to 0.7 micrograms/cubic cm and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal to 2.200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

  2. Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60K-lb Thrust Fastrac Rocket Engine

    NASA Technical Reports Server (NTRS)

    Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

    2000-01-01

    A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location approximately equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal 0.7 microgram/cc, and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal 2,200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

  3. Transmittance and Radiance Computations for Rocket Engine Plume Environments

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.

    2003-01-01

    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.

  4. Empirical Scaling Laws of Rocket Exhaust Cratering

    NASA Technical Reports Server (NTRS)

    Donahue, Carly M.; Metzger, Philip T.; Immer, Christopher D.

    2005-01-01

    When launching or landing a space craft on the regolith of a terrestrial surface, special attention needs to be paid to the rocket exhaust cratering effects. If the effects are not controlled, the rocket cratering could damage the spacecraft or other surrounding hardware. The cratering effects of a rocket landing on a planet's surface are not understood well, especially for the lunar case with the plume expanding in vacuum. As a result, the blast effects cannot be estimated sufficiently using analytical theories. It is necessary to develop physics-based simulation tools in order to calculate mission-essential parameters. In this work we test out the scaling laws of the physics in regard to growth rate of the crater depth. This will provide the physical insight necessary to begin the physics-based modeling.

  5. Lidar measurements of launch vehicle exhaust plumes

    NASA Astrophysics Data System (ADS)

    Dao, Phan D.; Curtis, David; Farley, Robert; Soletsky, Philip; Davidson, Gilbert; Gelbwachs, Jerry A.

    1997-10-01

    The Mobile Lidar Trailer (MLT) was developed and operated to characterize launch vehicle exhaust plume and its effects on the environment. Two recent applications of this facility are discussed in this paper. In the first application, the MLT was used to characterize plumes in the stratosphere up to 45 km in support of the Air Force Space and Missile Center's Rocket Impact on Stratospheric Ozone program. Solid rocket motors used by Titan IV and other heavy launch vehicles release large quantities of gaseous hydrochloric acid in the exhaust and cause concerns about a possible depletion of the ozone layer. The MLT was deployed to Cape Canaveral Air Station since October 1995 to monitor ozone and to investigate plume dynamics and properties. Six campaigns have been conducted and more are planned to provide unique data with the objective of addressing the environmental issues. The plume was observed to disperse rapidly into horizontally extended yet surprisingly thin layer with thickness recorded in over 700 lidar profiles to be less than 250 meters. MLT operates with the laser wavelengths of 532, 355 and 308 nm and a scanning receiving telescope. Data on particle backscattering at the three wavelengths suggest a consistent growth of particle size in the 2-3 hour observation sessions following the launch. In the second type of application, the MLT was used as a remote sensor of nitrogen dioxide, a caustic gaseous by-product of common liquid propellant oxidizer. Two campaigns were conducted at the Sol Se Mete Canyon test site in New Mexico in December 1996 an January 1997 to study the dispersion of nitrogen dioxide and rocket plume.

  6. Ship exhaust gas plume cooling

    NASA Astrophysics Data System (ADS)

    Schleijpen, H. M. A.; Neele, Filip P.

    2004-08-01

    The exhaust gas plume is an important and sometimes dominating contributor to the infrared signature of ships. Suppression of the infrared ship signatures has been studied by TNO for the Royal Netherlands Navy over considerable time. This study deals with the suppression effects, which can be achieved using a spray of cold water in the inner parts of the exhaust system. The effects are compared with the effect of cooling with air. A typical frigate size diesel engine serves as an example for gas flow, composition and temperature of the plume. The infrared emission of the cooled an un-cooled exhaust gases is calculated. Both the spectral behaviour and the integrated values over typical bands are discussed. Apart from the signature also some advantages of water exhaust gas cooling for the ship design are discussed.

  7. Two-dimensional calculation of chemical species and electrical properties in rocket plume flowfields

    NASA Astrophysics Data System (ADS)

    Zhang, Ping; Cui, Jisong; Liu, Qingyun

    1993-08-01

    A computational modeling technique and prediction method are presented for calculating two-dimensional profiles of chemical species mole fraction and electrical properties of rocket exhaust plumes. A comprehensive computer code has been programmed. The chemical reactions and radar attenuation which occur in a rocket plume can be predicted more truly by using this code. It is suitable to calculating parameters of rocket plumes under a near complete-expansion condition and for smokeless (or reduced smoke) propellant application. The calculation results indicate that evident errors will occur for prediction of chemical and electrical parameters in the plume flowfield if the chemical reactions in the plume are ignored.

  8. Effect of Soot Particles on Supersonic Rocket Plume Properties

    NASA Astrophysics Data System (ADS)

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

    2012-06-01

    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.

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

    NASA Astrophysics Data System (ADS)

    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

    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.

  10. Exhaust Nozzle Plume and Shock Wave Interaction

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

    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.

  11. Radiation from advanced solid rocket motor plumes

    NASA Technical Reports Server (NTRS)

    Farmer, Richard C.; Smith, Sheldon D.; Myruski, Brian L.

    1994-01-01

    The overall objective of this study was to develop an understanding of solid rocket motor (SRM) plumes in sufficient detail to accurately explain the majority of plume radiation test data. Improved flowfield and radiation analysis codes were developed to accurately and efficiently account for all the factors which effect radiation heating from rocket plumes. These codes were verified by comparing predicted plume behavior with measured NASA/MSFC ASRM test data. Upon conducting a thorough review of the current state-of-the-art of SRM plume flowfield and radiation prediction methodology and the pertinent data base, the following analyses were developed for future design use. The NOZZRAD code was developed for preliminary base heating design and Al2O3 particle optical property data evaluation using a generalized two-flux solution to the radiative transfer equation. The IDARAD code was developed for rapid evaluation of plume radiation effects using the spherical harmonics method of differential approximation to the radiative transfer equation. The FDNS CFD code with fully coupled Euler-Lagrange particle tracking was validated by comparison to predictions made with the industry standard RAMP code for SRM nozzle flowfield analysis. The FDNS code provides the ability to analyze not only rocket nozzle flow, but also axisymmetric and three-dimensional plume flowfields with state-of-the-art CFD methodology. Procedures for conducting meaningful thermo-vision camera studies were developed.

  12. Atmospheric scavenging of solid rocket exhaust effluents

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

    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.

  13. Numerical Analysis of Rocket Exhaust Cratering

    NASA Technical Reports Server (NTRS)

    2008-01-01

    Supersonic jet exhaust impinging onto a flat surface is a fundamental flow encountered in space or with a missile launch vehicle system. The flow is important because it can endanger launch operations. The purpose of this study is to evaluate the effect of a landing rocket s exhaust on soils. From numerical simulations and analysis, we developed characteristic expressions and curves, which we can use, along with rocket nozzle performance, to predict cratering effects during a soft-soil landing. We conducted a series of multiphase flow simulations with two phases: exhaust gas and sand particles. The main objective of the simulation was to obtain the numerical results as close to the experimental results as possible. After several simulating test runs, the results showed that packing limit and the angle of internal friction are the two critical and dominant factors in the simulations.

  14. Range safety signal propagation through the SRM exhaust plume of the space shuttle

    NASA Technical Reports Server (NTRS)

    Boynton, F. P.; Davies, A. R.; Rajasekhar, P. S.; Thompson, J. A.

    1977-01-01

    Theoretical predictions of plume interference for the space shuttle range safety system by solid rocket booster exhaust plumes are reported. The signal propagation was calculated using a split operator technique based upon the Fresnel-Kirchoff integral, using fast Fourier transforms to evaluate the convolution and treating the plume as a series of absorbing and phase-changing screens. Talanov's lens transformation was applied to reduce aliasing problems caused by ray divergence.

  15. Effects of entrained water and strong turbulence on afterburning within solid rocket motor plumes

    NASA Technical Reports Server (NTRS)

    Gomberg, R. I.; Wilmoth, R. G.

    1978-01-01

    During the first few seconds of the space shuttle trajectory, the solid rocket boosters will be in the proximity of the launch pad. Because of the launch pad structures and the surface of the earth, the turbulent mixing experienced by the exhaust gases will be greatly increased over that for the free flight situation. In addition, a system will be present, designed to protect the lifting vehicle from launch structure vibrations, which will inject quantities of liquid water into the hot plume. The effects of these two phenomena on the temperatures, chemical composition, and flow field present in the afterburning solid rocket motor exhaust plumes of the space shuttle were studied. Results are included from both a computational model of the afterburning and supporting measurements from Titan 3 exhaust plumes taken at Kennedy Space Center with infrared scanned radiometers.

  16. Particle Size Distributions Measured in the Stratospheric Plumes of Three Rockets During the ACCENT Missions

    NASA Astrophysics Data System (ADS)

    Wiedinmyer, C.; Brock, C. A.; Reeves, J. M.; Ross, M. N.; Schmid, O.; Toohey, D.; Wilson, J. C.

    2001-12-01

    The global impact of particles emitted by rocket engines on stratospheric ozone is not well understood, mainly due to the lack of comprehensive in situ measurements of the size distributions of these emitted particles. During the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) missions in 1999, the NASA WB-57F aircraft carried the University of Denver N-MASS and FCAS instruments into the stratospheric plumes from three rockets. Size distributions of particles with diameters from 4 to approximately 2000 nm were calculated from the instrument measurements using numerical inversion techniques. The data have been averaged over 30-second intervals. The particle size distributions observed in all of the rocket plumes included a dominant mode near 60 nm diameter, probably composed of alumina particles. A smaller mode at approximately 25 nm, possibly composed of soot particles, was seen in only the plumes of rockets that used liquid oxygen and kerosene as a propellant. Aircraft exhaust emitted by the WB-57F was also sampled; the size distributions within these plumes are consistent with prior measurements in aircraft plumes. The size distributions for all rocket intercepts have been fitted to bimodal, lognormal distributions to provide input for global models of the stratosphere. Our data suggest that previous estimates of the solid rocket motor alumina size distributions may underestimate the alumina surface area emission index, and so underestimate the particle surface area available for heterogeneous chlorine activation reactions in the global stratosphere.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    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.

  18. Solar rocket plume-mirror interactions

    SciTech Connect

    Yu, S.T.; Chang, C.L.; Merkle, C.L.

    1991-01-01

    The extent to which the plume from a solar thermal rocket will impinge on the solar collector is studied by flow field analysis. Such interaction can adversely affect collector performance through fouling, excessive heat loading, or pressure loads that deform the delicate structures. The geometrical shape of the collector is such that only the flow from the nozzle boundary layer can reach it, but the thrust levels of interest lead to very viscous nozzle flows with thick boundary layers. Reasonable accuracy in solving these flows requires a fully coupled viscous-inviscid procedure. Results show that the fraction of the plume that hits the collector can be well estimated by continuum theory, but that transitional and rarefied phenomena will have some impact on how it is distributed over the surface. Initial results for one representative condition show that approx. 4 percent of the total flow in the jet makes its way to the collector. The pressures on the collector, however, remain quite low because of its distance from the engine. Additional work is needed to document the effect of thrust scaling and wall cooling on impingement.

  19. Stratospheric aircraft exhaust plume and wake chemistry

    NASA Technical Reports Server (NTRS)

    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.

    1993-01-01

    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 is documented. A kinetic condensation model was implemented to predict heterogeneous condensation in the plume regime behind an HSCT flying in the lower stratosphere. Simulations were performed to illustrate the parametric dependence of contrail droplet growth on the exhaust condensation nuclei number density and size distribution. Model results indicate that the condensation of water vapor is strongly dependent on the number density of activated CN. Incorporation of estimates for dilution factors into a Lagrangian box model of the far-wake regime with scale-dependent diffusion indicates negligible decrease in ozone and enhancement of water concentrations of 6-13 times background, which decrease rapidly over 1-3 days. Radiative calculations indicate a net differential cooling rate of the plume about 3K/day at the beginning of the wake regime, with a total subsidence ranging between 0.4 and 1 km. Results from the Lagrangian plume model were used to estimate the effect of repeated superposition of aircraft plumes on the concentrations of water and NO(y) along a flight corridor. Results of laboratory studies of heterogeneous chemistry are also described. Kinetics of HCl, N2O5 and ClONO2 uptake on liquid sulfuric acid were measured as a function of composition and temperature. Refined measurements of the thermodynamics of nitric acid hydrates indicate that metastable dihydrate may play a role in the nucleation of more stable trihydrates PSC's.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

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

    PubMed

    Hu, Jichao; Chang, Juntao; Bao, Wen

    2014-01-01

    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.

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

    PubMed Central

    2014-01-01

    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

  4. Rocket exhaust ground cloud/atmospheric interactions

    NASA Technical Reports Server (NTRS)

    Hwang, B.; Gould, R. K.

    1978-01-01

    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.

  5. Rocket Plume Scaling for Orion Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

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

    2011-01-01

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

  6. Applicability of empirical data currently used in predicting solid propellant exhaust plumes

    NASA Technical Reports Server (NTRS)

    Tevepaugh, J. A.; Smith, S. D.; Penny, M. M.; Greenwood, T.; Roberts, B. B.

    1977-01-01

    Theoretical and experimental approaches to exhaust plume analysis are compared. A two-phase model is extended to include treatment of reacting gas chemistry, and thermodynamical modeling of the gaseous phase of the flow field is considered. The applicability of empirical data currently available to define particle drag coefficients, heat transfer coefficients, mean particle size, and particle size distributions is investigated. Experimental and analytical comparisons are presented for subscale solid rocket motors operating at three altitudes with attention to pitot total pressure and stagnation point heating rate measurements. The mathematical treatment input requirements are explained. The two-phase flow field solution adequately predicts gasdynamic properties in the inviscid portion of two-phase exhaust plumes. It is found that prediction of exhaust plume gas pressures requires an adequate model of flow field dynamics.

  7. Lander rocket exhaust effects on Europa regolith nitrogen assays

    NASA Astrophysics Data System (ADS)

    Lorenz, Ralph D.

    2016-08-01

    Soft-landings on large worlds such as Europa or our Moon require near-surface retropropulsion, which leads to impingement of the rocket plume on the surface. Surface modification by such plumes was documented on Apollo and Surveyor, and on Mars by Viking, Curiosity and especially Phoenix. The low temperatures of the Europan regolith may lead to efficient trapping of ammonia, a principal component of the exhaust from monopropellant hydrazine thrusters. Deposited ammonia may react with any trace organics, and may overwhelm the chemical and isotopic signatures of any endogenous nitrogen compounds, which are likely rare on Europa. An empirical correlation of the photometrically-altered regions ('blast zones') around prior lunar and Mars landings is made, indicating A=0.02T1.5, where A is the area in m2 and W is the lander weight (thus, ~thrust) at landing in N: this suggests surface alteration will occur out to a distance of ~9 m from a 200 kg lander on Europa.

  8. Surface composition of solid-rocket exhausted aluminum oxide particles

    NASA Technical Reports Server (NTRS)

    Cofer, Wesley R., III; Winstead, Edward L.; Key, Lawrence E.

    1989-01-01

    Particulate samples of aluminum oxide were collected on Teflon filters from the exhaust plume of the Space Shuttle (STS-61A, October 30, 1985) over the altitude interval 4.6-7.6 km immediately after launch. These particles were analyzed using SEM, energy-dispersive X-ray analysis, electron spectroscopy for chemical analysis, X-ray fluorescent spectroscopy, and conventional wet-chemical techniques. The samples were 0.6-1.0 percent surface-chlorided (chlorided meaning predominantly aluminum chlorides and oxychlorides, possibly including other adsorbed forms of chloride) by weight. This level of chloriding is about one-third of the amount determined previously from laboratory-prepared alumina and surface site samples of solid-rocket-produced alumina (SRPA) after both had been exposed to moist HCl vapor at temperatures down to ambient. This level is equivalent to previous laboratory results with samples exposed to moist HCl at temperatures above the boiling point of water. It is suggested that the present lower chloriding levels, determined for samples from a 'dry' Shuttle exhaust cloud, underscore the importance of a liquid water/hydrochloric acid phase in governing the extent of surface chloriding of SRPA. The reduced chloriding is not trivial with respect to potential physical/chemical modification of the SRPA particle surfaces and their corresponding interaction with the atmosphere.

  9. Combustion Model of Supersonic Rocket Exhausts in an Entrained Flow Enclosure

    NASA Technical Reports Server (NTRS)

    Vu, Bruce T.; Oliveira, Justin

    2011-01-01

    This paper describes the Computational Fluid Dynamics (CFD) model developed to simulate the supersonic rocket exhaust in an entrained flow cylinder. The model can be used to study the plume-induced environment due to static firing tests of the Taurus-II launch vehicle. The finite-rate chemistry is used to model the combustion process involving rocket propellant (RP-1) and liquid oxidizer (LOX). A similar chemical reacting model is also used to simulate the mixing of rocket plume and ambient air. The model provides detailed information on the gas concentration and other flow parameters within the enclosed region, thus allowing different operating scenarios to be examined in an efficient manner. It is shown that the real gas influence is significant and yields better agreement with the theory.

  10. Combustion Model of Supersonic Rocket Exhausts in an Entrained Flow Enclosure

    NASA Technical Reports Server (NTRS)

    Vu, Bruce; Oliveira, Justin

    2011-01-01

    This paper describes the Computation Fluid Dynamics (CFD) model developed to simulate the supersonic rocket exhaust in an entrained flow cylinder. The model can be used to study the plume-induced environment due to static firing test of the Taurus II launch vehicle. The finite rate chemistry is used to model the combustion process involving rocket propellant (RP 1) and liquid oxidizer (LOX). A similar chemical reacting model is also used to simulate the mixing of rocket plume and ambient air. The model provides detailed information on the gas concentration and other flow parameters within the enclosed region thus allowing different operating scenarios to be examined in an efficient manner. It is shown that the real gas influence is significant and yields better agreement with the theory.

  11. Rocket-Plume Spectroscopy Simulation for Hydrocarbon-Fueled Rocket Engines

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.

    2010-01-01

    The UV-Vis spectroscopic system for plume diagnostics monitors rocket engine health by using several analytical tools developed at Stennis Space Center (SSC), including the rocket plume spectroscopy simulation code (RPSSC), to identify and quantify the alloys from the metallic elements observed in engine plumes. Because the hydrocarbon-fueled rocket engine is likely to contain C2, CO, CH, CN, and NO in addition to OH and H2O, the relevant electronic bands of these molecules in the spectral range of 300 to 850 nm in the RPSSC have been included. SSC incorporated several enhancements and modifications to the original line-by-line spectral simulation computer program implemented for plume spectral data analysis and quantification in 1994. These changes made the program applicable to the Space Shuttle Main Engine (SSME) and the Diagnostic Testbed Facility Thruster (DTFT) exhaust plume spectral data. Modifications included updating the molecular and spectral parameters for OH, adding spectral parameter input files optimized for the 10 elements of interest in the spectral range from 320 to 430 nm and linking the output to graphing and analysis packages. Additionally, the ability to handle the non-uniform wavelength interval at which the spectral computations are made was added. This allowed a precise superposition of wavelengths at which the spectral measurements have been made with the wavelengths at which the spectral computations are done by using the line-by-line (LBL) code. To account for hydrocarbon combustion products in the plume, which might interfere with detection and quantification of metallic elements in the spectral region of 300 to 850 nm, the spectroscopic code has been enhanced to include the carbon-based combustion species of C2, CO, and CH. In addition, CN and NO have spectral bands in 300 to 850 nm and, while these molecules are not direct products of hydrocarbon-oxygen combustion systems, they can show up if nitrogen or a nitrogen compound is present

  12. Further Studies Using a Novel Free Molecule Rocket Plume Model

    NASA Astrophysics Data System (ADS)

    Woronowicz, Michael

    2003-05-01

    This paper describes some recent studies conducted using a set of analytic point source transient free molecule equations generated to model behavior ranging from molecular effusion to rocket plumes. These studies include comparisons to experimental data regarding steady flow from a sonic orifice and generation of a thruster backflow environment, followed by a transient development of plumes due to steady thruster operations and to a single pulse.

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

    NASA Astrophysics Data System (ADS)

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

    2013-09-01

    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.

  14. Predicting exhaust plume boundaries with supersonic external flows

    NASA Astrophysics Data System (ADS)

    Nash, Kyle L.; Whitaker, Kevin W.; Freeman, L. Michael

    1994-09-01

    Several methods for predicting exhaust plume boundaries with a surrounding external flow currently exist. Unfortunately, these methods are usually cumbersome and often expensive, since they may be computationally intensive. Also, these methods typically provide many flowfield details in addition to the plume boundary location. If only the latter is desired, then calculation of these other details is wasted effort. This concern resulted in the development of a simplified plume boundary prediction method capable of analyzing underexpanded nozzle flow exhausting into a supersonic external flow. This new method is based upon the well-established Latvala method and uses an iterative scheme that employs two-dimensional flowfield assumptions. However, the method is still applicable to axisymmetric plumes, and its simplicity permits efficient operation on personal computers. Predictions of boundaries for axisymmetric plumes surrounded by various high-speed external flows exhibit excellent agreement with empirical data, and parametric studies indicate that trends are correctly predicted.

  15. Improvement of Rocket Engine Plume Analysis Techniques

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1982-01-01

    A nozzle plume flow field code was developed. The RAMP code which was chosen as the basic code is of modular construction and has the following capabilities: two phase with two phase transonic solution; a two phase, reacting gas (chemical equilibrium reaction kinetics), supersonic inviscid nozzle/plume solution; and is operational for inviscid solutions at both high and low altitudes. The following capabilities were added to the code: a direct interface with JANNAF SPF code; shock capturing finite difference numerical operator; two phase, equilibrium/frozen, boundary layer analysis; a variable oxidizer to fuel ratio transonic solution; an improved two phase transonic solution; and a two phase real gas semiempirical nozzle boundary layer expansion.

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

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

    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.

  18. NTS-spill test facility wind tunnel exhaust plume characterization

    SciTech Connect

    Kerr, R.; Goldwire, H.; Smith, D.; Rawlings, J.; Schaffer, T.; Robson, J.

    1994-07-01

    The exhaust plume of the NTS-STF wind tunnel has been characterized to demonstrate its suitability as a target for CALIOPE experiments. Smoke from grenades has been released in multiple quantities and at different positions inside the tunnel. The smoke plumes have been recorded on video tape. The wind velocity profile has also been determined with a moveable array of miniature vane anemometers. These measurements will be used to determine the vapor concentration pathlength as part of the ground truth.

  19. Atmospheric scavenging of hydrochloric acid. [from rocket exhaust

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    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.

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

    PubMed

    Gelbwachs, J A

    1996-05-20

    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.

  1. Numerical Simulation of Rocket Exhaust Interaction with Lunar Soil

    NASA Technical Reports Server (NTRS)

    Liever, Peter; Tosh, Abhijit; Curtis, Jennifer

    2012-01-01

    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

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

    SciTech Connect

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

    2012-10-15

    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.

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

    NASA Technical Reports Server (NTRS)

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

    2009-01-01

    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.

  4. Ecological effects and environmental fate of solid rocket exhaust

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

    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.

  5. Ice nucleus activity measurements of solid rocket motor exhaust particles

    NASA Technical Reports Server (NTRS)

    Keller, V. W. (Compiler)

    1986-01-01

    The ice Nucleus activity of exhaust particles generated from combustion of Space Shuttle propellant in small rocket motors has been measured. The activity at -20 C was substantially lower than that of aerosols generated by unpressurized combustion of propellant samples in previous studies. The activity decays rapidly with time and is decreased further in the presence of moist air. These tests corroborate the low effectivity ice nucleus measurement results obtained in the exhaust ground cloud of the Space Shuttle. Such low ice nucleus activity implies that Space Shuttle induced inadvertent weather modification via an ice phase process is extremely unlikely.

  6. Assessment of analytical techniques for predicting solid propellant exhaust plumes

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    The calculation of solid propellant exhaust plume flow fields is addressed. Two major areas covered are: (1) the applicability of empirical data currently available to define particle drag coefficients, heat transfer coefficients, mean particle size and particle size distributions, and (2) thermochemical modeling of the gaseous phase of the flow field. Comparisons of experimentally measured and analytically predicted data are made. The experimental data were obtained for subscale solid propellant motors with aluminum loadings of 2, 10 and 15%. Analytical predictions were made using a fully coupled two-phase numerical solution. Data comparisons will be presented for radial distributions at plume axial stations of 5, 12, 16 and 20 diameters.

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

    PubMed

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

    2016-06-10

    This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air. PMID:27409023

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

    PubMed

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

    2016-06-10

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

  9. Studies of the exhaust products from solid propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Dawbarn, R.; Kinslow, M.

    1976-01-01

    This study was undertaken to determine the feasibility of conducting environmental chamber tests on the physical processes which occur when a solid rocket motor exhaust mixes with the ambient atmosphere. Of particular interest was the interaction between hydrogen chloride, aluminum oxide, and water vapor. The program consisted of three phases: (1) building a small rocket motor and using it to provide the exhaust species in a controlled environment; (2) evaluating instruments used to detect and measure HCl concentrations and if possible determining whether the HCl existed in the gaseous state or as an acid aerosol; (3) monitoring a series of 6.4-percent scale space shuttle motor tests and comparing the results to the environmental chamber studies. Eighteen firings were conducted in an environmental chamber with the initial ambient relative humidity set at values from 29 to 100 percent. Two additional firings were made in a large shed, and four were made on an open concrete apron. Six test firings at MSFC were monitored, and the ground level concentrations are reported. Evidence is presented which shows that the larger Al2O3 (5 to 50 micrometers) particles from the rocket motor can act as condensation nuclei. Under appropriate ambient conditions where there is sufficient water vapor this results in the formation of an acid aerosol. Droplets of this acid were detected both in the environmental chamber and in the scaled shuttle engine tests.

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

    NASA Astrophysics Data System (ADS)

    Chung, Chan H.; Kim, Suk C.; Stubbs, Robert M.; de Witt, Kenneth J.

    1993-06-01

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

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

    SciTech Connect

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

    1993-06-01

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

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

    NASA Technical Reports Server (NTRS)

    Chung, Chan H.; Kim, Suk C.; Stubbs, Robert M.; De Witt, Kenneth J.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Baran, W. J.

    1974-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    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.

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

    SciTech Connect

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

    1997-05-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

    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.

  17. Performance of reinforced polymer ablators exposed to a solid rocket motor exhaust. Technical report

    SciTech Connect

    Boyer, C.; Burgess, T.; Bowen, J.; Deloach, K.; Talmy, I.

    1992-10-01

    Summarized in this report is the effort by the Naval Surface Warfare Center Dahlgren Division (NSWCDD) and FMC Corporation (a launcher manufacturer) to identify new high performance ablators suitable for use on Navy guided missile launchers (GML) and ships' structures. The goal is to reduce ablator erosion by 25 to 50 percent compared to that of the existing ablators such as MXBE350 (rubbermodified phenolic containing glass fiber reinforcement). This reduction in erosion would significantly increase the number of new missiles with higher-thrust, longer burn rocket motors that can be launched prior to ablator refurbishment. In fact, there are a number of new Navy missiles being considered for development and introduction into existing GML: e.g., the Antisatellite Missile (ASM) and the Theater High-Altitude Area Defense (THAAD) Missile. The U.S. Navy experimentally evaluated the eight best fiber-reinforced, polymer composites from a possible field of 25 off-the-shelf ablators previously screened by FMC Corporation. They were tested by the Navy in highly aluminized solid rocket motor exhaust plumes to determine their ability to resist erosion and to insulate.... Ablator, Guided Missile Launchers, Erosion, Tactical missiles, Convective heating, Solid rocket motors, Aluminum oxide particles.

  18. Injection of Nuclear Rocket Engine Exhaust into Deep Unsaturated Zones

    NASA Astrophysics Data System (ADS)

    Cooper, C. A.; Decker, D.

    2008-05-01

    Nuclear rocket engine technology is being considered as a means of interplanetary vehicle propulsion for a manned mission to Mars. To achieve this, a test and development facility must be constructed to safely run nuclear engines. The testing of nuclear engines in the 1950's and 1960's was accomplished by exhausting the engine gases into the atmosphere, a practice that is no longer acceptable. Injection into deep unsaturated zones of radioactive exhaust gases and water vapor associated with the testing of nuclear rocket engines is being considered as a way of sequestering radionuclides from the environment. Numerical simulations were conducted to determine the ability of an unsaturated zone with the hydraulic properties of Frenchman Flat alluvium at the Nevada Test Site to contain gas-phase radionuclides. Gas and water vapor were injected for two hours at rates of 14.5 kg s-1 and 15 kg s-1, respectively, in an interval between 100 and 430 m below the land surface into alluvium with an intrinsic permeability of 10-11 m2 and porosity of 0.35. The results show that during a test of an engine, radionuclides with at least greater than 10-year half-lives may reach the land surface within several years after injection. Radionuclide transport is primarily controlled by the upward pressure gradient from the point of injection to the lower (atmospheric) pressure boundary condition at the land surface. Radionuclides with half-lives on the order of days should undergo enough decay prior to reaching the land surface. A cooling water vapor injected into the unsaturated zone simultaneously with the exhaust gas will condense within several meters of the injection point and drain downward toward the water table. However, the nearly horizontal hydraulic groundwater gradient present in several of the basins at NTS should limit lateral migration of radionuclides away from the vicinity of injection.

  19. Mass Spectra of Individual Aerosol Particles Acquired During Intercepts of a Space Shuttle Exhaust Plume

    NASA Astrophysics Data System (ADS)

    Cziczo, D. J.; Cziczo, D. J.; Murphy, D. M.; Thomson, D. S.; Thomson, D. S.

    2001-12-01

    The WB-57 aircraft accomplished fourteen distinct stratospheric intercepts of the exhaust plume from a space shuttle during ACCENT 2000. Liftoff of the shuttle Atlantis for STS-106 occurred at 8:46 am local (12:46 UTC) with intercepts occurring from 5 to 90 minutes afterward. The Particle Analysis by Laser Mass Spectrometry (PALMS) instrument, mounted in the nose of the aircraft, was used to acquire individual mass spectra of over 2500 particles during these intercepts. The majority of positive mass spectra indicate the presence of the metals Al, Fe, Zn, Ga, and V, all components found in the solid rocket fuel. Organic material, presumably from binding and curing agents, was also present. Negative mass spectra showed Cl from the oxidizer, ammonium perchlorate, as well as water. Rare exotic particles, for example those containing Ti and Ag and possibly formed during engine or seal ablation, were also detected. Particles originating from shuttle exhaust but also containing significant sulfuric acid were common toward the outer edge of the plume, especially during late encounters, suggesting that deposition or aerosol collision had occurred.

  20. Local and global effects on ozone from Titan rocket exhaust and deorbiting spacecraft debris

    SciTech Connect

    Connell, P.S.; Walton, J.J.; Penner, J.E.; O`Connor, C.

    1996-05-01

    Both the launching and deorbiting of spacecraft introduce foreign material directly into the stratosphere, a region of the atmosphere extending from around 12 to 50km above the earth`s surface. Launching of Titan and similar solid rocket motors adds to the stratospheric inorganic chlorine burden through emissions of HCl, atomic (Cl) and molecular (CL{sub 2}) directly into the stratosphere. Before the exhaust plume disperses, plume concentrations of these species are orders of magnitude above the background values (Denison et al, 1994). Dispersed through the stratosphere over the globe, however, the additional Cl burden is small compared to the background for currently envisioned launch frequencies. Inorganic chlorine is cleared from the atmosphere by wet deposition of HCl in rain after transport processes return air from the stratosphere to the troposphere, with an overall lifetime of a few years. After several year,a continuing fixed injection rate will produce a chlorine enhancement that reaches a steady state, balanced with loss via rainout. We report here on calculations in models in both two- and three dimensions that address three questions in rocket/spacecraft/stratospheric interactions. We have attempted to represent the early evolution (1-50 hours) of a vertical plume in the stratosphere with a Langrangian three- dimensional transport model driven by horizontal winds from a data- assimilating general circulation model. We have also conducted global calculations of the potential steady state effects of Cl injection from a specified rate of continuous launches in a current two- dimensional model of the stratosphere including all known important ozone production and loss processes. And, we have calculated the effect of increasing the steady state particulate surface area density in the stratosphere resulting from particle formation from satellite destruction on reentry.

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

    NASA Technical Reports Server (NTRS)

    Verma, Satyajit

    2006-01-01

    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.

  2. Active chlorine and nitric oxide formation from chemical rocket plume afterburning

    NASA Astrophysics Data System (ADS)

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

    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.

  3. Active chlorine and nitric oxide formation from chemical rocket plume afterburning

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

    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.

  4. Rocket Exhaust Cratering: Lessons Learned from Viking and Apollo

    NASA Technical Reports Server (NTRS)

    Metzger, Philip T.; Vu, Bruce T.

    2004-01-01

    During the Apollo and Viking programs NASA expended considerable effort to study the cratering of the regolith when a rocket launches or lands on it. That research ensured the success of those programs but also demonstrated that cratering will be a serious challenge for other mission scenarios. Unfortunately, because three decades have elapsed since NASA last performed a successful retro-rocket landing on a large planetary body - and ironically because Apollo and Viking were successful at minimizing the cratering effects - the space agency has a minimized sense of the seriousness of the issue. The most violent phase of a cratering event is when the static overpressure of the rocket exhaust exceeds the bearing capacity of the soil. This bearing capacity failure (BCF) punches a small and highly concave cup into the surface. The shape of the cup then redirects the supersonic jet - along with a large flux of high-velocity debris - directly toward the spacecraft. This has been observed in terrestrial experiments but never quantified analytically. The blast from such an event will be more than just quantitatively greater than the cratering that occurred in the Apollo and Viking programs. It will be qualitatively different, because BCF had been successfully avoided in all those missions. In fact, the Viking program undertook a significant research and development effort and redesigned the spacecraft specifically for the purpose of avoiding BCF [1]. (See Figure 1.) Because the Apollo and Viking spacecraft were successful at avoiding those cratering effects, it was unnecessary to understand them. As a result, the physics of a BCF-driven cratering event have never been well understood. This is a critical gap in our knowledge because BCF is unavoidable in the Martian environment with the large landers necessary for human exploration, and in Lunar landings it must also be addressed because it may occur depending upon the design specifics of the spacecraft and the weakening of

  5. Effects of rocket exhaust products in the thermosphere and ionsphere

    SciTech Connect

    Zinn, J.; Sutherland, C.D.

    1980-02-01

    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.

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

    NASA Astrophysics Data System (ADS)

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

    2016-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    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.

  8. Wedge Shock and Nozzle Exhaust Plume Interaction in a Supersonic Jet Flow

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Zaman, Khairul; Fagan, Amy; Heath, Christopher

    2014-01-01

    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 nozzle exhaust plume. Aft body shock waves that interact with the exhaust plume contribute to the near-field pressure signature of a vehicle. The plume and shock interaction was studied using computational fluid dynamics and compared with experimental data from a coaxial convergent-divergent nozzle flow in an open jet facility. A simple diamond-shaped wedge was used to generate the shock in the outer flow to study its impact on the inner jet flow. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the opposite plume boundary. The sonic boom pressure signature of the nozzle exhaust plume was modified by the presence of the wedge. Both the experimental results and computational predictions show changes in plume deflection.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    Film/dump cooling a rocket nozzle with fuel rich gas, as in the National Launch System (NLS) Space Transportation Main Engine (STME), adds potential complexities for integrating the engine with the vehicle. The chief concern is that once the film coolant is exhausted from the nozzle, conditions may exist during flight for the fuel-rich film gases to be recirculated to the vehicle base region. The result could be significantly higher base temperatures than would be expected from a regeneratively cooled nozzle. CFD analyses were conduced to augment classical scaling techniques for vehicle base environments. The FDNS code with finite rate chemistry was used to simulate a single, axisymmetric STME plume and the NLS base area. Parallel calculations were made of the Saturn V S-1 C/F1 plume base area flows. The objective was to characterize the plume/freestream shear layer for both vehicles as inputs for scaling the S-C/F1 flight data to NLS/STME conditions. The code was validated on high speed flows with relevant physics. This paper contains the calculations for the NLS/STME plume for the baseline nozzle and a modified nozzle. The modified nozzle was intended to reduce the fuel available for recirculation to the vehicle base region. Plumes for both nozzles were calculated at 10kFT and 50kFT.

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  11. 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)

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

    1976-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  13. Chance Encounter with a Stratospheric Kerosene Rocket Plume from Russia over California

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Wilson, J. C.; Ross, M. N.; Brock, C.; Sheridan, P.; Schoeberl, M. R.; Lait, L. R.; Bui, T. P.; Loewenstein, M.

    1999-01-01

    During a routine ER-2 aircraft high-altitude test flight on April 18, 1997, an unusual aerosol cloud was detected at 20 km altitude near the California coast at about 370 degrees N latitude. Not visually observed by the ER-2 pilot, the cloud was characterized bv high concentration of soot and sulfate aerosol in a region over 100 km in horizontal extent indicating that the source of the plume was a large hydrocarbon fueled vehicle, most likely a launch vehicle powered only by rocket motors burning liquid oxygen and kerosene. Two Russian Soyuz rockets could conceivably have produced the plume. The first was launched from the Baikonur Cosmodrome, Kazakhstan on April 6th; the second was launched from Plesetsk, Russia on April 9. Air parcel trajectory calculations and long-lived tracer gas concentrations in the cloud indicate that the Baikonur rocket launch is the most probable source of the plume. The parcel trajectory calculations do not unambiguously trace the transport of the Soyuz plume from Asia to North America, illustrating serious flaws in the point-to-point trajectory calculations. This chance encounter represents the only measurement of the stratospheric effects of emissions from a rocket powered exclusively with hydrocarbon fuel.

  14. Retro Rocket Motor Self-Penetrating Scheme for Heat Shield Exhaust Ports

    NASA Technical Reports Server (NTRS)

    Marrese-Reading, Colleen; St.Vaughn, Josh; Zell, Peter; Hamm, Ken; Corliss, Jim; Gayle, Steve; Pain, Rob; Rooney, Dan; Ramos, Amadi; Lewis, Doug; Shepherd, Joe; Inaba, Kazuaki

    2009-01-01

    A preliminary scheme was developed for base-mounted solid-propellant retro rocket motors to self-penetrate the Orion Crew Module heat shield for configurations with the heat shield retained during landings on Earth. In this system the motors propel impactors into structural push plates, which in turn push through the heat shield ablator material. The push plates are sized such that the remaining port in the ablator material is large enough to provide adequate flow area for the motor exhaust plume. The push plate thickness is sized to assure structural integrity behind the ablative thermal protection material. The concept feasibility was demonstrated and the performance was characterized using a gas gun to launch representative impactors into heat shield targets with push plates. The tests were conducted using targets equipped with Fiberform(R) and PICA as the heat shield ablator material layer. The PICA penetration event times were estimated to be under 30 ms from the start of motor ignition. The mass of the system (not including motors) was estimated to be less than 2.3 kg (5 lbs) per motor. The configuration and demonstrations are discussed.

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

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

    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.

  16. Flow field description of the Space Shuttle Vernier reaction control system exhaust plumes

    NASA Technical Reports Server (NTRS)

    Cerimele, Mary P.; Alred, John W.

    1987-01-01

    The flow field for the Vernier Reaction Control System (VRCS) jets of the Space Shuttle Orbiter has been calculated from the nozzle throat to the far-field region. The calculations involved the use of recently improved rocket engine nozzle/plume codes. The flow field is discussed, and a brief overview of the calculation techniques is presented. In addition, a proposed on-orbit plume measurement experiment, designed to improve future estimations of the Vernier flow field, is addressed.

  17. Analyzing rocket plume spectral data with neural networks

    NASA Technical Reports Server (NTRS)

    Whitaker, Kevin W.; Krishnakumar, K. S.; Benzing, Daniel A.

    1995-01-01

    The Optical Plume Anomaly Detection (OPAD) system is under development to provide early-warning failure detection in support of ground-level testing of the Space Shuttle Main Engine (SSME). Failure detection is to be achieved through the acquisition of spectrally resolved plume emissions and subsequent identification of abnormal levels indicative of engine corrosion or component failure. Two computer codes (one linear and the other non-linear) are used by the OPAD system to iteratively determine specific element concentrations in the SSME plume, given emission intensity and wavelength information. Since this analysis is extremely labor intensive, a study was initiated to develop neural networks that would model the 'inverse' of these computer codes. Optimally connected feed-forward networks with imperceptible prediction error have been developed for each element modeled by the linear code, SPECTRA4. Radial basis function networks were developed for the non-linear code, SPECTRA5, and predict combustion temperature in addition to element concentrations.

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

    NASA Astrophysics Data System (ADS)

    Ogungbemi, Kayode I.

    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

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

    NASA Technical Reports Server (NTRS)

    Sambamurthi, Jay K.

    1995-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    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.

  1. UN/visible absorption by OH radical in a hybrid rocket plume

    SciTech Connect

    Felix, T.M.; Teague, M.W.

    1995-12-01

    A spectrometer system was constructed for measurement of transient species in flames. A xenon arc lamp was used as the source of ultraviolet/visible radiation which was focused through the flame and onto a monochromator equipped with an ICCD detector. The system was used to measure absorption by OH radical around 306 nm in the plume of a hybrid rocket. Hydroxyl terminated polybutadiene (HTPB) was used as fuel and gaseous oxygen as oxidizer. The experimental spectra were analyzed by comparison with known vibration/rotational lines using a multiparameter curve-fitting program provided by Dr. Anthony Kotlar of the Army Research Laboratory, Aberdeen Proving Ground, Maryland. OH concentration and temperature profiles of the rocket plume have been determined and will be presented along with details of the spectrometer specifications.

  2. Axisymmetric computational fluid dynamics analysis of a film/dump-cooled rocket nozzle plume

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    Prediction of convective base heating rates for a new launch vehicle presents significant challenges to analysts concerned with base environments. The present effort seeks to augment classical base heating scaling techniques via a detailed investigation of the exhaust plume shear layer of a single H2/O2 Space Transportation Main Engine (STME). Use of fuel-rich turbine exhaust to cool the STME nozzle presented concerns regarding potential recirculation of these gases to the base region with attendant increase in the base heating rate. A pressure-based full Navier-Stokes computational fluid dynamics (CFD) code with finite rate chemistry is used to predict plumes for vehicle altitudes of 10 kft and 50 kft. Levels of combustible species within the plume shear layers are calculated in order to assess assumptions made in the base heating analysis.

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

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

    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.

  4. Numerical simulations of the flowfields of industrial ventilation systems and solar rocket plume

    SciTech Connect

    Yu, Shengtao.

    1989-01-01

    The motivation for this research is to incorporate modern numerical methods in modeling the flowfields of two systems: (1) industrial ventilation systems and (2) solar rocket plume. For both systems, calculations of the velocity, temperature, turbulence properties, and species concentration of flowfields were performed. Brief discussions of the two topics follow: (1) Industrial ventilation systems. An open vessel equipped with a push-pull ventilation system to control toxic vapor and a flanged suction inlet to control grinding particles and welding fumes has been analyzed. The computational method involves solving the two-dimensional turbulent flow equations for the conservation of mass, momentum, energy, turbulence properties, and chemical species in finite form. The method provides information needed by engineers to assess the effectiveness of their designs. In order to verify the accuracy of the theoretical analysis, a two-dimensional push-pull system prototype was set up and color schlieren photography and hot wire anemometry were performed. Favorable agreement was found between the experimental data and calculated results. (2) Solar rocket plume. The interaction of the solar rocket plume and the solar concentrator is studied by flow-field analysis. Such interaction can adversely affect the collector performance through fouling, excessive heat, or pressure loading. The geometrical shape of the concentrator is such that only the flow from the nozzle boundary layer can reach it, but the thrust levels of interest lead to very thick boundary layers. A time-marching Parabolized Navier-Stokes (PNS) scheme is developed to calculate the flowfields inside nozzles. The Method of Characteristics (MOC) is used to simulate the flow of rocket plume. Results show that both pressure and heat transfer effects are low, but that they increase as the chamber pressure or the thrust level size is reduced.

  5. Daytime midlatitude plasma depletions observed by Swarm: Topside signatures of the rocket exhaust

    NASA Astrophysics Data System (ADS)

    Park, Jaeheung; Kil, Hyosub; Stolle, Claudia; Lühr, Hermann; Coley, William R.; Coster, Anthea; Kwak, Young-Sil

    2016-03-01

    The daytime midlatitude plasma depletions (DMLPDs) observed on 22 May 2014 and 20 May 2015 by the Swarm constellation are not explained by any known natural phenomena. The DMLPDs were detected after rocket launches, and the DMLPD traces converged to the launch station. The event in 2015, for which sufficient total electron content (TEC) data are available, is accompanied with TEC depletion lasting for about 6 h. The persistence generally agrees with the lifetime expected for rocket exhaust depletions (REDs) which is determined by the recombination of the ionospheric oxygen ion with water molecules in the rocket exhaust. These results lead to the conclusion that DMLPDs are REDs in the topside. The RED characteristics identified from the observations on both days are (1) enhancement in electron temperature, (2) reduction in electron pressure, and (3) absence of substructures down to scale sizes of about 8 km (Nyquist's scale size).

  6. Chance Encounter with a Stratospheric Kerosene Rocket Plume From Russia Over California

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Wilson, J. C.; Ross, M. N.; Brock, C. A.; Sheridan, P. J.; Schoeberl, M. R.; Lait, L. R.; Bui, T. P.; Loewenstein, M.; Podolske, J. R.; Einaudi, Franco (Technical Monitor)

    2000-01-01

    A high-altitude aircraft flight on April 18, 1997 detected an enormous aerosol cloud at 20 km altitude near California (37 N). Not visually observed, the cloud had high concentrations of soot and sulfate aerosol, and was over 180 km in horizontal extent. The cloud was probably a large hydrocarbon fueled vehicle, most likely from rocket motors burning liquid oxygen and kerosene. One of two Russian Soyuz rockets could have produced the cloud: a launch from the Baikonur Cosmodrome, Kazakhstan on April 6; or from Plesetsk, Russia on April 9. Parcel trajectories and long-lived trace gas concentrations suggest the Baikonur launch as the cloud source. Cloud trajectories do not trace the Soyuz plume from Asia to North America, illustrating the uncertainties of point-to-point trajectories. This cloud encounter is the only stratospheric measurement of a hydrocarbon fuel powered rocket.

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

    NASA Technical Reports Server (NTRS)

    Panda, J.; Mosher, R.

    2010-01-01

    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

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

    NASA Astrophysics Data System (ADS)

    Bernhardt, P. A.

    2009-12-01

    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 exchange. Long-lived plasma irregularities are formed by the artificial hypersonic “neutral wind” passing through the ionosphere. Charge exchange between the fast neutrals and the ambient ions yields high-speed ion beams that excite electro-static plasma waves. Ground based radar has been used to detect both field aligned irregularities and electrostatic turbulence driven by the Space Shuttle OMS exhaust. Molecular ions produced by the charge exchange with molecules in the rocket exhaust recombine with a time scale of 10 minutes leaving a residual plasma depression. This ionospheric “hole” fills in by ambipolar diffusion leaving a depleted magnetic flux tube. This large scale reduction in Pedersen conductivity can provide a seed for plasma interchange instabilities. For instance, a rocket firing on the bottom side of the ionosphere near the equator can trigger a Rayleigh-Taylor instability that is naturally seen as equatorial Spread-F. The Naval Research Laboratory has been exploring these phenomena with dedicated burns of the Space Shuttle OMS engines and exhaust releases from rockets. The Shuttle Ionospheric Modification with Pulsed Localized Exhaust (SIMPLEX) series of experiments uses ground radars to probe the ionosphere affected by dedicated burns of the Space Shuttle OMS engines. Radars located at Millstone Hill, Massachusetts; Arecibo, Puerto Rico; Jicamarca, Peru; Kwajalein, Marshall Island; and Alice Springs, Australia have participated in the SIMPLEX program. A companion program called Shuttle Exhaust Ionospheric Turbulence Experiment has or will use satellites to fly through the turbulence

  9. Stratospheric aircraft exhaust plume and wake chemistry studies

    NASA Technical Reports Server (NTRS)

    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.

    1992-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    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.

  11. The washout of combustion-generated hydrogen chloride. [rocket exhaust raindrop scavenging quantification

    NASA Technical Reports Server (NTRS)

    Fenton, D. L.; Purcell, R. Y.; Hrdina, D.; Knutson, E. O.

    1980-01-01

    The coefficient for the washout from a rocket exhaust cloud of HCl generated by the combustion of an ammonium perchlorate-based solid rocket propellant such as that to be used for the Space Shuttle Booster is determined. A mathematical model of HCl scavenging by rain is developed taking into account rain droplet size, fall velocity and concentration under various rain conditions, partitioning of exhaust HCl between liquid and gaseous phases, the tendency of HCl to promote water vapor condensation and the concentration and size of droplets within the exhaust cloud. The washout coefficient is calculated as a function of total cloud water content, total HCl content at 100% relative humidity, condensation nuclei concentration and rain intensity. The model predictions are compared with experimental results obtained in scavenging tests with solid rocket exhaust and raindrops of different sizes, and the large reduction in washout coefficient at high relative humidities predicted by the model is not observed. A washout coefficient equal to 0.0000512 times the -0.176 power of the mass concentration of HCl times the 0.773 power of the rainfall intensity is obtained from the experimental data.

  12. A Transonic and Surpersonic Investigation of Jet Exhaust Plume Effects on the Afterbody and Base Pressures of a Body of Revolution

    NASA Technical Reports Server (NTRS)

    Andrews, C. D.; Cooper, C. E., Jr.

    1974-01-01

    An experimental aerodynamic investigation was conducted to provide data for studies to determine the criteria for simulating rocket engine plume induced aerodynamic effects in the wind tunnel using a simulated gaseous plume. Model surface and base pressure data were obtained in the presence of both a simulated and a prototype gaseous plume for a matrix of plume properties to enable investigators to determine the parameters that correlate the simulated and prototype plume-induced data. The test program was conducted in the Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel using two models, the first being a strut mounted cone-ogive-cylinder model with a fineness ratio of 9. Model exterior pressures, model plenum chamber and nozzle performance data were obtained at Mach numbers of 0.9, 1.2, 1.46, and 3.48. The exhaust plume was generated by using air as the simulant gas, or Freon-14 (CF4) as the prototype gas, over a chamber pressure range from 0 to 2,000 psia and a total temperature range from 50 to 600 F.

  13. Pseudo Color Densitometer Analysis-the Apollo 17/Saturn V Exhaust Plume.

    PubMed

    Orville, R E; Helsdon, J H

    1974-10-01

    Spectra of the Apollo 17/Saturn V exhaust plume have been obtained in the uv (300ndash;400 nm), visible (400-650 nm), and ir (750-790 nm) regions. Analysis of these data with a pseudo color densitometer reveals (1) a standing wave pattern in the exhaust plume characterized by a wavelength of 9 m, (2) a region of intense continuum within 40 m of the exit plane which supports previous reports of a continuum blackbody source with a peak temperature near 2600 K, (3) a region of continuum emission beyond 40 m that is not blackbody, and (4) line emissions beyond 40 m attributed to the sodium D lines and potassium. It is suggested that an interference filter centered on the sodium D lines could be used on a high speed framing camera to study the turbulent structure of the plume in the nonblackbody region.

  14. Remote measurement of the plume shape of aircraft exhausts at airports by passive FTIR spectrometry

    NASA Astrophysics Data System (ADS)

    Schafer, Klaus; Jahn, Carsten; Utzig, Selina; Flores-Jardines, Edgar; Harig, Roland; Rusch, Peter

    2004-11-01

    Information about the interaction between the exhaust plume of an aircraft jet engine and ambient air is required for the application of small-scale chemistry-transport models to investigate airport air quality. This interaction is not well understood. In order to study the interaction, spatial information about the plume is required. FTIR emission spectroscopy may be applied to analyze the aircraft exhausts. In order to characterize the plumes spatially, a scanning imaging FTIR system (SIGIS) has been improved. SIGIS is comprised of an interferometer (Bruker OPAG), an azimuth-elevation-scanning mirror, a data acquisition and control system with digital signal processors (DSP), an infrared camera and a personal computer. With this instrumentation it is possible to visualise the plume and to obtain information about the temperature distribution within the plume. Measurements are performed at low spectral resolution, because the dynamic environment of these measurements limits the measurement time to about 2 minutes. Measurements of the plume shapes of an APU and of main engines were performed.

  15. Range safety signal attenuation by the Space Shuttle main engine exhaust plumes

    NASA Technical Reports Server (NTRS)

    Pearce, B. E.

    1983-01-01

    An analysis of attenuation of the range safety signal at 416.5 MHz observed after SRB separation and ending at hand over to Bermuda, during which transmission must pass through the LOX/H2 propelled main engine exhaust plumes, is summarized. Absorption by free electrons in the exhaust plume can account for the nearly constant magnitude of the observed attenuation during this period; it does not explain the short term transient increases that occur at one or more times during this portion of the flight. It is necessary to assume that a trace amount (about 0.5 ppm) of easily ionizable impurity must be present in the exhaust flow. Other mechanisms of attenuation, such as scattering by turbulent fluctuations of both free and bound electrons and absorption by water vapor, were examined but found to be inadequate to explain the observations.

  16. 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)

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

    2014-01-01

    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.

  17. The effects of an ion-thruster exhaust plume on S-band carrier transmission

    NASA Technical Reports Server (NTRS)

    Ackerknecht, W. E.; Stanton, P. H.

    1976-01-01

    The study reported here was undertaken (1) to develop models of the effects of an ion-thruster exhaust plume on S-band signals, and (2) to measure the effects. The results show that an S-band signal passing through an ion-thruster plume is reduced in amplitude and advanced in phase. The mathematical models gave reasonable estimates of the average signal attenuation and phase shift. Negligible fluctuations in the signal amplitude and phase were measured during steady-state thruster operation. However, large jumps in phase occurred when changes were made in the thruster operating state. This study confirms that the thruster plume can have a significant effect on S-band communication link performance; hence the plume effects must be considered in S-band link calculations when electric thrusters are used for spacecraft propulsion.

  18. Rocket exhaust effluent modeling for tropospheric air quality and environmental assessments

    NASA Technical Reports Server (NTRS)

    Stephens, J. B.; Stewart, R. B.

    1977-01-01

    The various techniques for diffusion predictions to support air quality predictions and environmental assessments for aerospace applications are discussed in terms of limitations imposed by atmospheric data. This affords an introduction to the rationale behind the selection of the National Aeronautics and Space Administration (NASA)/Marshall Space Flight Center (MSFC) Rocket Exhaust Effluent Diffusion (REED) program. The models utilized in the NASA/MSFC REED program are explained. This program is then evaluated in terms of some results from a joint MSFC/Langley Research Center/Kennedy Space Center Titan Exhaust Effluent Prediction and Monitoring Program.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1995-01-01

    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.

  1. Modification of the upper atmosphere with chemicals found in rocket exhaust

    SciTech Connect

    Bernhardt, P.A.; Zinn, J.; Mendillo, M.; Baumgardner, J.

    1982-01-01

    Rockets, burning above 200 km altitude, release exhaust vapors which react chemically with the plasma comprising the F-region ionosphere. The two major types of atmospheric modification produced by rocket exhaust are: (1) the formation of large scale ionospheric holes, and (2) the enhancement of the airglow emissions. The ionospheric holes are regions tens of kilometers in diameter where the plasma concentration can be reduced by a factor of ten or more. Plasma instabilities may produce irregularities at the edges of the holes. Communication and navigation systems relying on radio propagation through the modified ionosphere may be affected. Airglow enhancements are a result of excited neutral species being produced by chemical reactions between the rocket exhaust and the ionospheric plasma. For example, the 630 nm line from atomic oxygen may increase twenty-fold in intensity over the ambient level. This paper reviews experimental observations and theoretical treatments of ionospheric modification produced by gas releases in the upper atmosphere. Recent experimental measurements of the ionospheric modification by an ATLAS-F launch vehicle are presented. The plans for future experiments are discussed.

  2. Crew Launch Vehicle Mobile Launcher Solid Rocket Motor Plume Induced Environment

    NASA Technical Reports Server (NTRS)

    Vu, Bruce T.; Sulyma, Peter

    2008-01-01

    The plume-induced environment created by the Ares 1 first stage, five-segment reusable solid rocket motor (RSRMV) will impose high heating rates and impact pressures on Launch Complex 39. The extremes of these environments pose a potential threat to weaken or even cause structural components to fail if insufficiently designed. Therefore the ability to accurately predict these environments is critical to assist in specifying structural design requirements to insure overall structural integrity and flight safety. This paper presents the predicted thermal and pressure environments induced by the launch of the Crew Launch Vehicle (CLV) from Launch Complex (LC) 39. Once the environments are predicted, a follow-on thermal analysis is required to determine the surface temperature response and the degradation rate of the materials. An example of structures responding to the plume-induced environment will be provided.

  3. Space Shuttle Solid Rocket Motor Plume Pressure and Heat Rate Measurements

    NASA Technical Reports Server (NTRS)

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

    2012-01-01

    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.

  4. Optical Measurements on Solid Specimens of Solid Rocket Motor Exhaust and Solid Rocket Motor Slag

    NASA Technical Reports Server (NTRS)

    Roberts, F. E., III

    1991-01-01

    Samples of aluminum slag were investigated to aid the Earth Science and Applications Division at the Marshall Space Flight Center (MSFC). Alumina from space motor propellant exhaust and space motor propellant slag was examined as a component of space refuse. Thermal emittance and solar absorptivity measurements were taken to support their comparison with reflectance measurements derived from actual debris. To determine the similarity between the samples and space motor exhaust or space motor slag, emittance and absorbance results were correlated with an examination of specimen morphology.

  5. Exhaust plumes and their interaction with missile airframes - A new viewpoint

    NASA Technical Reports Server (NTRS)

    Dash, S. M.; Sinha, N.

    1992-01-01

    The present, novel treatment of missile airframe-exhaust plume interactions emphasizes their simulation via a formal solution of the Reynolds-averaged Navier-Stokes (RNS) equation and is accordingly able to address the simulation requirements of novel missiles with nonconventional/integrated propulsion systems. The method is made possible by implicit RNS codes with improved artificial dissipation models, generalized geometric capabilities, and improved two-equation turbulence models, as well as by such codes' recent incorporation of plume thermochemistry and multiphase flow effects.

  6. The effect of exhaust plume/afterbody interaction on installed Scramjet performance

    NASA Technical Reports Server (NTRS)

    Edwards, Thomas Alan

    1988-01-01

    Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has

  7. Abatement of an aircraft exhaust plume using aerodynamic baffles.

    PubMed

    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

    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.

  8. Abatement of an aircraft exhaust plume using aerodynamic baffles.

    PubMed

    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

    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

  9. Some physical and thermodynamic properties of rocket exhaust clouds measured with infrared scanners

    NASA Technical Reports Server (NTRS)

    Gomberg, R. I.; Kantsios, A. G.; Rosensteel, F. J.

    1977-01-01

    Measurements using infrared scanners were made of the radiation from exhaust clouds from liquid- and solid-propellant rocket boosters. Field measurements from four launches were discussed. These measurements were intended to explore the physical and thermodynamic properties of these exhaust clouds during their formation and subsequent dispersion. Information was obtained concerning the initial cloud's buoyancy, the stabilized cloud's shape and trajectory, the cloud volume as a function of time, and it's initial and stabilized temperatures. Differences in radiation intensities at various wavelengths from ambient and stabilized exhaust clouds were investigated as a method of distinguishing between the two types of clouds. The infrared remote sensing method used can be used at night when visible range cameras are inadequate. Infrared scanning techniques developed in this project can be applied directly to natural clouds, clouds containing certain radionuclides, or clouds of industrial pollution.

  10. Far-Field Turbulent Vortex-Wake/Exhaust Plume Interaction for Subsonic and HSCT Airplanes

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Adam, Ihab; Wong, Tin-Chee

    1996-01-01

    Computational study of the far-field turbulent vortex-wake/exhaust plume interaction for subsonic and high speed civil transport (HSCT) airplanes is carried out. The Reynolds-averaged Navier-Stokes (NS) equations are solved using the implicit, upwind, Roe-flux-differencing, finite-volume scheme. The two-equation shear stress transport model of Menter is implemented with the NS solver for turbulent-flow calculation. For the far-field study, the computations of vortex-wake interaction with the exhaust plume of a single engine of a Boeing 727 wing in a holding condition and two engines of an HSCT in a cruise condition are carried out using overlapping zonal method for several miles downstream. These results are obtained using the computer code FTNS3D. The results of the subsonic flow of this code are compared with those of a parabolized NS solver known as the UNIWAKE code.

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

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

    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.

  12. Exhaust Plume Effects on Sonic Boom for a Delta Wing and a Swept Wing-Body Model

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Lake, Troy

    2012-01-01

    Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic boom. Research has been performed on sonic boom reduction and has included the contribution of the exhaust nozzle plume. Plume effect on sonic boom has progressed from the study of isolated nozzles to a study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline analysis of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field pressure profile. Reductions occurred in the peak-to-peak magnitude of the pressure profile for a swept wing-body vehicle. The exhaust plumes also had a favorable effect as the nozzles were moved outward along the wing-span.

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

    NASA Technical Reports Server (NTRS)

    Sambamurthi, Jay K.

    1995-01-01

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

  14. Real Time Diagnostics of Jet Engine Exhaust Plumes Using a Chirped QC Laser Spectrometer

    NASA Astrophysics Data System (ADS)

    Hay, K. G.; Duxbury, G.; Langford, N.

    2010-06-01

    Quantitative measurements of real-time variations of the chemical composition of a jet engine exhaust plume is demonstrated using a 4.86 μmn intra-pulse quantum cascade laser spectrometer. The measurements of the gas turbine exhaust were carried out in collaboration with John Black and Mark Johnson at Rolls Royce. The recording of five sets of averaged spectra a second has allowed us to follow the build up of the combustion products within the exhaust, and to demonstrate the large variation of the integrated absorption of these absorption lines with temperature. The absorption cross sections of the lines of both carbon monoxide and water increase with temperature, whereas those of the three main absorption lines of carbon dioxide decrease. At the steady state limit the absorption lines of carbon dioxide are barely visible, and the spectrum is dominated by absorption lines of carbon monoxide and water.

  15. Rocket engine plume diagnostics using video digitization and image processing - Analysis of start-up

    NASA Technical Reports Server (NTRS)

    Disimile, P. J.; Shoe, B.; Dhawan, A. P.

    1991-01-01

    Video digitization techniques have been developed to analyze the exhaust plume of the Space Shuttle Main Engine. Temporal averaging and a frame-by-frame analysis provide data used to evaluate the capabilities of image processing techniques for use as measurement tools. Capabilities include the determination of the necessary time requirement for the Mach disk to obtain a fully-developed state. Other results show the Mach disk tracks the nozzle for short time intervals, and that dominate frequencies exist for the nozzle and Mach disk movement.

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

    SciTech Connect

    Jenkins, R.A.

    1990-06-07

    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.

  17. In situ exhaust cloud measurements. [particle size distribution and cloud physics of rocket exhaust clouds

    NASA Technical Reports Server (NTRS)

    Wornom, D.

    1980-01-01

    Airborne in situ exhaust cloud measurements were conducted to obtain definitions of cloud particle size range, Cl2 content, and HCl partitioning. Particle size distribution data and Cl2 measurements were made during the May, August, and September 1977 Titan launches. The measurements of three basic effluents - HCl, NO sub X, and particles - against minutes after launch are plotted. The maximum observed HCl concentration to the maximum Cl2 concentration are compared and the ratios of the Cl2 to the HCl is calculated.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1971-01-01

    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.

  20. Modeling Macro- and Micro-Scale Turbulent Mixing and Chemistry in Engine Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Menon, Suresh

    1998-01-01

    Simulation of turbulent mixing and chemical processes in the near-field plume and plume-vortex regimes has been successfully carried out recently using a reduced gas phase kinetics mechanism which substantially decreased the computational cost. A detailed mechanism including gas phase HOx, NOx, and SOx chemistry between the aircraft exhaust and the ambient air in near-field aircraft plumes is compiled. A reduced mechanism capturing the major chemical pathways is developed. Predictions by the reduced mechanism are found to be in good agreement with those by the detailed mechanism. With the reduced chemistry, the computer CPU time is saved by a factor of more than 3.5 for the near-field plume modeling. Distributions of major chemical species are obtained and analyzed. The computed sensitivities of major species with respect to reaction step are deduced for identification of the dominant gas phase kinetic reaction pathways in the jet plume. Both the near field plume and the plume-vortex regimes were investigated using advanced mixing models. In the near field, a stand-alone mixing model was used to investigate the impact of turbulent mixing on the micro- and macro-scale mixing processes using a reduced reaction kinetics model. The plume-vortex regime was simulated using a large-eddy simulation model. Vortex plume behind Boeing 737 and 747 aircraft was simulated along with relevant kinetics. Many features of the computed flow field show reasonable agreement with data. The entrainment of the engine plumes into the wing tip vortices and also the partial detrainment of the plume were numerically captured. The impact of fluid mechanics on the chemical processes was also studied. Results show that there are significant differences between spatial and temporal simulations especially in the predicted SO3 concentrations. This has important implications for the prediction of sulfuric acid aerosols in the wake and may partly explain the discrepancy between past numerical studies

  1. Dynamic Analysis of a Building Under Rocket Engine Plume Acoustic Load

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

    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.

  2. Possible effect of the chlorine oxide dimer on transient ozone loss in rocket plumes. Technical report

    SciTech Connect

    Martin, L.R.

    1994-03-15

    Understanding transient, local ozone holes that may be produced by solid rocket boosters in the stratosphere puts special demands on models. One must consider the time scales as well as the rates for all of the pertinent chemical reactions involved in the destruction of ozone. In this report, we show that consideration of the existence of the chlorine oxide dimer, Cl2O2, and consideration of the necessary time scale for ozone loss are essential for prediction of a transient ozone hole. We argue that photolysis of this species is the major source of atomic chlorine in the plume at 20 km, and the ClO + 0 reaction is the major source at 30 km, although both processes play a role at the higher altitude. Inclusion of the chlorine oxide dimer ozone destruction cycle, which has not been considered in any of the full-scale models to date, predicts substantial ozone destruction on a scale of about 12-km diameter at 20-km altitude and the ClO cycle produces a 49-km-diameter hole at 30-km altitude. This analysis also suggests that the size of the hole at 20 km may be highly variable since it is sensitive to the variable ozone-to-methane ratio at that altitude. Ozone, Rocket launch, Stratospheric, Chlorine.

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

    NASA Astrophysics Data System (ADS)

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

    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 ΔCO2 and ΔT and were far above the local ambient atmospheric background-level of typically 15-50 pptv. Our observations indicate a lower limit for the efficiency ɛ for fuel-sulfur conversion to SA of 0.34 %.

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

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

    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.

  5. Exhaust plume and contamination characteristics of a bipropellant (MMH/N2O4) RCS thruster

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    Results are presented for three recent tests in a series of thruster contamination experiments made in liquid helium-cooled environmental facility. The contaminating effects encountered on various materials, surfaces, and components, due to the exhaust products from a 5-pound thrust, bipropellant (MMH/N2O4) thruster are investigated. The angular distribution of plume effects around the periphery of the thruster established by transmittance changes of quartz samples over the wavelength range from 0.2 to 2.0 micrometer is studied, along with mass deposition rates at a specific location measured with a quartz crystal microbalance for three different experiments. Quadrupole mass spectrometer measurements of the exhaust products over the mass number range from 12 to 75; infrared transmittance measurements of contaminated samples for the wavelength range from 2.5 to 15 microns; and infrared transmittance measurements of residue from the thruster nozzle are also considered.

  6. On the prediction of concentration variations in a dispersing heavy-duty truck exhaust plume using k- ɛ turbulent closure

    NASA Astrophysics Data System (ADS)

    Kim, Dong-Hee; Gautam, Mridul; Gera, Dinesh

    This work presents the computational fluid dynamic modeling of an exhaust plume dispersed from the exhaust pipe of a class-8 tractor truck powered by 330 hp Cummins M11 electronically controlled diesel engine. This effort utilizes an advanced CFD technique to accurately predict the variation of carbon dioxide concentration inside a turbulent plume using a k- ɛ eddy dissipation model. The simulation includes the "real-world" operation of a truck and its exhaust plume in a NASA, Langley aircraft testing wind tunnel, that had an effective volume of 226, 535 m 3 (8,000,000 ft 3). The predicted results show an excellent agreement with the experimentally measured values of CO 2 concentrations, dilution ratios, and the temperature variations inside the plume. A specific goal of this effort was to study the effect of recirculation region near the truck walls on dispersion of the plume. For this purpose, growth of the plume from the center of the exhaust pipe is also presented and discussed. This work also shows the benefits of CFD modeling in applications where dispersion correlations are not required a priori, instead the dispersion coefficients are calculated precisely by solving the turbulent kinetic energy and dissipation equations.

  7. Analytic model for washout of HCl(g) from dispersing rocket exhaust clouds

    NASA Technical Reports Server (NTRS)

    Pellett, G. L.

    1981-01-01

    The potential is investigated that precipitation scavenging of HCl from large solid rocket exhaust clouds may result in unacceptably acidic rain in the Cape Canaveral, Florida, area before atmospheric dispersion reduces HCl concentrations to safe limits. Several analytic expressions for HCl(g) and HCl(g + aq) washout are derived; a geometric mean washout coefficient is recommended. A previous HCl washout model is refined and applied to a space shuttle case (70 t HCl exhausted up to 4 km) and eight Titan 3 (60 percent less exhaust) dispersion cases. The vertical column density (sigma) decays were deduced by application of a multilayer Gaussian diffusion model to seven standard meteorological regimes for overland advection. The Titan 3 decays of sigma and initial rain pH differed greatly among regimes; e.g., a range of 2 pH units was spanned at x 100 km downwind and t = 2 hr. Environmentally significant pH's .5 for infrequent exposures were shown possible at X = 50 km and t 5 hr for the two least dispersive Titan 3 cases. Representative examples of downwind rainwater pH and G(X) are analyzed. Factors affecting the validity of the results are discussed.

  8. Coherent anti-Stokes Raman spectroscopy (CARS) and laser-induced fluorescence (LIF) measurements in a rocket engine plume

    SciTech Connect

    Williams, D.R.; McKeown, D.; Porter, F.M.; Baker, C.A.; Astill, A.G.; Rawley, K.M. . Combustion Dept. Epsilon Research, Buckinghamshire Defence Research Agency, Fort Halstead, Kent )

    1993-07-01

    Coherent anti-Stokes Raman spectroscopy (CARS) and laser-induced fluorescence (LIF) measurements in the plume of a liquid-fueled rocket engine are compared with the results predicted by a mathematical model of the plume. At most positions, high signal success rates were obtained. Success rates were lower during initial runs, while the system was optimized for operation in the rocket environment, and on axis close to the nozzle where the probing laser beams were severely deflected by the plume. For each position studied, the spectra taken were fitted for temperature and a mean temperature and standard deviation calculated from the results. The mean temperatures were compared with predicted temperature values obtained from a marching procedure parabolic computer program. CARS spectra from water vapor in the plume were also recorded and fitted for temperature and concentration. Excellent agreement between theory and experiment was obtained. Results showed a strong positive correlation between water vapor concentration and temperature at each measurement position--some contributions to this may arise from similarities of the effects of temperature and concentration on spectral shape. However, shear layer mixing and entrainment of cold gas into the plume may significantly affect the composition and temperature of the plume gases. LIF was used to visualize the plume structure. Imaging of the flow field was performed by detecting sodium fluorescence, after the oxidant was seeded with sodium. Images were obtained without excessively high background levels and large fluctuations in the plume structure were observed. This is consistent with the observations from the CARS experiments.

  9. Results of an investigation of jet plume effects on an 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 9 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82B), volume 1. [an exhaust flow simulation

    NASA Technical Reports Server (NTRS)

    Hawthorne, P. J.

    1976-01-01

    The base pressure environment was investigated for the first and second stage mated vehicle in a supersonic flow field from Mach 1.55 through 2.20 with simulated rocket engine exhaust plumes. The pressure environment was investigated for the orbiter at various vent port locations at these same freestream conditions. The Mach number environment around the base of the model with rocket plumes simulated was examined. 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 the tested configurations are shown.

  10. In situ observations in aircraft exhaust plumes in the lower stratosphere at midlatitudes

    NASA Technical Reports Server (NTRS)

    Fahey, D. W.; Keim, E. R.; Woodbridge, E. L.; Gao, R. S.; Boering, K. A.; Daube, B. C.; Wofsy, S. C.; Lohmann, R. P.; Hintsa, E. J.; Dessler, A. E.

    1995-01-01

    Instrumentation on the NASA ER-2 high-altitude aircraft has been used to observe engine exhaust from the same aircraft while operating in the lower stratosphere. Encounters with the exhaust plume occurred approximately 10 min after emission with spatial scales near 2 km and durations of up to 10 s. Measurements include total reactive nitrogen, NO(y), the component species NO and NO2, CO2, H2O, CO, N2O, condensation nuclei, and meteorological parameters. The integrated amounts of CO2 and H2O during the encounters are consistent with the stoichiometry of fuel combustion (1:1 molar). Emission indices (EI) for NO(x) (= NO + NO2), CO, and N2O are calculated using simultaneous measurements of CO2. EI values for NO(x) near 4 g/(kg fuel) are in good agreement with values scaled from limited ground-based tests of the ER-2 engine. Non-NO(x) species comprise less than about 20% of emitted reactive nitrogen, consistent with model evaluations. In addition to demonstrating the feasibility of aircraft plume detection, these results increase confidence in the projection of emissions from current and proposed supersonic aircraft fleets and hence in the assessment of potential long-term changes in the atmosphere.

  11. Design of Experiments for Both Experimental and Analytical Study of Exhaust Plume Effects on Sonic Boom

    NASA Technical Reports Server (NTRS)

    Castner, Raymond S.

    2009-01-01

    Computational fluid dynamics (CFD) analysis has been performed to study the plume effects on sonic boom signature for isolated nozzle configurations. The objectives of these analyses were to provide comparison to past work using modern CFD analysis tools, to investigate the differences of high aspect ratio nozzles to circular (axisymmetric) nozzles, and to report the effects of under expanded nozzle operation on boom signature. CFD analysis was used to address the plume effects on sonic boom signature from a baseline exhaust nozzle. Nearfield pressure signatures were collected for nozzle pressure ratios (NPRs) between 6 and 10. A computer code was used to extrapolate these signatures to a ground-observed sonic boom N-wave. Trends show that there is a reduction in sonic boom N-wave signature as NPR is increased from 6 to 10. As low boom designs are developed and improved, there will be a need for understanding the interaction between the aircraft boat tail shocks and the exhaust nozzle plume. These CFD analyses will provide a baseline study for future analysis efforts. For further study, a design of experiments has been conducted to develop a hybrid method where both CFD and small scale wind tunnel testing will validate the observed trends. The CFD and testing will be used to screen a number of factors which are important to low boom propulsion integration, including boat tail angle, nozzle geometry, and the effect of spacing and stagger on nozzle pairs. To design the wind tunnel experiment, CFD was instrumental in developing a model which would provide adequate space to observe the nozzle and boat tail shock structure without interference from the wind tunnel walls.

  12. A field study of solid rocket exhaust impacts on the near-field environment

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    Large solid rocket motors release large quantities of hydrogen chloride and aluminum oxide exhaust during launch and testing. Measurements and analysis of the interaction of this material with the deluge water spray and other environmental factors in the near field (within 1 km of the launch or test site) are summarized. Measurements of mixed solid and liquid deposition (typically 2 normal HCl) following space shuttle launches and 6.4 percent scale model tests are described. Hydrogen chloride gas concentrations measured in the hours after the launch of STS 41D and STS 51A are reported. Concentrations of 9 ppm, which are above the 5 ppm exposure limits for workers, were detected an hour after STS 51A. A simplified model which explains the primary features of the gas concentration profiles is included.

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

    NASA Technical Reports Server (NTRS)

    Scallion, William I.

    1991-01-01

    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.

  14. Search of archived data sources for rocket exhaust-induced modifications of the ionosphere

    SciTech Connect

    Chacko, C.C.; Mendillo, M.

    1980-09-01

    The emergence of the Satellite Power System (SPS) concept as a way of augmenting the dwindling energy sources available for commercial power usage involved such a large and unprecendented technological program that detailed assessment and feasibility studies were undertaken in an attempt to specify the true impact such a program would have. As part of the issues addressed, a comprehensive environmental impact study was initiated that involved an unprecedented scope of concerns ranging from ground-level noise and weather modifications to possible planetary-scale perturbations caused by SPS activity in distant Earth orbits. This report describes results of a study of an intermediate region of the Earth's environment (the ionosphere) where large-scale perturbations are caused by routine rocket activity. The SPS program calls for vast transportation demands into and out from the ionosphere (h approx. = 200 to 1000 km), and thus the well-known effect of chemical depletions of the ionosphere (so-called ionospheric holes) caused by rocket exhaust signaled a concern over the possible large-scale and long-term consequences of the induced effects.

  15. Space shuttle vehicle rocket plume impingement study for separation analysis. Tasks 2 and 3: Definition and preliminary plume impingement analysis for the MSC booster

    NASA Technical Reports Server (NTRS)

    Wojciechowski, C. J.; Penny, M. M.; Prozan, R. J.

    1970-01-01

    The results are presented of a space shuttle plume impingement study for the Manned Spacecraft Center configuration. This study was conducted as two tasks which were to (1) define the orbiter main stage engine exhaust plume flow field, and (2) define the plume impingement heating, force and resulting moment environments on the booster during the staging maneuver. To adequately define these environments during the staging maneuver and allow for deviation from the nominal separation trajectory, a multitude of relative orbiter/booster positions are analyzed which map the region that contains the separation trajectories. The data presented can be used to determine a separation trajectory which will result in acceptable impingement heating rates, forces, and the resulting moments. The data, presented in graphical form, include the effect of roll, pitch and yaw maneuvers for the booster. Quasi-steady state analysis methods were used with the orbiter engine operating at full thrust. To obtain partial thrust results, simple ratio equations are presented.

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

    SciTech Connect

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

    1980-10-01

    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.

  17. Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines: 1. Model calculation of the physical conditions in a jet exhaust

    NASA Astrophysics Data System (ADS)

    Platov, Yu. V.; Alpatov, V. V.; Klyushnikov, V. Yu.

    2014-01-01

    Model calculations have been performed for the temperature and pressure of combustion products in the jet exhaust of rocket engines of last stages of Proton, Molniya, and Start launchers operating in the upper atmosphere at altitudes above 120 km. It has been shown that the condensation of water vapor and carbon dioxide can begin at distances of 100-150 and 450-650 m away from the engine nozzle, respectively.

  18. Observation of the exhaust plume from the space shuttle main engine using the Microwave Limb Sounder

    NASA Astrophysics Data System (ADS)

    Pumphrey, H. C.; Lambert, A.; Livesey, N. J.

    2010-08-01

    A space shuttle launch deposits 700 t of water in the atmosphere. Some of this water is released into the upper mesosphere and lower thermosphere where it may be directly detected by a limb sounding satellite instrument. We report measurements of water vapour plumes from shuttle launches made by the Microwave Limb Sounder (MLS) on the Aura satellite. Approximately 50% of shuttle launches are detected by MLS. The signal appears at a similar level across the upper 10 km of the MLS limb scan, suggesting that the bulk of the observed water is above the top of the scan. Only a small fraction at best of smaller launches (Ariane, Proton) are detected. We conclude that the sensitivity of MLS is only just great enough to detect a shuttle sized launch, but that a suitably designed instrument of the same general type could detect the exhausts from a large proportion of heavy-lift launches.

  19. Observation of the exhaust plume from the space shuttle main engines using the microwave limb sounder

    NASA Astrophysics Data System (ADS)

    Pumphrey, H. C.; Lambert, A.; Livesey, N. J.

    2011-01-01

    A space shuttle launch deposits 700 tonnes of water in the atmosphere. Some of this water is released into the upper mesosphere and lower thermosphere where it may be directly detected by a limb sounding satellite instrument. We report measurements of water vapour plumes from shuttle launches made by the Microwave Limb Sounder (MLS) on the Aura satellite. Approximately 50%-65% of shuttle launches are detected by MLS. The signal appears at a similar level across the upper 10 km of the MLS limb scan, suggesting that the bulk of the observed water is above the top of the scan. Only a small fraction at best of smaller launches (Ariane 5, Proton) are detected. We conclude that the sensitivity of MLS is only just great enough to detect a shuttle sized launch, but that a suitably designed instrument of the same general type could detect the exhausts from a large proportion of heavy-lift launches.

  20. Analysis of Exhaust Plume Effects on Sonic Boom for a 59-Degree Wing Body Model

    NASA Technical Reports Server (NTRS)

    Castner, Raymond S.

    2011-01-01

    Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions are 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 focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analyses showed how the shock wave formed at the nozzle lip interacted with the nozzle boat-tail expansion wave. The nozzle lip shock moved with increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion. Lip shock movement caused a favorable change in the observed pressure signature. These results were applied to a simplified supersonic vehicle geometry with no inlets and no tail, in which the goal was to demonstrate how under-expanded nozzle operation reduced the sonic boom signature by twelve percent. A secondary goal was to demonstrate the use of the Cart3D inviscid code for off-body pressure signatures including the nozzle plume effect.

  1. Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Isolated Nozzles

    NASA Technical Reports Server (NTRS)

    Castner, Raymond S.

    2011-01-01

    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 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-ft Supersonic Wind Tunnel at the NASA Glenn Research Center to validate the computational study. Results demonstrated how the nozzle lip shock moved with increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion, causing a favorable change in the observed pressure signature. Experimental results were presented for comparison to the CFD results. The strong nozzle lip shock at high values of NPR intersected the nozzle boat-tail expansion and suppressed the expansion wave. Based on these results, it may be feasible to reduce the boat-tail expansion for a future supersonic aircraft with under-expanded nozzle exhaust flow by modifying nozzle pressure or nozzle divergent section geometry.

  2. Experimental research in the use of electrets in measuring effluents from rocket exhaust and a review of standard air quality measuring devices

    NASA Technical Reports Server (NTRS)

    Susko, M.

    1976-01-01

    Seven standard types of measuring devices used to obtain the chemical composition of rocket exhaust effluents were discussed. The electrets, a new measuring device, are investigated and compared with established measuring techniques. The preliminary results obtained show that electrets have multipollutant measuring capabilities, simplicity of deployment, speed of assessment or analysis, and may be an important and valuable tool in measuring pollutants from space vehicle rocket exhaust.

  3. Validation of Methods to Predict Vibration of a Panel in the Near Field of a Hot Supersonic Rocket Plume

    NASA Technical Reports Server (NTRS)

    Bremner, P. G.; Blelloch, P. A.; Hutchings, A.; Shah, P.; Streett, C. L.; Larsen, C. E.

    2011-01-01

    This paper describes the measurement and analysis of surface fluctuating pressure level (FPL) data and vibration data from a plume impingement aero-acoustic and vibration (PIAAV) test to validate NASA s physics-based modeling methods for prediction of panel vibration in the near field of a hot supersonic rocket plume. For this test - reported more fully in a companion paper by Osterholt & Knox at 26th Aerospace Testing Seminar, 2011 - the flexible panel was located 2.4 nozzle diameters from the plume centerline and 4.3 nozzle diameters downstream from the nozzle exit. The FPL loading is analyzed in terms of its auto spectrum, its cross spectrum, its spatial correlation parameters and its statistical properties. The panel vibration data is used to estimate the in-situ damping under plume FPL loading conditions and to validate both finite element analysis (FEA) and statistical energy analysis (SEA) methods for prediction of panel response. An assessment is also made of the effects of non-linearity in the panel elasticity.

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

    NASA Technical Reports Server (NTRS)

    Castner, Raymond

    2012-01-01

    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.

  5. Apollo 12 Lunar Module exhaust plume impingement on Lunar Surveyor III

    NASA Astrophysics Data System (ADS)

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

    2011-02-01

    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 μm (major diameter) × 74.5 μm (minor diameter) and the average estimated penetration depth was 88.4 μm. 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.

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

    NASA Technical Reports Server (NTRS)

    Bruckner, A. P.

    1972-01-01

    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.

  7. 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)

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

    2002-04-01

    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 ɛΔCO2 = 3.3 +/- 1.8% and ɛΔT = 2.9 +/- 1.6% for fuel sulfur conversion to sulfuric acid were inferred when relating the H2SO4 data to measurements of the plume tracers ΔCO2 and ΔT.

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

    NASA Astrophysics Data System (ADS)

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

    2011-05-01

    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.

  9. High altitude chemically reacting gas particle mixtures. Volume 2: Program manual for RAMP2. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    All of the elements used in the Reacting and Multi-Phase (RAMP2) computer code are described in detail. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields.

  10. Plume Mitigation: Soil Erosion and Lunar Prospecting Sensor Project

    NASA Technical Reports Server (NTRS)

    Metzger, Philip T.

    2014-01-01

    Demonstrate feasibility of the simplest, lowest-mass method of measuring density of a cloud of lunar soil ejected by rocket exhaust, using new math techniques with a small baseline laser/camera system. Focus is on exploring the erosion process that occurs when the exhaust plume of a lunar rocket impacts the regolith. Also, predicting the behavior of the lunar soil that would be blasted from a lunar landing/launch site shall assist in better design and protection of any future lunar settlement from scouring of structures and equipment. NASA is gathering experimental data to improve soil erosion models and understand how lunar particles enter the plume flow.

  11. Plume mass flow and optical damage distributions for an MMH/N2O4 RCS thruster. [exhaust plume contamination of spacecraft components

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    The data obtained from two recent experiments conducted in a continuing series of experiments at the Lewis Research Center into the contamination characteristics of a 5-pound thrust MMH/N2O4 engine are presented. The primary objectives of these experiments were to establish the angular distribution of condensible exhaust products within the plume and the corresponding optical damage angular distribution of transmitting optical elements attributable to this contaminant. The plume mass flow distribution was measured by five quartz crystal microbalances (QCM's) located at the engine axis evaluation. The fifth QCM was located above the engine and 15 deg behind the nozzle exit plane. The optical damage was determined by ex-situ transmittance measurements for the wavelength range from 0.2 to 0.6 microns on 2.54 cm diameter fused silica discs also located at engine centerline elevation. Both the mass deposition and optical damage angular distributions followed the expected trend of decreasing deposition and damage as the angle between sensor or sample and the nozzle axis increased. A simple plume gas flow equation predicted the deposition distribution reasonably well for angles of up to 55 degrees. The optical damage measurements also indicated significant effects at large angles.

  12. A computer program for thermal radiation from gaseous rocket exhuast plumes (GASRAD)

    NASA Technical Reports Server (NTRS)

    Reardon, J. E.; Lee, Y. C.

    1979-01-01

    A computer code is presented for predicting incident thermal radiation from defined plume gas properties in either axisymmetric or cylindrical coordinate systems. The radiation model is a statistical band model for exponential line strength distribution with Lorentz/Doppler line shapes for 5 gaseous species (H2O, CO2, CO, HCl and HF) and an appoximate (non-scattering) treatment of carbon particles. The Curtis-Godson approximation is used for inhomogeneous gases, but a subroutine is available for using Young's intuitive derivative method for H2O with Lorentz line shape and exponentially-tailed-inverse line strength distribution. The geometry model provides integration over a hemisphere with up to 6 individually oriented identical axisymmetric plumes, a single 3-D plume, Shading surfaces may be used in any of 7 shapes, and a conical limit may be defined for the plume to set individual line-of-signt limits. Intermediate coordinate systems may specified to simplify input of plumes and shading surfaces.

  13. Rocket noise - A review

    NASA Astrophysics Data System (ADS)

    McInerny, S. A.

    1990-10-01

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

  14. Survey of recent Al2O3 droplet size data in solid rocket chambers, nozzles, and plumes

    NASA Astrophysics Data System (ADS)

    Salita, Mark

    1994-10-01

    The size distribution of Al2O3 droplets in a solid propellant rocket is very different in the chamber, nozzle, and plume, primarily due to breakup and collision/coalescence in the nozzle. This paper attempts to summarize, evaluate, and relate the results of 19 recent experimental and analytical studies of droplet size in order to construct a unified model of droplet size evolution from the chamber through the nozzle and into the plume. It is believed that the distribution of droplet mass in the chamber is bimodal lognormal, with 1 micrometer smoke comprising about 80 percent of the mass and 50-100 micrometer caps comprising the remaining 20 percent. During passage through the nozzle, the caps shatter to product droplets whose diameters are about 10 percent of the caps but still 10 times those of smoke, while all but 1-2 percent of the smoke mass collides and coalesces with the shatter products to generate an essentially monomodal mass distribution at the nozzle exit whose D(sub 43) agrees with Hermsen's correlation and whose standard deviation sigma approximately equals 0.13 is smaller than both the chamber smoke (sigma approximately equals 0.40) or caps (sigma approximately equals 0.20).

  15. Survey of recent Al2O3 droplet size data in solid rocket chambers, nozzles, and plumes

    NASA Astrophysics Data System (ADS)

    Salita, Mark

    1994-10-01

    The size distribution of Al2O3 droplets in a solid propellant rocket is very different in the chamber, nozzle, and plume, primarily due to breakup and collision/coalescence in the nozzle. This paper attempts to summarize, evaluate, and relate the results of 19 recent experimental and analytical studies of droplet size in order to construct a unified model of droplet size evolution from the chamber through the nozzle and into the plume. It is believed that the distribution of droplet mass in the chamber is bimodal log-normal, with 1 micron smoke comprising about 80% of the mass and 50-100 micron caps comprising the remaining 20%. During passage through the nozzle, the caps shatter to product droplets whose diameters are about 10% of the caps but still 10 times those of smoke, while all but 1-2% of the smoke mass collides and coalesces with the shatter products to generate an essentially monomodal mass distribution at the nozzle exit whose D43 agrees with Hermsen's correlation and whose standard deviation alpha approximately or equal to 0.13 is smaller than both the chamber smoke (alpha approximately or equal to 0.40) or caps (alpha approximately or equal to 0.20).

  16. Modeling of Heat Transfer and Ablation of Refractory Material Due to Rocket Plume Impingement

    NASA Technical Reports Server (NTRS)

    Harris, Michael F.; Vu, Bruce T.

    2012-01-01

    CR Tech's Thermal Desktop-SINDA/FLUINT software was used in the thermal analysis of a flame deflector design for Launch Complex 39B at Kennedy Space Center, Florida. The analysis of the flame deflector takes into account heat transfer due to plume impingement from expected vehicles to be launched at KSC. The heat flux from the plume was computed using computational fluid dynamics provided by Ames Research Center in Moffet Field, California. The results from the CFD solutions were mapped onto a 3-D Thermal Desktop model of the flame deflector using the boundary condition mapping capabilities in Thermal Desktop. The ablation subroutine in SINDA/FLUINT was then used to model the ablation of the refractory material.

  17. Digital filtering of plume emission spectra

    NASA Technical Reports Server (NTRS)

    Madzsar, George C.

    1990-01-01

    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.

  18. The chemistry and diffusion of aircraft exhausts in the lower stratosphere during the first few hours after fly-by. [with attention to ozone depletion by SST exhaust plumes

    NASA Technical Reports Server (NTRS)

    Hilst, G. R.

    1974-01-01

    An analysis of the hydrogen-nitrogen-oxygen reaction systems in the lower stratosphere as they are initially perturbed by individual aircraft engine exhaust plumes was conducted in order to determine whether any significant chemical reactions occur, either among exhaust chemical species, or between these species and the environmental ozone, while the exhaust products are confined to intact plume segments at relatively high concentrations. The joint effects of diffusive mixing and chemical kinetics on the reactions were also studied, using the techniques of second-order closure diffusion/chemistry models. The focus of the study was on the larger problem of the potential depletion of ozone by supersonic transport aircraft exhaust materials emitted into the lower stratosphere.

  19. Prediction of space shuttle fluctuating pressure environments, including rocket plume effects

    NASA Technical Reports Server (NTRS)

    Plotkin, K. J.; Robertson, J. E.

    1973-01-01

    Preliminary estimates of space shuttle fluctuating pressure environments have been made based on prediction techniques developed by Wyle Laboratories. Particular emphasis has been given to the transonic speed regime during launch of a parallel-burn space shuttle configuration. A baseline configuration consisting of a lightweight orbiter and monolithic SRB, together with a typical flight trajectory, have been used as models for the predictions. Critical fluctuating pressure environments are predicted at transonic Mach numbers. Comparisons between predicted environments and wind tunnel test results, in general, showed good agreement. Predicted one-third octave band spectra for the above environments were generally one of three types: (1) attached turbulent boundary layer spectra (typically high frequencies); (2) homogeneous separated flow and shock-free interference flow spectra (typically intermediate frequencies); and (3) shock-oscillation and shock-induced interference flow spectra (typically low frequencies). Predictions of plume induced separated flow environments were made. Only the SRB plumes are important, with fluctuating levels comparable to compression-corner induced separated flow shock oscillation.

  20. High altitude chemically reacting gas particle mixtures. Volume 1: A theoretical analysis and development of the numerical solution. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    The overall contractual effort and the theory and numerical solution for the Reacting and Multi-Phase (RAMP2) computer code are described. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. Fundamental equations for steady flow of reacting gas-particle mixtures, method of characteristics, mesh point construction, and numerical integration of the conservation equations are considered herein.

  1. One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

    NASA Technical Reports Server (NTRS)

    Sakowski, Barbara; Edwards, Daryl; Dickens, Kevin

    2014-01-01

    Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation [Ref 1]. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet

  2. One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

    NASA Technical Reports Server (NTRS)

    Sakowski, Barbara A.; Edwards, Daryl; Dickens, Kevin

    2014-01-01

    Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet conduction as

  3. Ionospheric shock waves triggered by rockets

    NASA Astrophysics Data System (ADS)

    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

    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.

  4. Equations of motion for the variable mass flow-variable exhaust velocity rocket

    NASA Technical Reports Server (NTRS)

    Tempelman, W. H.

    1972-01-01

    An equation of motion for a one dimensional rocket is derived as a function of the mass flow rate into the acceleration chamber and the velocity distribution along the chamber, thereby including the transient flow changes in the chamber. The derivation of the mass density requires the introduction of the special time coordinate. The equation of motion is derived from both classical force and momentum approaches and is shown to be consistent with the standard equation expressed in terms of flow parameters at the exit to the acceleration chamber.

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

    NASA Technical Reports Server (NTRS)

    Castner, Raynold S.

    2010-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

    Castner, Raymond S.

    2009-01-01

    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.

  7. Velocity and temperature characteristics of two-stream, coplanar jet exhaust plumes

    NASA Technical Reports Server (NTRS)

    Von Glahn, U.; Goodykoontz, J.; Wasserbauer, C.

    1984-01-01

    The subsonic jet exhaust velocity and temperature characteristics of model scale, two stream coplanar nozzles were obtained experimentally. The data obtained included the effects of fan to primary stream velocity and temperature ratios on the jet axial and radial flow characteristics. Empirical parameters were developed to correlate the measured data. The resultant equations were shown to be extensions of a previously published single stream jet velocity and temperature correlation.

  8. Size distribution of unburned aluminum particles in solid propellant rocket motor exhaust

    SciTech Connect

    Larson, R.S.

    1986-06-01

    The size distribution of particles of unburned aluminum exiting a solid propellant rocket chamber is calculated by extending a previously developed theoretical model. Both one-dimensional and two-dimensional approximations to the chamber flow field are considered, but particle velocity lags are neglected. Results of the one-dimensional analysis differ from the more realistic two-dimensional results in that they predict a lower overall combustion efficiency and a most probable particle size which is always greater than zero. It is argued that these observations can be explained by the fact that the one-dimensional flow field allows many particles to pass through the chamber with a very short residence time.

  9. Hydrocarbon-Fueled Rocket Plume Measurement Using Polarized UV Raman Spectroscopy

    NASA Technical Reports Server (NTRS)

    Wehrmeyer, Joseph A.

    2002-01-01

    The influence of pressure upon the signal strength and polarization properties of UV Raman signals has been investigated experimentally up to pressures of 165 psia (11 atm). No significant influence of pressure upon the Raman scattering cross section or depolarization ratio of the N2 Raman signal was found. The Raman scattering signal varied linearly with pressure for the 300 K N2 samples examined, thus showing no enhancement of cross section with increasing pressure. However at the highest pressures associated with rocket engine combustion, there could be an increase in the Raman scattering cross section, based upon others' previous work at higher pressures than those examined in this work. The influence of pressure upon thick fused silica windows, used in the NASA Modular Combustion Test Article, was also investigated. No change in the transmission characteristics of the windows occurred as the pressure difference across the windows increased from 0 psig up to 150 psig. A calibration was performed on the UV Raman system at Vanderbilt University, which is similar to the one at the NASA-Marshall Test Stand 115. The results of this calibration are described in the form of temperature-dependent functions, f(T)'s, that account for the increase in Raman scattering cross section with an increase in temperature and also account for the reduction in collected Raman signal if wavelength integration does not occur across the entire wavelength range of the Raman signal. These functions generally vary only by approximately 10% across their respective temperature ranges, except for the case Of CO2, where there is a factor of three difference in its f(T) from 300 K to 2500 K. However this trend for CO2 is consistent with the experimental work of others, and is expected based on the low characteristic vibrational temperature Of CO2. A time-averaged temperature measurement technique has been developed, using the same equipment as for the work mentioned above, that is based upon

  10. Recommended launch-hold criteria for protecting public health from hydrogen chloride (HC1) gas produced by rocket exhaust

    SciTech Connect

    Daniels, J.I.; Baskett, R.L.

    1995-11-01

    Solid-fuel rocket motors used by the United States Air Force (USAF) to launch missiles and spacecraft can produce ambient-air concentrations of hydrogen chloride (HCI) gas. The HCI gas is a reaction product exhausted from the rocket motor during normal launch or emitted as a result of a catastrophic abort destroying the launch vehicle. Depending on the concentration in ambient air, the HCI gas can be irritating or toxic to humans. The diagnostic and complex-terrain wind field and particle dispersion model used by the Lawrence Livermore National Laboratory`s (LLNL`s) Atmospheric Release Advisory Capability (ARAC) Program was applied to the launch of a Peacekeeper missile from Vandenberg Air Force Base (VAFB) in California. Results from this deterministic model revealed that under specific meteorological conditions, cloud passage from normal-launch and catastropic-abort situations can yield measureable ground-level air concentrations of HCI where the general public is located. To protect public health in the event of such cloud passage, scientifically defensible, emergency ambient-air concentration limits for HCI were developed and recommended to the USAF for use as launch-hold criteria. Such launch-hold criteria are used to postpone a launch unless the forecasted meteorological conditions favor the prediction of safe ground-level concentrations of HCl for the general public. The recommended concentration limits are a 2 ppM 1-h time-weighted average (TWA) concentration constrained by a 1-min 10-ppM average concentration. This recommended criteria is supported by human dose-response information, including data for sensitive humans (e.g., asthmatics), and the dose response exhibited experimentally by animal models with respiratory physiology or responses considered similar to humans.

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

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    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.

  12. Characteristics of aerosol particles and trace gases in ship exhaust plumes

    NASA Astrophysics Data System (ADS)

    Drewnick, F.; Diesch, J.; Borrmann, S.

    2011-12-01

    Gaseous and particulate matter from marine vessels gain increasing attention due to their significant contribution to the anthropogenic burden of the atmosphere, implying the change of the atmospheric composition and the impact on local and regional air quality and climate (Eyring et al., 2010). As ship emissions significantly affect air quality of onshore regions, this study deals with various aspects of gas and particulate plumes from marine traffic measured near the Elbe river mouth in northern Germany. In addition to a detailed investigation of the chemical and physical particle properties from different types of commercial marine vessels, we will focus on the chemistry of ship plumes and their changes while undergoing atmospheric processing. Measurements of the ambient aerosol, various trace gases and meteorological parameters using a mobile laboratory (MoLa) were performed on the banks of the Lower Elbe which is passed on average, daily by 30 ocean-going vessels reaching the port of Hamburg, the second largest freight port of Europe. During 5 days of sampling from April 25-30, 2011 170 commercial marine vessels were probed at a distance of about 1.5-2 km with high temporal resolution. Mass concentrations in PM1, PM2.5 and PM10 and number as well as PAH and black carbon (BC) concentrations in PM1 were measured; size distribution instruments covered the size range from 6 nm up to 32 μm. The chemical composition of the non-refractory aerosol in the submicron range was measured by means of an Aerosol Mass Spectrometer (Aerodyne HR-ToF-AMS). Gas phase species analyzers monitored various trace gas concentrations in the air and a weather station provided meteorological parameters. Additionally, a wide spectrum of ship information for each vessel including speed, size, vessel type, fuel type, gross tonnage and engine power was recorded via Automatic Identification System (AIS) broadcasts. Although commercial marine vessels powered by diesel engines consume high

  13. Modeling a VASIMR rocket at UVSC

    NASA Astrophysics Data System (ADS)

    Matheson, Phil; Gray, William; Page, Leland

    2004-10-01

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1960-01-01

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

  15. 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)

    Smith, S. D.

    1984-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

    Dougherty, N. Sam; Liu, Baw-Lin

    1991-01-01

    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.

  17. 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

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

    1991-12-09

    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.

  18. Another Look at Rocket Thrust

    ERIC Educational Resources Information Center

    Hester, Brooke; Burris, Jennifer

    2012-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Bowyer, J. M.

    1977-01-01

    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.

  20. Plume Mitigation for Mars Terminal Landing: Soil Stabilization Project

    NASA Technical Reports Server (NTRS)

    Hintze, Paul E.

    2014-01-01

    Kennedy Space Center (KSC) has led the efforts for lunar and Martian landing site preparation, including excavation, soil stabilization, and plume damage prediction. There has been much discussion of sintering but until our team recently demonstrated it for the lunar case there was little understanding of the serious challenges. Simplistic sintering creates a crumbly, brittle, weak surface unsuitable for a rocket exhaust plume. The goal of this project is to solve those problems and make it possible to land a human class lander on Mars, making terminal landing of humans on Mars possible for the first time.

  1. 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)

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

    1975-01-01

    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.

  2. Research in the use of electrets in measuring effluents from rocket exhaust of the space shuttle (6.4 percent scaled model) and Viking 1 launch

    NASA Technical Reports Server (NTRS)

    Susko, M.

    1977-01-01

    Electrets used to detect the chemical composition of rocket exhaust effluents were investigated. The effectiveness of electrets was assessed while comparisons were made with hydrogen chloride measuring devices from chamber and field tests, and computed results from a multilayer diffusion model. The experimental data used were obtained from 18 static test firings, chamber tests, and the Viking 1 launch to Mars. Results show that electrets have multipollutant measuring capabilities, simplicity of deployment, and speed of assessment. The electrets compared favorably with other hydrogen chloride measuring devices. The summary of the measured data from the electrets and the hydrogen chloride detectors was within the upper and lower bounds of the computed hydrogen chloride concentrations from the multilayer diffusion model.

  3. Rocket pollution reduction system

    SciTech Connect

    Geisler, R.L.

    1994-01-04

    A system is provided for reducing the emissions of hydrochloric acid (HCl) from solid fuel rockets, especially during ground disposal. An aqueous solution of an alkali metal hydroxide is injected as a mist into the rocket chamber as the rocket fuel is burned. The reaction of the alkali metal with hydrogen chloride (HCl) produces a salt and thereby minimizes the presence of hydrochloric acid in the rocket exhaust. An injected neutralizing material which reduces hydrochloric acid, but which produces less thrust than an equal weight of rocket fuel, can be injected into an operating rocket which carries a payload high above the earth, with the injected material being injected only while the rocket is at a lower altitude when hydrochloric acid is most undesirable. The injected material can be produced by a small auxiliary rocket device whose exhaust is delivered directly to the main rocket chamber, and with the exhaust of the auxiliary rocket device including a high proportion of magnesium to react with the hydrochloric acid with minimal degradation of rocket performance. 4 figs.

  4. Plume interference with space shuttle range safety signals

    NASA Technical Reports Server (NTRS)

    Boynton, F. P.; Rajaseknar, P. S.

    1979-01-01

    The computational procedure for signal propagation in the presence of an exhaust plume is presented. Comparisons with well-known analytic diffraction solutions indicate that accuracy suffers when mesh spacing is inadequate to resolve the first unobstructed Fresnel zone at the plume edge. Revisions to the procedure to improve its accuracy without requiring very large arrays are discussed. Comparisons to field measurements during a shuttle solid rocket motor (SRM) test firing suggest that the plume is sharper edged than one would expect on the basis of time averaged electron density calculations. The effects, both of revisions to the computational procedure and of allowing for a sharper plume edge, are to raise the signal level near tail aspect. The attenuation levels then predicted are still high enough to be of concern near SRM burnout for northerly launches of the space shuttle.

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

    NASA Technical Reports Server (NTRS)

    1990-01-01

    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.

  6. 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)

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

    1973-01-01

    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.

  7. Preliminary Analysis of the Effect of Flow Separation Due to Rocket Jet Pluming on Aircraft Dynamic Stability During Atmospheric Exit

    NASA Technical Reports Server (NTRS)

    Dryer, Murray; North, Warren J.

    1959-01-01

    A theoretical investigation was conducted to determine the effects of body boundary-layer separation resulting from a highly underexpanded jet on the dynamic stability of a typical rocket aircraft during an atmospheric exit trajectory. The particular flight condition studied on a digital computer for five degrees of freedom was at Mach 6.0 and 150,000 feet. In view of the unknown character of the separated flow field, two estimates of the pressures in the separated region were made to calculate the unbalanced forces and moments. These estimates, based on limited fundamental zero-angle-of-attack studies and observations, are believed to cover what may be the actual case. In addition to a fixed control case, two simulated pilot control inputs were studied: rate-limited and instantaneous responses. The resulting-motions with and without boundary-layer separation were compared for various initial conditions. The lower of the assumed misalinement forces and moments led to a situation whereby a slowly damped motion could be satisfactorily controlled with rate-limited control input. The higher assumption led to larger amplitude, divergent motions when the same control rates were used. These motions were damped only when the instantaneous control responses were assumed.

  8. Instrumentation of UALR labscale hybrid rocket motor

    NASA Astrophysics Data System (ADS)

    Wright, Andrew B.; Teague, Warfield; Wright, Ann M.; Wilson, Edmond W.

    2006-05-01

    The Central Arkansas Combustion Group has used a NASA EPSCoR grant to improve the instrumentation and control of its labscale hybrid rocket facility. The research group investigates fundamental aspects of combustion in hybrid rocket motors. This paper describes the new instrumentation, provides examples of measurements taken, and describes novel instrumentation which is in the process of development. A six degree-of-freedom thrust system measures the total work done during a burn to compare the efficiency of fuels and fuel additives. The new system measures the forces and moments in three spatial dimensions. An accurate measure of thrust oscillations will lead to better understanding of the cause and eventual minimization of the oscillations. Plume spectrometers are employed to determine and measure the reaction intermediates and products of combustion at the exhaust. The new control system features an oxygen mass flow controller, which allows the accurate measurement of the oxidant introduced into the motor.

  9. Results of the NASA/MSFC FA-23 plume technology test program performed in the NASA/Ames unitary wind tunnels

    NASA Technical Reports Server (NTRS)

    Hendershot, K. C.

    1977-01-01

    A 2.25% scale model of the space shuttle external tank and solid rocket boosters was tested in the NASA/Ames Unitary 11 x 11 foot transonic and 9 x 7 foot supersonic tunnels to obtain base pressure data with firing solid propellant exhaust plumes. Data system difficulties prevented the acquisition of any useful data in the 9 x 7 tunnel. However, 28 successful rocket test firings were made in the 11 x 11 tunnel, providing base pressure data at Mach numbers of 0.5, 0.9, 1.05, 1.2, and 1.3 and at plume pressure ratios ranging from 11 to 89.

  10. Shock propagation in the exhaust gas handling system of the proposed large altitude rocket cell: methods and preliminary analysis

    SciTech Connect

    Sutton, S.B.; Pierce, R.E.

    1984-10-04

    Numerical predictions are to be performed of the shock pressures that would result from the detonation of 100,000 lbm TNT. The initial phase of the project was to develop the methodology for analyzing the problem, develop a preliminary conceptual design to use in initial simulations, and estimate over-pressures, inside the conceptual facility, resulting from the propellant detonation. This report discusses the methods of analysis used to study the problem of the detonation of the propellant, and the propagation of the shock wave inside the exhaust gas processing system, and presents preliminary results. The KOVEC computer code was used to simulate the detonation of 100,000 lbm TNT and develop a boundary prescription for use in the gas dynamics code GASP which models the propagation of the shock wave through the LARC exhaust gas processing system. Results are presented showing the effect of cross-sectional area changes and variations in the initial pressure in the gas processing system on the shock wave peak pressure and propagation speed.

  11. Ground and Space-Based Measurement of Rocket Engine Burns in the Ionosphere

    NASA Technical Reports Server (NTRS)

    Bernhardt, P. A.; Ballenthin, J. O.; Baumgardner, J. L.; Bhatt, A.; Boyd, I. D.; Burt, J. M.; Caton, R. G.; Coster, A.; Erickson, P. J.; Huba, J. D.; Earle, G. D.; Kaplan, C. R.; Foster, J. C.; Groves, K. M.; Haaser, R. A.; Heelis, R. A.; Hunton, D. E.; Hysell, D. L.; Klenzing, J. H.; Larsen, M. F.; Lind, F. D.; Pedersen, T. R.; Pfaff, R. F.; Stoneback, R. A.; Roddy, P. A.; Rodriguez, S. P.; San Antonio, G. S.; Schuck, P. W.; Siefring, C. L.; Selcher, C. A.; Smith, S. M.; Talaat, E. R.; Thomason, J. F.; Tsunoda, R. T.; Varney, R. H.

    2013-01-01

    On-orbit firings of both liquid and solid rocket motors provide localized disturbances to the plasma in the upper atmosphere. Large amounts of energy are deposited to ionosphere in the form of expanding exhaust vapors which change the composition and flow velocity. Charge exchange between the neutral exhaust molecules and the background ions (mainly O+) yields energetic ion beams. The rapidly moving pickup ions excite plasma instabilities and yield optical emissions after dissociative recombination with ambient electrons. Line-of-sight techniques for remote measurements rocket burn effects include direct observation of plume optical emissions with ground and satellite cameras, and plume scatter with UHF and higher frequency radars. Long range detection with HF radars is possible if the burns occur in the dense part of the ionosphere. The exhaust vapors initiate plasma turbulence in the ionosphere that can scatter HF radar waves launched from ground transmitters. Solid rocket motors provide particulates that become charged in the ionosphere and may excite dusty plasma instabilities. Hypersonic exhaust flow impacting the ionospheric plasma launches a low-frequency, electromagnetic pulse that is detectable using satellites with electric field booms. If the exhaust cloud itself passes over a satellite, in situ detectors measure increased ion-acoustic wave turbulence, enhanced neutral and plasma densities, elevated ion temperatures, and magnetic field perturbations. All of these techniques can be used for long range observations of plumes in the ionosphere. To demonstrate such long range measurements, several experiments were conducted by the Naval Research Laboratory including the Charged Aerosol Release Experiment, the Shuttle Ionospheric Modification with Pulsed Localized Exhaust experiments, and the Shuttle Exhaust Ionospheric Turbulence Experiments.

  12. Laboratory studies of the stratospheric effects of rocket exhaust. Final technical report, 1 Jan 1996--31 December 1998

    SciTech Connect

    Molina, M.J.; Haider, M.; Mantz, Y.; Gutzwiller, L.; Molina, L.T.

    1998-12-01

    The most important chlorine activation reaction that takes place in the stratosphere is the heterogeneous reaction of chlorine nitrate with hydrogen chloride to produce molecular chlorine and nitric acid. The reaction is catalytic, promoted by surfaces that are not themselves affected by the reaction. This process was investigated in the laboratory. The reaction probability was found to have a value of about 0.02 on glass, on laboratory alumina, and on alumina particles emitted by solid rocket motors. The reaction probability on ice surfaces has a value larger than 0.2 and a negligible value on halocarbon wax or Teflon surfaces. The measurements were carried out under reactant partial pressure, temperature and humidity conditions covering those that are encountered in the mid-latitude lower stratosphere. The reaction mechanism appears do be determined by the water layers adsorbed on the solid surface. Measurements of the amount of water taken up by alumina surfaces were also carried out; the results indicate that under stratospheric conditions several water layers will indeed cover the alumina surfaces.

  13. 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)

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

    2005-01-01

    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.

  14. Space Shuttle Plume Simulation Effect on Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hair, L. M.

    1978-01-01

    Technology for simulating plumes in wind tunnel tests was not adequate to provide the required confidence in test data where plume induced aerodynamic effects might be significant. A broad research program was undertaken to correct the deficiency. Four tasks within the program are reported. Three of these tasks involve conducting experiments, related to three different aspects of the plume simulation problem: (1) base pressures; (2) lateral jet pressures; and (3) plume parameters. The fourth task involves collecting all of the base pressure test data generated during the program. Base pressures were measured on a classic cone ogive cylinder body as affected by the coaxial, high temperature exhaust plumes of a variety of solid propellant rockets. Valid data were obtained at supersonic freestream conditions but not at transonic. Pressure data related to lateral (separation) jets at M infinity = 4.5, for multiple clustered nozzles canted to the freestream and operating at high dynamic pressure ratios. All program goals were met although the model hardware was found to be large relative to the wind tunnel size so that operation was limited for some nozzle configurations.

  15. 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)

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

    1978-01-01

    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.

  16. Site Alteration Effects from Rocket Exhaust Impingement During a Simulated Viking Mars Landing. Part 2: Chemical and Biological Site Alteration

    NASA Technical Reports Server (NTRS)

    Husted, R. R.; Smith, I. D.; Fennessey, P. V.

    1977-01-01

    Chemical and biological alteration of a Mars landing site was investigated experimentally and analytically. The experimental testing was conducted using a specially designed multiple nozzle configuration consisting of 18 small bell nozzles. The chemical test results indicate that an engine using standard hydrazine fuel will contaminate the landing site with ammonia (50-500ppm), nitrogen (5-50ppm), aniline (0.01-0.5ppm), hydrogen cyanide (0.01-0.5ppm), and water. A purified fuel, with impurities (mostly aniline) reduced by a factor of 50-100, limits the amount of hydrogen cyanide and aniline to below detectable limits for the Viking science investigations and leaves the amounts of ammonia, nitrogen, and water in the soil unchanged. The large amounts of ammonia trapped in the soil will make interpretation of the organic analysis investigation results more difficult. The biological tests indicate that the combined effects of plume gases, surface heating, surface erosion, and gas composition resulting from the retrorockets will not interfere with the Viking biology investigation.

  17. An analytical and experimental investigation of resistojet plumes

    NASA Technical Reports Server (NTRS)

    Zana, Lynnette M.; Hoffman, David J.; Breyley, Loranell R.; Serafini, John S.

    1987-01-01

    As a part of the electrothermal propulsion plume research program at the NASA Lewis Research Center, efforts have been initiated to analytically and experimentally investigate the plumes of resistojet thrusters. The method of Simons for the prediction of rocket exhaust plumes is developed for the resistojet. Modifications are made to the source flow equations to account for the increased effects of the relatively large nozzle boundary layer. Additionally, preliminary mass flux measurements of a laboratory resistojet using CO2 propellant at 298 K have been obtained with a cryogenically cooled quartz crystal microbalance (QCM). There is qualitative agreement between analysis and experiment, at least in terms of the overall number density shape functions in the forward flux region.

  18. An analytical and experimental investigation of resistojet plumes

    NASA Technical Reports Server (NTRS)

    Zana, L. M.; Hoffman, D. J.; Breyley, L. R.; Serafini, J. S.

    1987-01-01

    As a part of the electrothermal propulsion plume research program at the NASA Lewis Research Center, efforts have been initiated to analytically and experimentally investigate the plumes of resistojet thrusters. The method of G.A. Simons for the prediction of rocket exhaust plumes is developed for the resistojet. Modifications are made to the source flow equations to account for the increased effects of the relatively large nozzle boundary layer. Additionally, preliminary mass flux measurements of a laboratory resistojet using CO2 propellant at 298 K have been obtained with a cryogenically cooled quartz crystal microbalance (QCM). There is qualitative agreement between analysis and experiment, at least in terms of the overall number density shape functions in the forward flux region.

  19. Pressure Loads Produced on a Flat-Plate Wing By Rocket Jets Exhausting in a Spanwise Direction Below the Wing and Perpendicular to a Free-Stream Flow of Mach Number 2.0

    NASA Technical Reports Server (NTRS)

    Falanga, Ralph A.; Janos, Joseph J.

    1961-01-01

    An investigation at a Reynolds number per foot of 14.4 x 10(exp 6) was made to determine the pressure loads produced on a flat-plate wing by rocket jets exhausting in a spanwise direction beneath the wing and perpendicular to a free-stream flow of Mach number 2.0. The ranges of the variables involved were (1) nozzle types - one sonic (jet Mach number of 1.00), two supersonic (jet Mach numbers of 1.74 and 3.04),. and one two-dimensional supersonic (jet Mach number of 1.71); (2) vertical nozzle positions beneath the wing of 4, 8 and 12 nozzle-throat diameters; and (3) ratios of rocket-chamber total pressure to free- stream static pressure from 0 to 130. The incremental normal force due to jet interference on the wing varied from one to two times the rocket thrust and generally decreased as the pressure ratio increased. The chordwise coordinate of the incremental-normal-force center of pressure remained upstream of the nozzle center line for the nozzle positions and pressure ratios of the investigation. The chordwise coordinate approached zero as the jet vertical distance beneath the wing increased. In the spanwise direction there was little change due to varying rocket-jet position and pressure ratio. Some boundary-layer flow separation on the wing was observed for the rocket jets close to the wing and at the higher pressure ratios. The magnitude of the chordwise and spanwise pressure distributions due to jet interference was greatest for rocket jets close to the wing and decreased as the jet was displaced farther from the wing. The design procedure for the rockets used is given in the appendix.

  20. Scaled Rocket Testing in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  1. Towards Understanding the Fluid Dynamic Phenomenon of Interest to Rocket Base Heating: A Review

    NASA Technical Reports Server (NTRS)

    Venkatapathy, E.; Park, C.; Palmer, G.; Arnold, James O. (Technical Monitor)

    1994-01-01

    The significance of the base heating problem for rockets during ascent is due to the complex interaction between the rocket nozzle plumes and the external-flow which can change the flow characteristics in the base region dramatically. At lower altitudes the external-flow merges with the plume-flow, without the formation of a large separated flow region, and the cooler external-flow promotes convective cooling of the base wall. Under these conditions the majority of the base heating is due to radiative heating from the shock heated plume gases. At higher altitudes, however, the process of base heating is not so straightforward. The plume and the base flow expands dramatically and separated flow regions occur in the base area. Hot exhaust gases from the rocket nozzle will be entrained into the separated flow regions and produce a convective component to the base wall heating. Further, if the rocket exhaust-gas contains soot, the soot can increase the emission from the gas and dramatically increase the wall absorption coefficient for radiative heating if it is deposited on the walls . In addition, if the rocket exhaust gas is fuel rich, the fuel can bum in the separated flow regions and further increase the base heating. The base burning phenomenon, and the increased base heating caused by it at higher altitudes, have been observed for the Space Shuttle and Saturn Rocket. Under these conditions, the total heating is significantly higher than the heating without separated flow in the base region, and the increase in heating is directly attributable to the fluid dynamic complexity of the base region. Realistic simulation of the base heating requires that the calculated flow environment reproduce the fluid dynamic flow features accurately. Thus, it will be necessary to introduce into the CFD codes the capability for the flow to respond to the complex vehicle geometry, the effect of turbulence, the ability to accurately reproduce the plume shock/shear layer structures and

  2. Coupled simulation of CFD-flight-mechanics with a two-species-gas-model for the hot rocket staging

    NASA Astrophysics Data System (ADS)

    Li, Yi; Reimann, Bodo; Eggers, Thino

    2016-11-01

    The hot rocket staging is to separate the lowest stage by directly ignite the continuing-stage-motor. During the hot staging, the rocket stages move in a harsh dynamic environment. In this work, the hot staging dynamics of a multistage rocket is studied using the coupled simulation of Computational Fluid Dynamics and Flight Mechanics. Plume modeling is crucial for a coupled simulation with high fidelity. A 2-species-gas model is proposed to simulate the flow system of the rocket during the staging: the free-stream is modeled as "cold air" and the exhausted plume from the continuing-stage-motor is modeled with an equivalent calorically-perfect-gas that approximates the properties of the plume at the nozzle exit. This gas model can well comprise between the computation accuracy and efficiency. In the coupled simulations, the Navier-Stokes equations are time-accurately solved in moving system, with which the Flight Mechanics equations can be fully coupled. The Chimera mesh technique is utilized to deal with the relative motions of the separated stages. A few representative staging cases with different initial flight conditions of the rocket are studied with the coupled simulation. The torque led by the plume-induced-flow-separation at the aft-wall of the continuing-stage is captured during the staging, which can assist the design of the controller of the rocket. With the increasing of the initial angle-of-attack of the rocket, the staging quality becomes evidently poorer, but the separated stages are generally stable when the initial angle-of-attack of the rocket is small.

  3. Design and Testing of a Liquid Nitrous Oxide and Ethanol Fueled Rocket Engine

    SciTech Connect

    Youngblood, Stewart

    2015-08-01

    A small-scale, bi-propellant, liquid fueled rocket engine and supporting test infrastructure were designed and constructed at the Energetic Materials Research and Testing Center (EMRTC). This facility was used to evaluate liquid nitrous oxide and ethanol as potential rocket propellants. Thrust and pressure measurements along with high-speed digital imaging of the rocket exhaust plume were made. This experimental data was used for validation of a computational model developed of the rocket engine tested. The developed computational model was utilized to analyze rocket engine performance across a range of operating pressures, fuel-oxidizer mixture ratios, and outlet nozzle configurations. A comparative study of the modeling of a liquid rocket engine was performed using NASA CEA and Cantera, an opensource equilibrium code capable of being interfaced with MATLAB. One goal of this modeling was to demonstrate the ability of Cantera to accurately model the basic chemical equilibrium, thermodynamics, and transport properties for varied fuel and oxidizer operating conditions. Once validated for basic equilibrium, an expanded MATLAB code, referencing Cantera, was advanced beyond CEAs capabilities to predict rocket engine performance as a function of supplied propellant flow rate and rocket engine nozzle dimensions. Cantera was found to comparable favorably to CEA for making equilibrium calculations, supporting its use as an alternative to CEA. The developed rocket engine performs as predicted, demonstrating the developedMATLAB rocket engine model was successful in predicting real world rocket engine performance. Finally, nitrous oxide and ethanol were shown to perform well as rocket propellants, with specific impulses experimentally recorded in the range of 250 to 260 seconds.

  4. Delta 2 Explosion Plume Analysis Report

    NASA Technical Reports Server (NTRS)

    Evans, Randolph J.

    2000-01-01

    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.

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

    NASA Astrophysics Data System (ADS)

    Toohey, D. W.; Ross, M.

    2009-12-01

    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

  6. Particle Rotation Effects in Rarefied Two-Phase Plume Flows

    NASA Astrophysics Data System (ADS)

    Burt, Jonathan M.; Boyd, Iain D.

    2005-05-01

    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.

  7. Linear Spectral Analysis of Plume Emissions Using an Optical Matrix Processor

    NASA Technical Reports Server (NTRS)

    Gary, C. K.

    1992-01-01

    Plume spectrometry provides a means to monitor the health of a burning rocket engine, and optical matrix processors provide a means to analyze the plume spectra in real time. By observing the spectrum of the exhaust plume of a rocket engine, researchers have detected anomalous behavior of the engine and have even determined the failure of some equipment before it would normally have been noticed. The spectrum of the plume is analyzed by isolating information in the spectrum about the various materials present to estimate what materials are being burned in the engine. Scientists at the Marshall Space Flight Center (MSFC) have implemented a high resolution spectrometer to discriminate the spectral peaks of the many species present in the plume. Researchers at the Stennis Space Center Demonstration Testbed Facility (DTF) have implemented a high resolution spectrometer observing a 1200-lb. thrust engine. At this facility, known concentrations of contaminants can be introduced into the burn, allowing for the confirmation of diagnostic algorithms. While the high resolution of the measured spectra has allowed greatly increased insight into the functioning of the engine, the large data flows generated limit the ability to perform real-time processing. The use of an optical matrix processor and the linear analysis technique described below may allow for the detailed real-time analysis of the engine's health. A small optical matrix processor can perform the required mathematical analysis both quicker and with less energy than a large electronic computer dedicated to the same spectral analysis routine.

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

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

    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.

  9. 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)

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

    1979-01-01

    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.

  10. Contrail formation in the tropopause region caused by emissions from an Ariane 5 rocket

    NASA Astrophysics Data System (ADS)

    Voigt, Ch.; Schumann, U.; Graf, K.

    2016-07-01

    Rockets directly inject water vapor and aerosol into the atmosphere, which promotes the formation of ice clouds in ice supersaturated layers of the atmosphere. Enhanced mesospheric cloud occurrence has frequently been detected near 80-kilometer altitude a few days after rocket launches. Here, unique evidence for cirrus formation in the tropopause region caused by ice nucleation in the exhaust plume from an Ariane 5-ECA rocket is presented. Meteorological reanalysis data from the European Centre for Medium-Range Weather Forecasts show significant ice supersaturation at the 100-hectopascal level in the American tropical tropopause region on November 26, 2011. Near 17-kilometer altitudes, the temperatures are below the Schmidt-Appleman threshold temperature for rocket condensation trail formation on that day. Immediately after the launch from the Ariane 5-ECA at 18:39 UT (universal time) from Kourou, French Guiana, the formation of a rocket contrail is detected in the high resolution visible channel from the SEVIRI (Spinning Enhanced Visible and InfraRed Imager) on the METEOSAT9 satellite. The rocket contrail is transported to the south and its dispersion is followed in SEVIRI data for almost 2 h. The ice crystals predominantly nucleated on aluminum oxide particles emitted by the Ariane 5-ECA solid booster and further grow by uptake of water vapor emitted from the cryogenic main stage and entrained from the ice supersaturated ambient atmosphere. After rocket launches, the formation of rocket contrails can be a frequent phenomenon under ice supersaturated conditions. However, at present launch rates, the global climate impact from rocket contrail cirrus in the tropopause region is small.

  11. Atmospheric scavenging exhaust

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. The airborne HCl concentration varied from 0.2 to 10.0 ppm and the raindrop sizes tested included 0.55 mm, 1.1 mm, and 3.0 mm. 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 employed. 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.

  12. Another Look at Rocket Thrust

    NASA Astrophysics Data System (ADS)

    Hester, Brooke; Burris, Jennifer

    2012-12-01

    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force—the thrust—on the rocket.1,2 Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of conservation of momentum is generally used in textbooks3,4 and elsewhere5 in deriving the rocket thrust equation. In this paper we take the somewhat different approach of explicitly applying Newton's second law to the rocket. This method provides a good opportunity to show the importance of choosing carefully the system to which Newton's second law is applied.

  13. 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)

    1989-01-01

    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.

  14. Liquid Booster Module (LBM) plume flowfield model

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1981-01-01

    A complete definition of the LBM plume is important for many Shuttle design criteria. The exhaust plume shape has a significant effect on the vehicle base pressure. The LBM definition is also important to the Shuttle base heating, aerodynamics and the influence of the exhaust plume on the launch stand and environment. For these reasons a knowledge of the LBM plume characteristics is necessary. A definition of the sea level LBM plume as well as at several points along the Shuttle trajectory to LBM, burnout is presented.

  15. Underexpanded Supersonic Plume Surface Interactions: Applications for Spacecraft Landings on Planetary Bodies

    NASA Technical Reports Server (NTRS)

    Mehta, M.; Sengupta, A.; Renno, N. O.; Norman, J. W.; Gulick, D. S.

    2011-01-01

    Numerical and experimental investigations of both far-field and near-field supersonic steady jet interactions with a flat surface at various atmospheric pressures are presented in this paper. These studies were done in assessing the landing hazards of both the NASA Mars Science Laboratory and Phoenix Mars spacecrafts. Temporal and spatial ground pressure measurements in conjunction with numerical solutions at altitudes of approx.35 nozzle exit diameters and jet expansion ratios (e) between 0.02 and 100 are used. Data from steady nitrogen jets are compared to both pulsed jets and rocket exhaust plumes at Mach approx.5. Due to engine cycling, overpressures and the plate shock dynamics are different between pulsed and steady supersonic impinging jets. In contrast to highly over-expanded (e <1) and underexpanded exhaust plumes, results show that there is a relative ground pressure load maximum for moderately underexpanded (e approx.2-5) jets which demonstrate a long collimated plume shock structure. For plumes with e much >5 (lunar atmospheric regime), the ground pressure is minimal due to the development of a highly expansive shock structure. We show this is dependent on the stability of the plate shock, the length of the supersonic core and plume decay due to shear layer instability which are all a function of the jet expansion ratio. Asymmetry and large gradients in the spatial ground pressure profile and large transient overpressures are predominantly linked to the dynamics of the plate shock. More importantly, this study shows that thruster plumes exhausting into martian environments possess the largest surface pressure loads and can occur at high spacecraft altitudes in contrast to the jet interactions at terrestrial and lunar atmospheres. Theoretical and analytical results also show that subscale supersonic cold gas jets adequately simulate the flow field and loads due to rocket plume impingement provided important scaling parameters are in agreement. These

  16. Measurements of temperature profiles at the exit of small rockets.

    PubMed

    Griggs, M; Harshbarger, F C

    1966-02-01

    The sodium line reversal technique was used to determine the reversal temperature profile across the exit of small rockets. Measurements were made on one 73-kg thrust rocket, and two 23-kg thrust rockets with different injectors. The large rocket showed little variation of reversal temperature across the plume. However, the 23-kg rockets both showed a large decrease of reversal temperature from the axis to the edge of the plume. In addition, the sodium line reversal technique of temperature measurement was compared with an infrared technique developed in these laboratories.

  17. Congreve Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The British fired Congreve rockets against the United States in the War of 1812. As a result Francis Scott Key coined the phrase the 'rocket's red glare.' Congreve had used a 16-foot guide stick to help stabilize his rocket. William Hale, another British inventor, invented the stickless rocket in 1846. The U.S. Army used the Hale rocket more than 100 years ago in the war with Mexico. Rockets were also used to a limited extent by both sides in the American Civil War.

  18. Safe testing nuclear rockets economically

    SciTech Connect

    Howe, S. D.; Travis, B. J.; Zerkle, D. K.

    2002-01-01

    Several studies over the past few decades have recognized the need for advanced propulsion to explore the solar system. As early as the 1960s, Werner Von Braun and others recognized the need for a nuclear rocket for sending humans to Mars. The great distances, the intense radiation levels, and the physiological response to zero-gravity all supported the concept of using a nuclear rocket to decrease mission time. These same needs have been recognized in later studies, especially in the Space Exploration Initiative in 1989. One of the key questions that has arisen in later studies, however, is the ability to test a nuclear rocket engine in the current societal environment. Unlike the 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.

  19. Rocket Flight.

    ERIC Educational Resources Information Center

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

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

  20. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  1. Altitude-Compensating Nozzle (ACN) Project: Planning for Dual-Bell Rocket Nozzle Flight Testing on the NASA F-15B

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

    For more than a half-century, several types of altitude-compensating nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Although the dual-bell rocket nozzle has been thoroughly studied, this nozzle has still not been tested in a relevant flight environment. This poster presents the top-level rationale and preliminary plans for conducting flight research with the dual-bell rocket nozzle, while exhausting the plume into the freestream flow field at various altitudes. The primary objective is to gain a greater understanding of the nozzle plume sensitivity to freestream flight effects, which will also include detailed measurements of the plume mode transition within the nozzle. To accomplish this goal, the NASA F-15B is proposed as the testbed for advancing the technology readiness level of this greatly-needed capability. All proposed tests include the quantitative performance analysis of the dual-bell rocket nozzle as compared with the conventional-bell nozzle.

  2. Prediction of Acoustic Environments from Horizontal Rocket Firings

    NASA Technical Reports Server (NTRS)

    Giacomoni, Clothilde

    2014-01-01

    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

  3. Carrier rockets

    NASA Astrophysics Data System (ADS)

    Aleksandrov, V. A.; Vladimirov, V. V.; Dmitriev, R. D.; Osipov, S. O.

    This book takes into consideration domestic and foreign developments related to launch vehicles. General information concerning launch vehicle systems is presented, taking into account details of rocket structure, basic design considerations, and a number of specific Soviet and American launch vehicles. The basic theory of reaction propulsion is discussed, giving attention to physical foundations, the various types of forces acting on a rocket in flight, basic parameters characterizing rocket motion, the effectiveness of various approaches to obtain the desired velocity, and rocket propellants. Basic questions concerning the classification of launch vehicles are considered along with construction and design considerations, aspects of vehicle control, reliability, construction technology, and details of structural design. Attention is also given to details of rocket motor design, the basic systems of the carrier rocket, and questions of carrier rocket development.

  4. The production of nitric oxide in the troposphere as a result of solid-rocket-motor afterburning

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    As part of an ongoing assessment of the environmental effects of solid-rocket-motor operations in the troposphere, estimates were made of the nitric oxide produced in the troposphere by the space shuttle and Titan 3-C boosters. Calculations were made with the low-altitude plume computer program and included the effects of coupled finite-rate chemistry and turbulent mixing. A recent measurement of nitric oxide taken in the effluent cloud of a Titan 3-C booster is compared with calculations made with this computer code. The various chemical reactions of the exhaust gases are listed in tabular form.

  5. Large-eddy simulations of a solid-rocket booster jet

    NASA Astrophysics Data System (ADS)

    Paoli, Roberto; Poubeau, Adele; Cariolle, Daniel

    2014-11-01

    Emissions from solid-rocket boosters are responsible for a severe decrease in ozone concentration in the rocket plume during the first hours after a launch. The main source of ozone depletion is due to hydrogen chloride that is converted into chlorine in the high temperature regions of the jet (afterburning). The objective of this study is to evaluate the active chlorine concentration in the plume of a solid-rocket booster using large-eddy simulations. The gas is injected through the entire nozzle of the booster and a local time-stepping method based on coupling multi-instances of a fluid solver is used to extend the computational domain up to 600 nozzle exit diameters. The methodology is validated for a non-reactive case by analyzing the flow characteristics of supersonic co-flowing under expanded jets. Then, the chemistry of chlorine is studied offline using a complex chemistry solver and the LES data extracted from the mean trajectories of sample fluid particles. Finally, the online chemistry is analyzed by means of the multispecies version of the LES solver using a reduced chemistry scheme. The LES are able to capture the mixing of the exhaust with ambient air and the species concentrations, which is also useful to initialize atmospheric simulations on larger domains.

  6. Orbital Maneuvering Vehicle (OMV) plume and plume effects study

    NASA Technical Reports Server (NTRS)

    Smith, Sheldon D.

    1991-01-01

    The objective was to characterize the Orbital Maneuvering Vehicle (OMV) propulsion and attitude control system engine exhaust plumes and predict the resultant plume impingement pressure, heat loads, forces, and moments. Detailed description is provided of the OMV gaseous nitrogen (GN2) thruster exhaust plume flow field characteristics calculated with the RAMP2 snd SFPGEN computer codes. Brief descriptions are included of the two models, GN2 thruster characteristics and RAMP2 input data files. The RAMP2 flow field could be recalculated by other organizations using the information presented. The GN2 flow field can be readily used by other organizations who are interested in GN2 plume induced environments which require local flow field properties which can be supplied using the SFPGEN GN2 model.

  7. Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode

    NASA Technical Reports Server (NTRS)

    Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.

    2000-01-01

    are being conducted at Marshall Space Flight Center to benchmark the FDNS code for RBCC engine operations for such configurations. The primary fluid physics of interests are the mixing and interaction of the rocket plume and secondary flow, subsequent combustion of the fuel rich rocket exhaust with the secondary flow and combustion of the injected afterburner flow. The CFD results are compared to static pressure along the RBCC duct walls, Raman Spectroscopy specie distribution data at several axial locations, net engine thrust and entrained air for the SLS cases. The CFD results compare reasonably well with the experimental results.

  8. Ionospheric modification by rocket effluents. Final report

    SciTech Connect

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

    1980-06-01

    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.

  9. POD Analysis of Jet-Plume/Afterbody-Wake Interaction

    NASA Astrophysics Data System (ADS)

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

    2009-11-01

    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.

  10. Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

    NASA Technical Reports Server (NTRS)

    Anderson, William E.

    2005-01-01

    The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

  11. Solid rocket booster thermal radiation model, volume 1

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    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.

  12. Methylhydrazinium nitrate. [rocket plume deposit chemistry

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

    Boyle, W.; Conine, B.

    1978-01-01

    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.

  14. Experimentation in the low-density plume of a simulated electrothermal thruster for computer code validation

    NASA Technical Reports Server (NTRS)

    Meissner, Dana L.

    1993-01-01

    Pressures and flow angles are measured in the plume of a 20 deg half-angle, conical nozzle in vacuum with Pitot tubes and conical probes. The area of measurement in the plume ranges from the nozzle exit plane to 480 mm axially downstream and from the plume centerline to 60 mm radially. The nozzle has an exit-to-throat area ratio of 100:1 and a throat diameter of 3.2 mm. The nozzle flow exhausts to a vacuum of order 10(exp -2) Pa to simulate a resistojet (an electrothermal rocket of less than 1 N of thrust) operating in space. Experimental data are given for flows of nitrogen at 55 and 68 mg/s, stagnation temperatures between 695 and 921 K, and stagnation pressures ranging from 5600 to 7100 Pa. Data are also given for argon at a rate of 68 mg/s, a stagnation temperature of 648 K, and stagnation pressures of 4500, 4750, and 4770 Pa. Measurements in the nitrogen plume are compared with computational results from a direct-simulation Monte Carlo method.

  15. Helping HAN for hybrid rockets

    SciTech Connect

    Ramohalli, K.; Dowler, W.

    1995-01-01

    Hydroxyl amine nitrate (HAN) is a powerful oxidizer for hybrid rocket flight motors. Miscible with water up to 95% by mass, it also has high density and has been extensively characterized for materials compatibility, safety, transportation, storage and handling. Before any serious attempt to use the proposed oxidizer in hybrids, though, the usual performance figures must first be obtained. The simplest are time-independent, equilibrium rocket performance numbers that include chamber temperature, temperature at the nozzle throat, and key species in the exhaust. These numbers must be followed by several other important performance evaluation, including burning rates, pressure dependence, susceptibility to instabilities and temperature sensitivity.

  16. Rockets Away!

    ERIC Educational Resources Information Center

    Kaahaaina, Nancy

    1997-01-01

    Describes a project that involved a rocket-design competition where students played the roles of McDonnell Douglas employees competing for NASA contracts. Provides a real world experience involving deadlines, design and performance specifications, and budgets. (JRH)

  17. Measuring Fluctuating Pressure Levels and Vibration Response in a Jet Plume

    NASA Technical Reports Server (NTRS)

    Osterholt, Douglas J.; Knox, Douglas M.

    2011-01-01

    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.

  18. Multispectral imaging of aircraft exhaust

    NASA Astrophysics Data System (ADS)

    Berkson, Emily E.; Messinger, David W.

    2016-05-01

    Aircraft pollutants emitted during the landing-takeoff (LTO) cycle have significant effects on the local air quality surrounding airports. There are currently no inexpensive, portable, and unobtrusive sensors to quantify the amount of pollutants emitted from aircraft engines throughout the LTO cycle or to monitor the spatial-temporal extent of the exhaust plume. We seek to thoroughly characterize the unburned hydrocarbon (UHC) emissions from jet engine plumes and to design a portable imaging system to remotely quantify the emitted UHCs and temporally track the distribution of the plume. This paper shows results from the radiometric modeling of a jet engine exhaust plume and describes a prototype long-wave infrared imaging system capable of meeting the above requirements. The plume was modeled with vegetation and sky backgrounds, and filters were selected to maximize the detectivity of the plume. Initial calculations yield a look-up chart, which relates the minimum amount of emitted UHCs required to detect the presence of a plume to the noise-equivalent radiance of a system. Future work will aim to deploy the prototype imaging system at the Greater Rochester International Airport to assess the applicability of the system on a national scale. This project will help monitor the local pollution surrounding airports and allow better-informed decision-making regarding emission caps and pollution bylaws.

  19. Air-Powered Rockets.

    ERIC Educational Resources Information Center

    Rodriguez, Charley; Raynovic, Jim

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

  20. Rain scavenging of solid rocket exhaust clouds

    NASA Technical Reports Server (NTRS)

    Dingle, A. N.

    1978-01-01

    An explicit model for cloud microphysics was developed for application to the problem of co-condensation/vaporization of HCl and H2O in the presence of Al2O3 particulate nuclei. Validity of the explicit model relative to the implicit model, which has been customarily applied to atmospheric cloud studies, was demonstrated by parallel computations of H2O condensation upon (NH4)2 SO4 nuclei. A mesoscale predictive model designed to account for the impact of wet processes on atmospheric dynamics is also under development. Input data specifying the equilibrium state of HC1 and H2O vapors in contact with aqueous HC1 solutions were found to be limited, particularly in respect to temperature range.

  1. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2005-01-01

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

  2. A smoke producing rocket motor for atmospheric wind profiling

    SciTech Connect

    Grubelich, M.C. ); Rowland, J. . Applied Physics Lab.)

    1991-01-01

    A composite propellant was developed to produce a dense plume from a rocket motor. The development of this propellant combined the smoke producing capabilities of a smoke generator with a rocket motor, thereby integrating the separate systems into one unit. A rocket motor was designed for use with this propellant to produce a high density particulate plume. This plume could then be used to determine the wind profile in the atmosphere by using a light detection and ranging system. Additionally, this smoke producing propellant could be used for rapid screening or identification. The burn rate characteristics of the propellant were measured and static firings of rocket motors were conducted to determine the performance of the propellant. The results of these tests will be presented as well as theoretical performance predictions of a flight vehicle.

  3. A smoke producing rocket motor for atmospheric wind profiling

    SciTech Connect

    Grubelich, M.C.; Rowland, J.

    1991-12-31

    A composite propellant was developed to produce a dense plume from a rocket motor. The development of this propellant combined the smoke producing capabilities of a smoke generator with a rocket motor, thereby integrating the separate systems into one unit. A rocket motor was designed for use with this propellant to produce a high density particulate plume. This plume could then be used to determine the wind profile in the atmosphere by using a light detection and ranging system. Additionally, this smoke producing propellant could be used for rapid screening or identification. The burn rate characteristics of the propellant were measured and static firings of rocket motors were conducted to determine the performance of the propellant. The results of these tests will be presented as well as theoretical performance predictions of a flight vehicle.

  4. 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)

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

    1993-01-01

    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.

  5. An in-situ measurement of particulates from solid rocket motors fired in space

    NASA Technical Reports Server (NTRS)

    Alred, J. W.

    1986-01-01

    Current models exist that predict the damage caused by the impact of aluminum oxide exhaust particles as well as their lifetime in useable space. In these models, two necessary inputs are the size and flux of the particles. An experiment, referred to as the Plume Witness Plate, was designed for the Remote Manipulator System of the space shuttle orbiter to measure in-situ the flux and material effects of a solid rocket motor (SRM) firing in space. Five different types of samples were used to provide a broad range of substances: (1) fused quartz glass (representative of orbiter windows); (2) germanium micrometeroid capture cells; (3) orbiter HRTS tiles from the thermal protection system; (4) Kapton foil; and (5) metallic disks of aluminum, copper, titanium, graphite epoxy, and gold. The analyses of the data show excellent agreement with ground-based SRM firings in terms of particle size distribution and mass distribution. The Particle Impact Damage Integrator computer model used to calculate potential damage of orbiter surfaces by SRM exhaust plumes agrees favorable with the results in terms of particle size and velocity distributions though it may be conservative by as much as 20%.

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

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

    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.

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

    NASA Technical Reports Server (NTRS)

    1990-01-01

    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.

  8. Cratering Soil by Impinging Jets of Gas, with Application to Landing Rockets on Planetary Surfaces

    NASA Technical Reports Server (NTRS)

    Metzger, Philip T.; Vu, B. T.; Taylor, D. E.; Kromann, M. J.; Fuchs, M.; Yutko, B.; Dokos, A.; Immer, Christopher D.; Lane, J. E.; Dunkel, Michael B.; Donahue, Carly M.; Latta, R. C., III

    2007-01-01

    Several physical mechanisms are involved in excavating granular materials beneath a vertical jet of gas. These occur, for example, beneath the exhaust plume of a rocket landing on the soil of the Moon or Mars. A series of experiments and simulations have been performed to provide a detailed view of the complex gas/soil interactions. Measurements have also been taken from the Apollo lunar landing videos and from photographs of the resulting terrain, and these help to demonstrate how the interactions extrapolate into the lunar environment. It is important to understand these processes at a fundamental level to support the ongoing design of higher-fidelity numerical simulations and larger-scale experiments. These are needed to enable future lunar exploration wherein multiple hardware assets will be placed on the Moon within short distances of one another. The high-velocity spray of soil from landing spacecraft must be accurately predicted and controlled lest it erosively damage the surrounding hardware.

  9. Tvashtar's Plume

    NASA Technical Reports Server (NTRS)

    2007-01-01

    This dramatic image of Io was taken by the Long Range Reconnaissance Imager (LORRI) on New Horizons at 11:04 Universal Time on February 28, 2007, just about 5 hours after the spacecraft's closest approach to Jupiter. The distance to Io was 2.5 million kilometers (1.5 million miles) and the image is centered at 85 degrees west longitude. At this distance, one LORRI pixel subtends 12 kilometers (7.4 miles) on Io.

    This processed image provides the best view yet of the enormous 290-kilometer (180-mile) high plume from the volcano Tvashtar, in the 11 o'clock direction near Io's north pole. The plume was first seen by the Hubble Space Telescope two weeks ago and then by New Horizons on February 26; this image is clearer than the February 26 image because Io was closer to the spacecraft, the plume was more backlit by the Sun, and a longer exposure time (75 milliseconds versus 20 milliseconds) was used. Io's dayside was deliberately overexposed in this picture to image the faint plumes, and the long exposure also provided an excellent view of Io's night side, illuminated by Jupiter. The remarkable filamentary structure in the Tvashtar plume is similar to details glimpsed faintly in 1979 Voyager images of a similar plume produced by Io's volcano Pele. However, no previous image by any spacecraft has shown these mysterious structures so clearly.

    The image also shows the much smaller symmetrical fountain of the plume, about 60 kilometers (or 40 miles) high, from the Prometheus volcano in the 9 o'clock direction. The top of a third volcanic plume, from the volcano Masubi, erupts high enough to catch the setting Sun on the night side near the bottom of the image, appearing as an irregular bright patch against Io's Jupiter-lit surface. Several Everest-sized mountains are highlighted by the setting Sun along the terminator, the line between day and night.

    This is the last of a handful of LORRI images that New Horizons is sending 'home' during its busy close

  10. Measurements of Aged Aircraft Exhaust in the ACCENT Mission

    NASA Technical Reports Server (NTRS)

    Friedl, R.; Ross, A.

    2000-01-01

    The Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) mission is a multi-agency sponsored effort to evaluate the roles of aircraft and rocket exhaust in perturbing ozone chemistry and modifying aerosols and clouds.

  11. 60. Historic plan of Building 202 exhaust scrubber, June 18, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    60. Historic plan of Building 202 exhaust scrubber, June 18, 1955. NASA GRC drawing no. CD-101261. (On file at NASA Glenn Research Center). - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  12. 13. Building 202 exhaust scrubber water detention tank, looking southeast ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. Building 202 exhaust scrubber water detention tank, looking southeast from bed of Abram Creek. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  13. Atomic hydrogen rocket engine

    NASA Technical Reports Server (NTRS)

    Etters, R. D.; Flurchick, K.

    1981-01-01

    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.

  14. Ice nuclei measurements from solid rocket motor effluents

    NASA Technical Reports Server (NTRS)

    Hindman, E. E., II

    1980-01-01

    The ice crystal forming nuclei (IN) measured in solid rocket motor (SRM) exhaust products is discussed in relation to space shuttle exhaust. Preliminary results from laboratory investigations and flight preparations for March 1978 Titan launch are discussed. The work necessary to provide adequate measurements of IN and cloud condensation nuclei (CCN) in the stabilized ground clouds from SRM's is studied.

  15. Space shuttle exhaust cloud properties

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    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.

  16. Preliminary study of a hydrogen peroxide rocket for use in moving source jet noise tests

    NASA Technical Reports Server (NTRS)

    Plencner, R. M.

    1977-01-01

    A preliminary investigation was made of using a hydrogen peroxide rocket to obtain pure moving source jet noise data. The thermodynamic cycle of the rocket was analyzed. It was found that the thermodynamic exhaust properties of the rocket could be made to match those of typical advanced commercial supersonic transport engines. The rocket thruster was then considered in combination with a streamlined ground car for moving source jet noise experiments. When a nonthrottlable hydrogen peroxide rocket was used to accelerate the vehicle, propellant masses and/or acceleration distances became too large. However, when a throttlable rocket or an auxiliary system was used to accelerate the vehicle, reasonable propellant masses could be obtained.

  17. Isotopic mapping of groundwater perchlorate plumes.

    PubMed

    Sturchio, Neil C; Hoaglund, John R; Marroquin, Roy J; Beloso, Abelardo D; Heraty, Linnea J; Bortz, Sarah E; Patterson, Thomas L

    2012-01-01

    Analyses of stable isotope ratios of chlorine and oxygen in perchlorate can, in some cases, be used for mapping and source identification of groundwater perchlorate plumes. This is demonstrated here for large, intersecting perchlorate plumes in groundwater from a region having extensive groundwater perchlorate contamination and a large population dependent on groundwater resources. The region contains both synthetic perchlorate derived from rocket fuel manufacturing and testing activities and agricultural perchlorate derived predominantly from imported Chilean (Atacama) nitrate fertilizer, along with a likely component of indigenous natural background perchlorate from local wet and dry atmospheric deposition. Most samples within each plume reflect either a predominantly synthetic or a predominantly agricultural perchlorate source and there is apparently a minor contribution from the indigenous natural background perchlorate. The existence of isotopically distinct perchlorate plumes in this area is consistent with other lines of evidence, including groundwater levels and flow paths as well as the historical land use and areal distribution of potential perchlorate sources.

  18. Computer model predictions of the local effects of large, solid-fuel rocket motors on stratospheric ozone. Technical report

    SciTech Connect

    Zittel, P.F.

    1994-09-10

    The solid-fuel rocket motors of large space launch vehicles release gases and particles that may significantly affect stratospheric ozone densities along the vehicle's path. In this study, standard rocket nozzle and flowfield computer codes have been used to characterize the exhaust gases and particles through the afterburning region of the solid-fuel motors of the Titan IV launch vehicle. The models predict that a large fraction of the HCl gas exhausted by the motors is converted to Cl and Cl2 in the plume afterburning region. Estimates of the subsequent chemistry suggest that on expansion into the ambient daytime stratosphere, the highly reactive chlorine may significantly deplete ozone in a cylinder around the vehicle track that ranges from 1 to 5 km in diameter over the altitude range of 15 to 40 km. The initial ozone depletion is estimated to occur on a time scale of less than 1 hour. After the initial effects, the dominant chemistry of the problem changes, and new models are needed to follow the further expansion, or closure, of the ozone hole on a longer time scale.

  19. Hybrid Rocket Propulsion for Sounding Rocket Applications

    NASA Technical Reports Server (NTRS)

    1991-01-01

    A discussion of the H-225K hybrid rocket motor, produced by the American Rocket Company, is given. The H-225K motor is presented in terms of the following topics: (1) hybrid rocket fundamentals; (2) hybrid characteristics; and (3) hybrid advantages.

  20. 40 CFR Appendix Vi to Part 266 - Stack Plume Rise

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 40 Protection of Environment 28 2012-07-01 2012-07-01 false Stack Plume Rise VI Appendix VI to Part 266 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED... FACILITIES Pt. 266, App. VI Appendix VI to Part 266—Stack Plume Rise Flow rate (m3/s) Exhaust Temperature...

  1. 40 CFR Appendix Vi to Part 266 - Stack Plume Rise

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 28 2013-07-01 2013-07-01 false Stack Plume Rise VI Appendix VI to Part 266 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED... FACILITIES Pt. 266, App. VI Appendix VI to Part 266—Stack Plume Rise Flow rate (m3/s) Exhaust Temperature...

  2. 40 CFR Appendix Vi to Part 266 - Stack Plume Rise

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 40 Protection of Environment 27 2011-07-01 2011-07-01 false Stack Plume Rise VI Appendix VI to Part 266 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED... FACILITIES Pt. 266, App. VI Appendix VI to Part 266—Stack Plume Rise Flow rate (m3/s) Exhaust Temperature...

  3. 40 CFR Appendix Vi to Part 266 - Stack Plume Rise

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 40 Protection of Environment 27 2014-07-01 2014-07-01 false Stack Plume Rise VI Appendix VI to Part 266 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED... FACILITIES Pt. 266, App. VI Appendix VI to Part 266—Stack Plume Rise Flow rate (m3/s) Exhaust Temperature...

  4. 40 CFR Appendix Vi to Part 266 - Stack Plume Rise

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 40 Protection of Environment 26 2010-07-01 2010-07-01 false Stack Plume Rise VI Appendix VI to Part 266 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED... FACILITIES Pt. 266, App. VI Appendix VI to Part 266—Stack Plume Rise Flow rate (m3/s) Exhaust Temperature...

  5. Crater Formation Due to Lunar Plume Impingement

    NASA Technical Reports Server (NTRS)

    Marsell, Brandon

    2011-01-01

    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.

  6. Skylon Aerodynamics and SABRE Plumes

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel; Afosmis, Michael; Bowles, Jeffrey; Pandya, Shishir

    2015-01-01

    An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown-a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

  7. Gas Emission Measurements from the RD 180 Rocket Engine

    NASA Technical Reports Server (NTRS)

    Ross, H. R.

    2001-01-01

    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.

  8. Challenger Rocket Booster

    NASA Technical Reports Server (NTRS)

    1986-01-01

    At about 76 seconds, fragments of the Orbiter can be seen tumbling against a background of fire, smoke and vaporized propellants from the External Tank. The left Solid Rocket Booster (SRB) flys rampant, still thrusting. The reddish-brown cloud envelops the disintergrating Orbiter. The color is indicative of the nitrogen tetroxide oxidizer propellant in the Orbiter Reaction Control System. On January 28, 1986 frigid overnight temperatures caused normally pliable rubber O-ring seals and putty that are designed to seal and establish joint integrity between the Solid Rocket Booster (SRB) joint segments, to become hard and non- flexible. At the instant of SRB ignition, tremendous stresses and pressures occur within the SRB casing and especially at the joint attachment points. The failure of the O-rings and putty to 'seat' properly at motor ignition, caused hot exhaust gases to blow by the seals and putty. During Challenger's ascent, this hot gas 'blow by' ultimately cut a swath completely through the steel booster casing; and like a welder's torch, began cutting into the External Tank (ET). It is believed that the ET was compromised in several locations starting in the aft at the initial point where SRB joint failure occured. The ET hydrogen tank is believed to have been breached first, with continuous rapid incremental failure of both the ET and SRB. The chain reaction of events occurring in milliseconds culminated in a massive explosion. The orbiter Challenger was instantly ejected by the blast and went askew into the supersonic air flow. These aerodynamic forces caused structural shattering and complete destruction of the orbiter. Though it was concluded that the G-forces experienced during orbiter ejection and break-up were survivable, impact with the ocean surface was not. Tragically, all seven crewmembers perished.

  9. Magnetic Detachment and Plume Control in Escaping Magnetized Plasma

    SciTech Connect

    P. F. Schmit and N. J. Fisch

    2008-11-05

    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.

  10. The 1991 version of the plume impingement computer program. Volume 2: User's input guide

    NASA Technical Reports Server (NTRS)

    Bender, Robert L.; Somers, Richard E.; Prendergast, Maurice J.; Clayton, Joseph P.; Smith, Sheldon D.

    1991-01-01

    The Plume Impingement Program (PLIMP) is a computer code used to predict impact pressures, forces, moments, heating rates, and contamination on surfaces due to direct impingement flowfields. Typically, it has been used to analyze the effects of rocket exhaust plumes on nearby structures from ground level to the vacuum of space. The program normally uses flowfields generated by the MOC, RAMP2, SPF/2, or SFPGEN computer programs. It is capable of analyzing gaseous and gas/particle flows. A number of simple subshapes are available to model the surfaces of any structure. The original PLIMP program has been modified many times of the last 20 years. The theoretical bases for the referenced major changes, and additional undocumented changes and enhancements since 1988 are summarized in volume 1 of this report. This volume is the User's Input Guide and should be substituted for all previous guides when running the latest version of the program. This version can operate on VAX and UNIX machines with NCAR graphics ability.

  11. Contact diagnostics of combustion products of rocket engines, their units, and systems

    NASA Astrophysics Data System (ADS)

    Ivanov, N. N.; Ivanov, A. N.

    2013-12-01

    This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors.

  12. STS-98 Emits Plume of Smoke

    NASA Technical Reports Server (NTRS)

    2001-01-01

    This awesome image depicts the full moon, sunset launch of the Space Shuttle Orbiter Atlantis STS-98 mission on February 7, 2001 at 6:13 p.m. eastern time. The large white plume is the pillar of smoke and stream left behind by the solid rocket boosters. The very bright dot that exists above the plume is the flame still visible at the base of the rocket boosters. The top of the plume is being directly illuminated by sunlight whereas the bottom portion lies within the Earth's shadow. The bright orb in the lower right-hand corner of the image is the full sunlit face of the moon which has already risen above the eastern horizon. The dark cone-shaped feature extending downward towards the moon is the smoke plume shadow, known as the Bugeron Effect (common during sunrise and sunset launches). The Earth, Moon, and Sun were naturally in alignment causing the shadow to appear to end at the moon. (Photo courtesy Patrick McCracken, NASA Headquarters)

  13. Langmuir probe surveys of an arcjet exhaust

    NASA Technical Reports Server (NTRS)

    Zana, Lynnette M.

    1987-01-01

    Electrostatic (Langmuir) probes of both spherical and cylindrical geometry have been used to obtain electron number density and temperature in the exhaust of a laboratory arcjet. The arcjet thruster operated on nitrogen and hydrogen mixtures to simulate fully decomposed hydrazine in a vacuum environment with background pressures less than 0.05 Pa. The exhaust appears to be only slightly ionized (less than 1 percent) with local plasma potentials near facility ground. The current-voltage characteristics of the probes indicate a Maxwellian temperature distribution. Plume data are presented as a function of arcjet operating conditions and also position in the exhaust.

  14. Effects of rocket engines on laser during lunar landing

    NASA Astrophysics Data System (ADS)

    Wan, Xiong; Shu, Rong; Huang, Genghua

    2013-11-01

    In the Chinese moon exploration project “ChangE-3”, the laser telemeter and lidar are important equipments on the lunar landing vehicle. A low-thrust vernier rocket engine works during the soft landing, whose plume may influence on the laser equipments. An experiment has first been accomplished to evaluate the influence of the plume on the propagation characteristics of infrared laser under the vacuum condition. Combination with our theoretical analysis has given an appropriate assessment of the plume's effects on the infrared laser hence providing a valuable basis for the design of lunar landing systems.

  15. PHYSICAL AND NUMERICAL MODELING OF ASD EXHAUST DISPERSION AROUND HOUSES

    EPA Science Inventory

    The report discusses the use of a wind tunnel to physically model the dispersion of exhaust plumes from active soil depressurization (ASD) radon mitigation systems in houses. he testing studied the effects of exhaust location (grade level vs. above the eave), as house height, roo...

  16. Program Computes Sound Pressures at Rocket Launches

    NASA Technical Reports Server (NTRS)

    Ogg, Gary; Heyman, Roy; White, Michael; Edquist, Karl

    2005-01-01

    Launch Vehicle External Sound Pressure is a computer program that predicts the ignition overpressure and the acoustic pressure on the surfaces and in the vicinity of a rocket and launch pad during launch. The program generates a graphical user interface (GUI) that gathers input data from the user. These data include the critical dimensions of the rocket and of any launch-pad structures that may act as acoustic reflectors, the size and shape of the exhaust duct or flame deflector, and geometrical and operational parameters of the rocket engine. For the ignition-overpressure calculations, histories of the chamber pressure and mass flow rate also are required. Once the GUI has gathered the input data, it feeds them to ignition-overpressure and launch-acoustics routines, which are based on several approximate mathematical models of distributed sources, transmission, and reflection of acoustic waves. The output of the program includes ignition overpressures and acoustic pressures at specified locations.

  17. Liquid rocket engine nozzles

    NASA Technical Reports Server (NTRS)

    1976-01-01

    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.

  18. 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...

  19. Validation of scramjet exhaust simulation technique

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    Scramjet/airframe integration design philosophy for hypersonic aircraft results in configurations having lower aft surfaces that serve as exhaust nozzles. There is a strong coupling between the exhaust plume and the aerodynamics of the vehicle, making accurate simulation of the engine exhaust mandatory. The experimental verification of the simulation procedure is described. The detonation tube simulator was used to produce an exact simulation of the scramjet exhaust for a Mach 8 flight condition. The pressure distributions produced by the exact exhaust flow were then duplicated by a cool mixture Argon and Freon 13B1. Such a substitute gas mixture validated by the detonation tube technique could be used in conventional wind tunnel tests. The results presented show the substitute gas simulation technique to be valid for shockless expansions.

  20. Pulse Detonation Rocket MHD Power Experiment

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

    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

  1. Sounding rockets in Antarctica

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

    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.

  2. Rockets for spin recovery

    NASA Technical Reports Server (NTRS)

    Whipple, R. D.

    1980-01-01

    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.

  3. Environmentally compatible solid rocket propellants

    NASA Technical Reports Server (NTRS)

    Jacox, James L.; Bradford, Daniel J.

    1995-01-01

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

  4. Solid propellant exhausted aluminum oxide and hydrogen chloride - Environmental considerations

    NASA Technical Reports Server (NTRS)

    Cofer, W. R., III; Winstead, E. L.; Purgold, G. C.; Edahl, R. A.

    1993-01-01

    Measurements of gaseous hydrogen chloride (HCl) and particulate aluminum oxide (Al2O3) were made during penetrations of five Space Shuttle exhaust clouds and one static ground test firing of a shuttle booster. Instrumented aircraft were used to penetrate exhaust clouds and to measure and/or collect samples of exhaust for subsequent analyses. The focus was on the primary solid rocket motor exhaust products, HCl and Al2O3, from the Space Shuttle's solid boosters. Time-dependent behavior of HCl was determined for the exhaust clouds. Composition, morphology, surface chemistry, and particle size distributions were determined for the exhausted Al2O3. Results determined for the exhaust cloud from the static test firing were complicated by having large amounts of entrained alkaline ground debris (soil) in the lofted cloud. The entrained debris may have contributed to neutralization of in-cloud HCl.

  5. An analytical approach for the prediction of gamma-to-alpha phase transformation of aluminum oxide (Al2O3) particles in the Space Shuttle ASRM and RSRM exhausts

    NASA Technical Reports Server (NTRS)

    Oliver, S. M.; Moylan, B. E.

    1992-01-01

    The analytical approach developed here utilizes the flow-field output from industry standard nozzle and plume codes as input into a particle phase conversion code which predicts the amount of gamma-to-alpha conversion in SRM exhausts. Sixty different cases were considered which varied such parameters as particle size, degree of undercooling, motor type, and altitude. On-centerline calculations were made for both the ASRM and RSRM at an altitude of 100,000 feet with particle sizes varying from 3.5 to 9.1 micron radius and undercooling varying from 0 to 20 percent. Additional calculations were made for the ASRM at 100,000 feet off centerline and at an altitude of 60,000 feet on centerline. The results indicate that significant amounts of metastable alumina will be present in ASRM and RSRM exhausts. Though not significant to motor performance, this may be important in such issues as environmental effects of rocket exhausts, plume radiative heating predictions, and particle size determination by laser scattering.

  6. Response of selected plant and insect species to simulated SRM exhaust mixtures and to exhaust components from SRM fuels

    NASA Technical Reports Server (NTRS)

    Heck, W. W.

    1980-01-01

    The possible biologic effects of exhaust products from solid rocket motor (SRM) burns associated with the space shuttle are examined. The major components of the exhaust that might have an adverse effect on vegetation, HCl and Al2O3 are studied. Dose response curves for native and cultivated plants and selected insects exposed to simulated exhaust and component chemicals from SRM exhaust are presented. A system for dispensing and monitoring component chemicals of SRM exhaust (HCl and Al2O3) and a system for exposing test plants to simulated SRM exhaust (controlled fuel burns) are described. The effects of HCl, Al2O3, and mixtures of the two on the honeybee, the corn earworm, and the common lacewing and the effects of simulated exhaust on the honeybee are discussed.

  7. 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)

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

    1976-01-01

    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.

  8. Rocket effluent - Its ice nucleation activity and related properties

    NASA Technical Reports Server (NTRS)

    Parungo, F. P.; Allee, P. A.

    1978-01-01

    To investigate the possibility of inadvertent weather modification from rocket effluent, aerosol samples were collected from an instrumented aircraft subsequent to the Voyager I and II launches. The aerosol's morphology, concentration and size distribution were examined with an electron microscope. The elemental compositions of individual particles were analyzed with an X-ray energy spectrometer. Ice nucleus concentration was measured with a subfreezing thermal diffusion chamber. The particles' physical and chemical properties were related to their ice nucleation activity. A laboratory experiment on rocket propellant exhaust was conducted under controlled conditions. Both laboratory and field experimental results indicated that rocket propellant exhaust can produce active ice nuclei. Their consequences for potential inadvertant weather modification demand additional study.

  9. Rocket effluent: Its ice nucleation activity and related properties

    NASA Technical Reports Server (NTRS)

    Parungo, F. P.; Allee, P. A.

    1978-01-01

    To investigate the possibility of inadvertent weather modification from rocket effluent, aerosol samples were collected from an instrumented aircraft subsequent to the Voyager 1 and 2 launches. The aerosol's morphology, concentration, and size distribution were examined with an electron microscope. The elemental compositions of individual particles were analyzed with an X-ray energy spectrometer. Ice nucleus concentration was measured with a thermal diffusion chamber. The particles' physical and chemical properties were related to their ice nucleation activity. A laboratory experiment on rocket propellant exhaust was conducted under controlled conditions. Both laboratory and field experimental results indicated that rocket propellant exhaust can produce active ice nuclei and modify local weather in suitable meteorological conditions.

  10. Imaging Fourier transform spectrometry of chemical plumes

    NASA Astrophysics Data System (ADS)

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

    2009-05-01

    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.

  11. Modification of the Simons model for calculation of nonradial expansion plumes

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    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.

  12. Turbulent Plumes in Nature

    NASA Astrophysics Data System (ADS)

    Woods, Andrew W.

    2010-01-01

    This review describes a range of natural processes leading to the formation of turbulent buoyant plumes, largely relating to volcanic processes, in which there are localized, intense releases of energy. Phenomena include volcanic eruption columns, bubble plumes in lakes, hydrothermal plumes, and plumes beneath the ice in polar oceans. We assess how the dynamics is affected by heat transfer, particle fallout and recycling, and Earth's rotation, as well as explore some of the mixing of the ambient fluid produced by plumes in a confined geometry.

  13. Modeling Europa's dust plumes

    NASA Astrophysics Data System (ADS)

    Southworth, B. S.; Kempf, S.; Schmidt, J.

    2015-12-01

    The discovery of Jupiter's moon Europa maintaining a probably sporadic water vapor plume constitutes a huge scientific opportunity for NASA's upcoming mission to this Galilean moon. Measuring properties of material emerging from interior sources offers a unique chance to understand conditions at Europa's subsurface ocean. Exploiting results obtained for the Enceladus plume, we simulate possible Europa plume configurations, analyze particle number density and surface deposition results, and estimate the expected flux of ice grains on a spacecraft. Due to Europa's high escape speed, observing an active plume will require low-altitude flybys, preferably at altitudes of 5-100 km. At higher altitudes a plume may escape detection. Our simulations provide an extensive library documenting the possible structure of Europa dust plumes, which can be quickly refined as more data on Europa dust plumes are collected.

  14. Analysis on Impulse Characteristics of PDRE with Exhaust Measurements

    NASA Astrophysics Data System (ADS)

    Hu, Hong-bo; Weng, Chun-sheng; Lv, Xiao-jing; Li, Ning

    2014-06-01

    The exhaust characteristics related to impulse was investigated in a pulse detonation rocket engine (PDRE) by tunable diode laser absorption sensing system. The instantaneous parameters of temperature, velocity and pressure were obtained for exhaust at engine exit. Analysis on impulse characteristics based on control volume of the PDRE was conducted for a single operation circle with experimental results. It was concluded that the impulse (3.26 N·s) achieved by exhaust measurements was in agreement with that (3.09 N·s) by a load cell. The impulse caused by exhaust momentum experienced an extremely sharp ascending, a steep rising and a slow increment in sequence. The exhausts during the sharp ascending and steep rising were under expansion with high mass weighted average temperature (>1266 K), so there was a possible promotion for exhausts utilizing.

  15. Sounding rocket lessons learned

    NASA Technical Reports Server (NTRS)

    Wessling, Francis C.; Maybee, George W.

    1991-01-01

    Programmatic, applicatory, developmental, and operational aspects of sounding rocket utilization for materials processing studies are discussed. Lessons learned through the experience of 10 sounding rocket missions are described. Particular attention is given to missions from the SPAR, Consort, and Joust programs. Successful experiments on Consort include the study of polymer membranes and resins, biological processes, demixing of immiscible liquids, and electrodeposition.

  16. The Rocket Project.

    ERIC Educational Resources Information Center

    Winemiller, Jake; And Others

    1991-01-01

    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)

  17. Life Saving Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    By 1870, American and British inventors had found other ways to use rockets. For example, the Congreve rocket was capable of carrying a line over 1,000 feet to a stranded ship. In 1914, an estimated 1,000 lives were saved by this technique.

  18. Model Rockets and Microchips.

    ERIC Educational Resources Information Center

    Fitzsimmons, Charles P.

    1986-01-01

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

  19. Rockets -- Part II.

    ERIC Educational Resources Information Center

    Leitner, Alfred

    1982-01-01

    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)

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

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

    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.

  1. Development of a 12-Thrust Chamber Kerosene /Oxygen Primary Rocket Sub-System for an Early (1964) Air-Augmented Rocket Ground-Test System

    NASA Technical Reports Server (NTRS)

    Pryor, D.; Hyde, E. H.; Escher, W. J. D.

    1999-01-01

    Airbreathing/Rocket combined-cycle, and specifically rocket-based combined- cycle (RBCC), propulsion systems, typically employ an internal engine flow-path installed primary rocket subsystem. To achieve acceptably short mixing lengths in effecting the "air augmentation" process, a large rocket-exhaust/air interfacial mixing surface is needed. This leads, in some engine design concepts, to a "cluster" of small rocket units, suitably arrayed in the flowpath. To support an early (1964) subscale ground-test of a specific RBCC concept, such a 12-rocket cluster was developed by NASA's Marshall Space Flight Center (MSFC). The small primary rockets used in the cluster assembly were modified versions of an existing small kerosene/oxygen water-cooled rocket engine unit routinely tested at MSFC. Following individual thrust-chamber tests and overall subsystem qualification testing, the cluster assembly was installed at the U. S. Air Force's Arnold Engineering Development Center (AEDC) for RBCC systems testing. (The results of the special air-augmented rocket testing are not covered here.) While this project was eventually successfully completed, a number of hardware integration problems were met, leading to catastrophic thrust chamber failures. The principal "lessons learned" in conducting this early primary rocket subsystem experimental effort are documented here as a basic knowledge-base contribution for the benefit of today's RBCC research and development community.

  2. Automated and Manual Rocket Crater Measurement Software

    NASA Technical Reports Server (NTRS)

    Metzger, Philip; Immer, Christopher

    2012-01-01

    An update has been performed to software designed to do very rapid automated measurements of craters created in sandy substrates by rocket exhaust on liftoff. The previous software was optimized for pristine lab geometry and lighting conditions. This software has been enhanced to include a section for manual measurements of crater parameters; namely, crater depth, crater full width at half max, and estimated crater volume. The tools provide a very rapid method to measure these manual parameters to ease the burden of analyzing large data sets. This software allows for rapid quantization of the rocket crater parameters where automated methods may not work. The progress of spreadsheet data is continuously saved so that data is never lost, and data can be copied to clipboards and pasted to other software for analysis. The volume estimation of a crater is based on the central max depth axis line, and the polygonal shape of the crater is integrated around that axis.

  3. Acid droplet generation in SRM exhaust clouds

    NASA Technical Reports Server (NTRS)

    Dingle, A. N.

    1983-01-01

    A free energy analysis is applied to the co-condensation/evaporation of H2O and HCl vapors on wettable particles in open air in order to model droplet nucleation in solid rocket motor (SRM) exhaust clouds. Formulations are defined for the free energy change, the drop radius, the saturation ratio, the total number of molecules, and the mean molecular radius in solution, as well as the molecular volume and the concentration range. The free energy release in the phase transition for the AL2O3 nuclei in the SRM exhaust is examined as a function of the HCl molefraction and nucleating particle radius, based on Titan III launch exhaust cloud conditions 90 sec after ignition. The most efficient droplet growth is determined to occur at an HCl molefraction of 0.082 and a particle radius of 0.0000013 cm, i.e. a molality of 5.355.

  4. 61. Historic elevation and section drawing of Building 202 exhaust ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    61. Historic elevation and section drawing of Building 202 exhaust scrubber, July 18, 1955. NASA GRC drawing no. CD-101263. (On file at NASA Glenn Research Center). - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  5. 63. Historic detail drawing of inlet duct cone on exhaust ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    63. Historic detail drawing of inlet duct cone on exhaust scrubber at building 202, June 18, 1955. NASA GRC drawing no. CD-101266. (On file at NASA Glenn Research Center). - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  6. 28. Historic view of Building 202 exhaust scrubber stack, detail, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    28. Historic view of Building 202 exhaust scrubber stack, detail, July 31, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45648. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  7. 27. Historic view of Building 202 exhaust scrubber stack, July ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    27. Historic view of Building 202 exhaust scrubber stack, July 31, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45650. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  8. Simulation of wake vortex radiometric detection via jet exhaust proxy

    NASA Astrophysics Data System (ADS)

    Daniels, Taumi S.

    2015-06-01

    This paper describes an analysis of the potential of an airborne hyperspectral imaging IR instrument to infer wake vortices via turbine jet exhaust as a proxy. The goal was to determine the requirements for an imaging spectrometer or radiometer to effectively detect the exhaust plume, and by inference, the location of the wake vortices. The effort examines the gas spectroscopy of the various major constituents of turbine jet exhaust and their contributions to the modeled detectable radiance. Initially, a theoretical analysis of wake vortex proxy detection by thermal radiation was realized in a series of simulations. The first stage used the SLAB plume model to simulate turbine jet exhaust plume characteristics, including exhaust gas transport dynamics and concentrations. The second stage used these plume characteristics as input to the Line By Line Radiative Transfer Model (LBLRTM) to simulate responses from both an imaging IR hyperspectral spectrometer or radiometer. These numerical simulations generated thermal imagery that was compared with previously reported wake vortex temperature data. This research is a continuation of an effort to specify the requirements for an imaging IR spectrometer or radiometer to make wake vortex measurements. Results of the two-stage simulation will be reported, including instrument specifications for wake vortex thermal detection. These results will be compared with previously reported results for IR imaging spectrometer performance.

  9. Simulation of Wake Vortex Radiometric Detection via Jet Exhaust Proxy

    NASA Technical Reports Server (NTRS)

    Daniels, Taumi S.

    2015-01-01

    This paper describes an analysis of the potential of an airborne hyperspectral imaging IR instrument to infer wake vortices via turbine jet exhaust as a proxy. The goal was to determine the requirements for an imaging spectrometer or radiometer to effectively detect the exhaust plume, and by inference, the location of the wake vortices. The effort examines the gas spectroscopy of the various major constituents of turbine jet exhaust and their contributions to the modeled detectable radiance. Initially, a theoretical analysis of wake vortex proxy detection by thermal radiation was realized in a series of simulations. The first stage used the SLAB plume model to simulate turbine jet exhaust plume characteristics, including exhaust gas transport dynamics and concentrations. The second stage used these plume characteristics as input to the Line By Line Radiative Transfer Model (LBLRTM) to simulate responses from both an imaging IR hyperspectral spectrometer or radiometer. These numerical simulations generated thermal imagery that was compared with previously reported wake vortex temperature data. This research is a continuation of an effort to specify the requirements for an imaging IR spectrometer or radiometer to make wake vortex measurements. Results of the two-stage simulation will be reported, including instrument specifications for wake vortex thermal detection. These results will be compared with previously reported results for IR imaging spectrometer performance.

  10. Indians Repulse British With Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    During the early introduction of rockets to Europe, they were used only as weapons. Enemy troops in India repulsed the British with rockets. Later, in Britain, Sir William Congreve developed a rocket that could fire to about 9,000 feet. The British fired Congreve rockets against the United States in the War of 1812.

  11. Quantification of Plume-Soil Interaction and Excavation Due to the Sky Crane Descent Stage

    NASA Technical Reports Server (NTRS)

    Vizcaino, Jeffrey; Mehta, Manish

    2015-01-01

    The quantification of the particulate erosion that occurs as a result of a rocket exhaust plume impinging on soil during extraterrestrial landings is critical for future robotic and human lander mission design. The aerodynamic environment that results from the reflected plumes results in dust lifting, site alteration and saltation, all of which create a potentially erosive and contaminant heavy environment for the lander vehicle and any surrounding structures. The Mars Science Lab (MSL), weighing nearly one metric ton, required higher levels of thrust from its retro propulsive systems and an entirely new descent system to minimize these effects. In this work we seek to quantify plume soil interaction and its resultant soil erosion caused by the MSL's Sky Crane descent stage engines by performing three dimensional digital terrain and elevation mapping of the Curiosity rover's landing site. Analysis of plume soil interaction altitude and time was performed by detailed examination of the Mars Descent Imager (MARDI) still frames and reconstructed inertial measurement unit (IMU) sensor data. Results show initial plume soil interaction from the Sky Crane's eight engines began at ground elevations greater than 60 meters and more than 25 seconds before the rovers' touchdown event. During this time, viscous shear erosion (VSE) was dominant typically resulting in dusting of the surface with flow propagating nearly parallel to the surface. As the vehicle descended and decreased to four powered engines plume-plume and plume soil interaction increased the overall erosion rate at the surface. Visibility was greatly reduced at a height of roughly 20 meters above the surface and fell to zero ground visibility shortly after. The deployment phase of the Sky Crane descent stage hovering at nearly six meters above the surface showed the greatest amount of erosion with several large particles of soil being kicked up, recirculated, and impacting the bottom of the rover chassis. Image

  12. Oxidation Behavior of Copper Alloy Candidates for Rocket Engine Applications (Technical Poster)

    NASA Technical Reports Server (NTRS)

    Ogbuji, Linus U. J.; Humphrey, Donald H.; Barrett, Charles A.; Greenbauer-Seng, Leslie (Technical Monitor); Gray, Hugh R. (Technical Monitor)

    2002-01-01

    A rocket engine's combustion chamber is lined with material that is highly conductive to heat in order to dissipate the huge thermal load (evident in a white-hot exhaust plume). Because of its thermal conductivity copper is the best choice of liner material. However, the mechanical properties of pure copper are inadequate to withstand the high stresses, hence, copper alloys are needed in this application. But copper and its alloys are prone to oxidation and related damage, especially "blanching" (an oxidation-reduction mode of degradation). The space shuttle main engine combustion chamber is lined with a Cu-Ag-Zr alloy, "NARloy-Z", which exhibits blanching. A superior liner is being sought for the next generation of RLVs (Reusable Launch Vehicles) It should have improved mechanical properties and higher resistance to oxidation and blanching, but without substantial penalty in thermal conductivity. GRCop84, a Cu-8Cr-4Nb alloy (Cr2Nb in Cu matrix), developed by NASA Glenn Research Center (GRC) and Case Western Reserve University, is a prime contender for RLV liner material. In this study, the oxidation resistance of GRCop-84 and other related/candidate copper alloys are investigated and compared

  13. Baking Soda and Vinegar Rockets

    NASA Astrophysics Data System (ADS)

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

    2009-02-01

    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors1,2 that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the experimentally measured rocket height. Baking soda and vinegar rockets present fewer safety concerns and require a smaller launch area than rapid combustion chemical rockets. Both kits were of nearly identical design, costing ˜20. The rockets required roughly 30 minutes of assembly time consisting of mostly taping the soft plastic fuselage to the Styrofoam nose cone.

  14. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

    Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads several hundred miles in altitude. These missions return a variety of scientific data including; chemical makeup and physical processes taking place in the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices used on satellites and other spacecraft prior to their use in more expensive activities. This paper addresses the NASA Wallops Island history of GPS Sounding Rocket experience since 1994 and the development of highly accurate and useful system.

  15. Modeling Europa's Dust Plumes

    NASA Astrophysics Data System (ADS)

    Southworth, B.; Kempf, S.; Schmidt, J.

    2015-12-01

    The discovery of Europa maintaining a probably sporadic water vapor plume constitutes a huge scientific opportunity for NASA's upcoming mission to this Galilean moon. Measuring the properties of material emerging from interior sources offers a unique chance to understand conditions at Europa's subsurface ocean. Exploiting results obtained for the Enceladus plume, we adjust the ejection model by Schmidt et al. [2008] to the conditions at Europa. In this way, we estimate properties of a possible, yet unobserved dust component of the Europa plume. For a size-dependent speed distribution of emerging ice particles we use the model from Kempf et al. [2010] for grain dynamics, modified to run simulations of plumes on Europa. Specifically, we model emission from the two plume locations determined from observations by Roth et al. [2014] and also from other locations chosen at the closest approach of low-altitude flybys investigated in the Europa Clipper study. This allows us to estimate expected fluxes of ice grains on the spacecraft. We then explore the parameter space of Europa dust plumes with regard to particle speed distribution parameters, plume location, and spacecraft flyby elevation. Each parameter set results in a 3-dimensional particle density structure through which we simulate flybys, and a map of particle fallback ('snowfall') on the surface of Europa. Due to the moon's high escape speed, a Europa plume will eject few to no particles that can escape its gravity, which has several further consequences: (i) For given ejection velocity a Europa plume will have a smaller scale height, with a higher particle number densities than the plume on Enceladus, (ii) plume particles will not feed the diffuse Galilean dust ring, (iii) the snowfall pattern on the surface will be more localized about the plume location, and will not induce a global m = 2 pattern as seen on Enceladus, and (iv) safely observing an active plume will require low altitude flybys, preferably at 50

  16. Characterization of rocket propellant combustion products

    SciTech Connect

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

    1991-12-09

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    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.

  18. Highlights of Transient Plume Impingement Model Validation and Applications

    NASA Technical Reports Server (NTRS)

    Woronowicz, Michael

    2011-01-01

    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.

  19. Rocket University at KSC

    NASA Technical Reports Server (NTRS)

    Sullivan, Steven J.

    2014-01-01

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

  20. 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...

  1. Rocketing into Adaptive Inquiry.

    ERIC Educational Resources Information Center

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

    2002-01-01

    Defines adaptive inquiry and argues for employing this method which allows lessons to be shaped in response to student needs. Illustrates this idea by detailing an activity in which teams of students build rockets. (DDR)

  2. The History of Rockets.

    ERIC Educational Resources Information Center

    Newby, J. C.

    1988-01-01

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

  3. Rocket engine numerical simulator

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in viewgraph form and include the following: a rocket engine numerical simulator (RENS) definition; objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusion.

  4. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

  5. Monitoring Engine Vibrations And Spectrum Of Exhaust

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

    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).

  6. Rocket Motor Microphone Investigation

    NASA Technical Reports Server (NTRS)

    Pilkey, Debbie; Herrera, Eric; Gee, Kent L.; Giraud, Jerom H.; Young, Devin J.

    2010-01-01

    At ATK's facility in Utah, large full-scale solid rocket motors are tested. The largest is a five-segment version of the reusable solid rocket motor, which is for use on the Ares I launch vehicle. As a continuous improvement project, ATK and BYU investigated the use of microphones on these static tests, the vibration and temperature to which the instruments are subjected, and in particular the use of vent tubes and the effects these vents have at low frequencies.

  7. Numerical simulation of helicopter engine plume in forward flight

    NASA Technical Reports Server (NTRS)

    Dimanlig, Arsenio C. B.; Vandam, Cornelis P.; Duque, Earl P. N.

    1994-01-01

    Flowfields around helicopters contain complex flow features such as large separated flow regions, vortices, shear layers, blown and suction surfaces and an inherently unsteady flow imposed by the rotor system. Another complicated feature of helicopters is their infrared signature. Typically, the aircraft's exhaust plume interacts with the rotor downwash, the fuselage's complicated flowfield, and the fuselage itself giving each aircraft a unique IR signature at given flight conditions. The goal of this project was to compute the flow about a realistic helicopter fuselage including the interaction of the engine air intakes and exhaust plume. The computations solve the Think-Layer Navier Stokes equations using overset type grids and in particular use the OVERFLOW code by Buning of NASA Ames. During this three month effort, an existing grid system of the Comanche Helicopter was to be modified to include the engine inlet and the hot engine exhaust. The engine exhaust was to be modeled as hot air exhaust. However, considerable changes in the fuselage geometry required a complete regriding of the surface and volume grids. The engine plume computations have been delayed to future efforts. The results of the current work consists of a complete regeneration of the surface and volume grids of the most recent Comanche fuselage along with a flowfield computation.

  8. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

    Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads several hundred miles in altitude. These missions return a variety of scientific data including; chemical makeup and physical processes taking place In the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices used on satellites and other spacecraft prior to their use in more expensive activities. The NASA Sounding Rocket Program is managed by personnel from Goddard Space Flight Center Wallops Flight Facility (GSFC/WFF) in Virginia. Typically around thirty of these rockets are launched each year, either from established ranges at Wallops Island, Virginia, Poker Flat Research Range, Alaska; White Sands Missile Range, New Mexico or from Canada, Norway and Sweden. Many times launches are conducted from temporary launch ranges in remote parts of the world requi6ng considerable expense to transport and operate tracking radars. An inverse differential GPS system has been developed for Sounding Rocket. This paper addresses the NASA Wallops Island history of GPS Sounding Rocket experience since 1994 and the development of a high accurate and useful system.

  9. Infrasound Rocket Signatures

    NASA Astrophysics Data System (ADS)

    Olson, J.

    2012-09-01

    This presentation reviews the work performed by our research group at the Geophysical Institute as we have applied the tools of infrasound research to rocket studies. This report represents one aspect of the effort associated with work done for the National Consortium for MASINT Research (NCMR) program operated by the National MASINT Office (NMO) of the Defense Intelligence Agency (DIA). Infrasound, the study of acoustic signals and their propagation in a frequency band below 15 Hz, enables an investigator to collect and diagnose acoustic signals from distant sources. Absorption of acoustic energy in the atmosphere decreases as the frequency is reduced. In the infrasound band signals can propagate hundreds and thousands of kilometers with little degradation. We will present an overview of signatures from rockets ranging from small sounding rockets such as the Black Brandt and Orion series to larger rockets such as Delta 2,4 and Atlas V. Analysis of the ignition transients provides information that can uniquely identify the motor type. After the rocket ascends infrasound signals can be used to characterize the rocket and identify the various events that take place along a trajectory such as staging and maneuvering. We have also collected information on atmospheric shocks and sonic booms from the passage of supersonic vehicles such as the shuttle. This review is intended to show the richness of the unique signal set that occurs in the low-frequency infrasound band.

  10. Volcanic Plume Measurements with UAV (Invited)

    NASA Astrophysics Data System (ADS)

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

    2013-12-01

    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

  11. SRB Environment Evaluation and Analysis. Volume 3: ASRB Plume Induced Environments

    NASA Technical Reports Server (NTRS)

    Bender, R. L.; Brown, J. R.; Reardon, J. E.; Everson, J.; Coons, L. W.; Stuckey, C. I.; Fulton, M. S.

    1991-01-01

    Contract NAS8-37891 was expanded in late 1989 to initiate analysis of Shuttle plume induced environments as a result of the substitution of the Advanced Solid Rocket Booster (ASRB) for the Redesigned Solid Rocket Booster (RSRB). To support this analysis, REMTECH became involved in subscale and full-scale solid rocket motor test programs which further expanded the scope of work. Later contract modifications included additional tasks to produce initial design cycle environments and to specify development flight instrumentation. Volume 3 of the final report describes these analyses and contains a summary of reports resulting from various studies.

  12. Stealth Plumes on Io

    NASA Technical Reports Server (NTRS)

    Johnson, T. V.; Matson, Dennis L.; Blaney, Diana L.; Veeder, Glenn J.; Davies, Ashley

    1995-01-01

    We suggest that Io's eruptive activity may include a class of previously undetected SO2 geysers. The thermodynamic models for the eruptive plumes discovered by Voyager 'involve low to moderate entropy SO2 eruptions. The resulting plumes are a mixture of solid and gas which emerge from the vent and follow essentially ballistic trajectories. We show that intrusion of silicate magma into buried SO2 deposits can create the required conditions for high entropy eruptions which proceed entirely in the vapor phase. These purely gaseous plumes would have been invisible to Voyager's instruments. Hence, we call them "stealth" plumes. Such eruptions could explain the "patchy" SO2 atmosphere inferred from recent UV and micro-wave spectral observations. The magma intrusion rate required to support the required gas production for these plumes is a negligible fraction of estimated global magma intrusion rates.

  13. Rocket exhaust effects as active space plasma experiments of opportunity

    NASA Astrophysics Data System (ADS)

    Mendillo, M.

    1983-07-01

    Examples of how photometer and wide-angle airglow imaging systems can be used to study diffusive and photochemical properties of the upper atmosphere are given. Incoherent scatter measurements of a large-scale ionospheric hole are shown to yield estimates of dynamical and chemical rate constants associated with the plasma perturbations themselsves. The Spacelab-2 series of shuttle engine burn experiments are summarized.

  14. Apollo video photogrammetry estimation of plume impingement effects

    NASA Astrophysics Data System (ADS)

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

    2011-07-01

    Future missions to the Moon may require 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 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 modern 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°. The lofted particle density is estimated at 10 8-10 13 particles/m 3. 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.

  15. Identification of Noise Sources During Rocket Engine Test Firings and a Rocket Launch Using a Microphone Phased-Array

    NASA Technical Reports Server (NTRS)

    Panda, Jayanta; Mosher, Robert N.; Porter, Barry J.

    2013-01-01

    A 70 microphone, 10-foot by 10-foot, microphone phased array was built for use in the harsh environment of rocket launches. The array was setup at NASA Wallops launch pad 0A during a static test firing of Orbital Sciences' Antares engines, and again during the first launch of the Antares vehicle. It was placed 400 feet away from the pad, and was hoisted on a scissor lift 40 feet above ground. The data sets provided unprecedented insight into rocket noise sources. The duct exit was found to be the primary source during the static test firing; the large amount of water injected beneath the nozzle exit and inside the plume duct quenched all other sources. The maps of the noise sources during launch were found to be time-dependent. As the engines came to full power and became louder, the primary source switched from the duct inlet to the duct exit. Further elevation of the vehicle caused spilling of the hot plume, resulting in a distributed noise map covering most of the pad. As the entire plume emerged from the duct, and the ondeck water system came to full power, the plume itself became the loudest noise source. These maps of the noise sources provide vital insight for optimization of sound suppression systems for future Antares launches.

  16. ROCKET PORT CLOSURE

    DOEpatents

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  17. Prometheus: Io's wandering plume.

    PubMed

    Kieffer, S W; Lopes-Gautier, R; McEwen, A; Smythe, W; Keszthelyi, L; Carlson, R

    2000-05-19

    Unlike any volcanic behavior ever observed on Earth, the plume from Prometheus on Io has wandered 75 to 95 kilometers west over the last 20 years since it was first discovered by Voyager and more recently observed by Galileo. Despite the source motion, the geometric and optical properties of the plume have remained constant. We propose that this can be explained by vaporization of a sulfur dioxide and/or sulfur "snowfield" over which a lava flow is moving. Eruption of a boundary-layer slurry through a rootless conduit with sonic conditions at the intake of the melted snow can account for the constancy of plume properties. PMID:10817989

  18. General view of the Solid Rocket Booster's (SRB) Solid Rocket ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of the Solid Rocket Booster's (SRB) Solid Rocket Motor Segments in the Surge Building of the Rotation Processing and Surge Facility at Kennedy Space Center awaiting transfer to the Vehicle Assembly Building and subsequent mounting and assembly on the Mobile Launch Platform. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  19. The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics

    NASA Technical Reports Server (NTRS)

    Deans, Matthew C.; Reed, Brian D.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; Kinzbach, McKenzie I.; McLean, Christopher H.

    2014-01-01

    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 model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters are currently being tested in a small rocket, altitude facility at NASA GRC. A suite of diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, and Schlieren imaging are being used to acquire plume measurements of AF-M315E 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.

  20. Rockets in World War I

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  1. Heat Exhaustion, First Aid

    MedlinePlus

    ... rashes clinical tools newsletter | contact Share | Heat Exhaustion, First Aid A A A Heat exhaustion signs and symptoms ... specific to the other stages of heat illness. First Aid Guide Use a combination of the following measures ...

  2. Laser Rayleigh and Raman Diagnostics for Small Hydrogen/oxygen Rockets

    NASA Technical Reports Server (NTRS)

    Degroot, Wilhelmus A.; Zupanc, Frank J.

    1993-01-01

    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.

  3. 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).

  4. Mars Methane Plume Tracer

    NASA Astrophysics Data System (ADS)

    Mischna, M. A.; Banfield, D.; Sykes, I.

    2014-07-01

    Putative releases of methane from the martian surface may be challenging to detect from orbit. Successful detections depend on the character of the plume itself (duration, magnitude, expanse), but also on the observing platform.

  5. Methane Plumes on Mars

    NASA Video Gallery

    Spectrometer instruments attached to several telescopes detect plumes of methane emitted from Mars during its summer and spring seasons. High levels of methane are indicated by warmer colors. The m...

  6. Sulfur plumes off Namibia

    NASA Technical Reports Server (NTRS)

    2002-01-01

    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

  7. Baking Soda and Vinegar Rockets

    ERIC Educational Resources Information Center

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

    2009-01-01

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

  8. Overview of rocket engine control

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Musgrave, Jeffrey L.

    1991-01-01

    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.

  9. Rocket center Peenemuende - Personal memories

    NASA Technical Reports Server (NTRS)

    Dannenberg, Konrad; Stuhlinger, Ernst

    1993-01-01

    A brief history of Peenemuende, the rocket center where Von Braun and his team developed the A-4 (V-2) rocket under German Army auspices, and the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes, is presented. Emphasis is placed on the expansion of operations beginning in 1942.

  10. Controlling plume deflection by acoustic excitation - An experimental demonstration

    NASA Astrophysics Data System (ADS)

    Ahuja, K. K.

    1990-10-01

    Effect of imposing an external sound field on a Coanda jet was investigated experimentally. It was found that the exhaust angle of a Coanda plume can be varied by changing the level of excitation. Limited experiments were also performed in a wind tunnel to study the effects of flight simulation on plume deflection controllability by sound using a hollow airfoil fitted with a Coanda jet. Pressure coefficients are measured over this airfoil with and without acoustic excitation of the Coanda Jet. This exploratory study provided a number of new ideas for future work for controlling flow over curved surfaces.

  11. Water Rocket Workout.

    ERIC Educational Resources Information Center

    Esler, William K.; Sanford, Daniel

    1989-01-01

    Water rockets are used to present Newton's three laws of motion to high school physics students. Described is an outdoor activity which uses four students per group. Provides a launch data sheet to record height, angle of elevation, amount of water used, and launch number. (MVL)

  12. This "Is" Rocket Science!

    ERIC Educational Resources Information Center

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

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

  13. The Relativistic Rocket

    ERIC Educational Resources Information Center

    Antippa, Adel F.

    2009-01-01

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

  14. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

  15. Solid rocket motors

    NASA Technical Reports Server (NTRS)

    Carpenter, Ronn L.

    1993-01-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  16. Thiokol Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Graves, S. R.

    2000-01-01

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

  17. Solid rocket motors

    NASA Astrophysics Data System (ADS)

    Carpenter, Ronn L.

    1993-02-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  18. Dr. Goddard Transports Rocket

    NASA Technical Reports Server (NTRS)

    1974-01-01

    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.

  19. Liquid rocket valve assemblies

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.

  20. Liquid rocket valve components

    NASA Technical Reports Server (NTRS)

    1973-01-01

    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.

  1. This Is Rocket Science!

    NASA Astrophysics Data System (ADS)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

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

  2. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  3. Rockets using Liquid Oxygen

    NASA Technical Reports Server (NTRS)

    Busemann, Adolf

    1947-01-01

    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.

  4. Infrared Signature Modeling and Analysis of Aircraft Plume

    NASA Astrophysics Data System (ADS)

    Rao, Arvind G.

    2011-09-01

    In recent years, the survivability of an aircraft has been put to task more than ever before. One of the main reasons is the increase in the usage of Infrared (IR) guided Anti-Aircraft Missiles, especially due to the availability of Man Portable Air Defence System (MANPADS) with some terrorist groups. Thus, aircraft IR signatures are gaining more importance as compared to their radar, visual, acoustic, or any other signatures. The exhaust plume ejected from the aircraft is one of the important sources of IR signature in military aircraft that use low bypass turbofan engines for propulsion. The focus of the present work is modelling of spectral IR radiation emission from the exhaust jet of a typical military aircraft and to evaluate the aircraft susceptibility in terms of the aircraft lock-on range due to its plume emission, for a simple case against a typical Surface to Air Missile (SAM). The IR signature due to the aircraft plume is examined in a holistic manner. A comprehensive methodology of computing IR signatures and its affect on aircraft lock-on range is elaborated. Commercial CFD software has been used to predict the plume thermo-physical properties and subsequently an in-house developed code was used for evaluating the IR radiation emitted by the plume. The LOWTRAN code has been used for modeling the atmospheric IR characteristics. The results obtained from these models are in reasonable agreement with some available experimental data. The analysis carried out in this paper succinctly brings out the intricacy of the radiation emitted by various gaseous species in the plume and the role of atmospheric IR transmissivity in dictating the plume IR signature as perceived by an IR guided SAM.

  5. Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  6. Collapse in Thermal Plumes

    NASA Astrophysics Data System (ADS)

    Pears, M. I.; Lithgow-Bertelloni, C. R.; Dobson, D. P.; Davies, R.

    2013-12-01

    Collapsing thermal plumes have been investigated through experimental and numerical simulations. Collapsing plumes are an uncommon fluid dynamical phenomenon, usually seen when the buoyancy source is turned off. A series of fluid dynamical experiments were conducted on thermal plumes at a variety of temperature and viscosity contrasts, in a 26.5 cm^3 cubic tank heated by a constant temperature heater 2 cm in diameter and no-slip bottom and top surfaces. Working fluids included Lyle's Golden Syrup and ADM's Liquidose 436 syrup, which have strongly-temperature dependent viscosity and high Pr number (10^3-10^7 at experimental conditions). Visualisation included white light shadowgraphs and PIV of the central plane. Temperature contrasts ranged from 3-60°C, and two differing forms of collapse were identified. At very low temperature differences 'no rise' collapse was discovered, where the plumes stagnate in the lower third of the tank before collapsing. At temperature differences between 10-23°C normal evolution occurred until 'lens shape' collapse developed between midway and two-thirds of the distance from the base. The lens shape originated in the top of the conduit and was present throughout collapse. At temperatures above ΔT=23°C the plumes follow the expected growth and shape and flatten out at the top of the tank. Thermal collapse remains difficult to explain given experimental conditions (continuous heating). Instead it is possible that small density differences arising from crystallization at ambient temperatures changes plume buoyancy-inducing collapse. We show results on the evolution of the refractive index of the syrup through time to ascertain this possibility. Preliminary numerical results using Fluidity will be presented to explore a greater parameter range of viscosity contrasts and tank aspect ratios.

  7. Rapid Mars transits with exhaust-modulated plasma propulsion

    NASA Technical Reports Server (NTRS)

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

    1995-01-01

    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.

  8. Plumes Do Not Exist

    NASA Astrophysics Data System (ADS)

    Hamilton, W. B.; Anderson, D. L.; Foulger, G. R.; Winterer, E. L.

    Hypothetical plumes from the deep mantle are widely assumed to provide an abso- lute hotspot reference frame, inaugurate rifting, drive plates, and profoundly influence magmatic and tectonic evolution of oceans and continents. Many papers on local to global tectonics, magmatism, and geochemistry invoke plumes, and assign to the man- tle whatever properties, dynamics, and composition are needed to enable them. The fixed-plume concept arose from the Emperor-Hawaii seamount-and-island province, the 45 Ma inflection in which was assumed to record a 60-degree change in direction by the Pacific plate. Paleomagnetic latitudes and smooth Pacific spreading patterns show that such a change did not occur. Other Pacific chains once assumed to be syn- chronous with, and Euler-parallel to, Hawaii have proved to be neither. Thermal and physical properties of Hawaiian lithosphere falsify plume predictions. Rationales for fixed hotspots elsewhere also have become untenable as databases enlarged. Astheno- sphere is everywhere near solidus temperature, so buoyant melt does not require a local heat source but, rather, needs a thin roof or crack or tensional setting for egress. MORB and ocean-island basalt (OIB) broadly intergrade in composition, but MORB typically is richer in refractory elements and their radiogenic daughters, whereas OIB commonly is richer in fusible elements and their daughters. MORB and OIB contrasts are required by melt behavior and do not indicate unlike source reservoirs. MORB melts rise, with minimal reaction, through hot asthenosphere, whereas OIB melts re- act, and thereby lose substance, by crystallizing refractories and retaining and assim- ilating subordinate fusibles, with thick, cool lithosphere and crust. There is no need for hypotheses involving chaotic plume behavior or thousands of km of lateral flow of plume material, nor for postulates of SprimitiveT lower mantle contrary to cos- & cedil;mological and thermodynamic considerations. Plume

  9. Energy saving exhaust siphon

    SciTech Connect

    Baldwin, N.B.

    1982-04-06

    A device is disclosed for attachment to the tailpipe of an exhaust system comprising a body portion placed around the tailpipe, but spaced apart from the tailpipe, in a manner that air may easily flow between the body portion and the tailpipe when the vehicle is moving in a forward direction, a narrowing portion operative to compress the air flow, and an exhaust discharge portion operative for the exhaust from the tailpipe and the air to be discharged therethrough.

  10. Exhaust gas purification device

    SciTech Connect

    Fujiwara, H.; Hibi, T.; Sayo, S.; Sugiura, Y.; Ueda, K.

    1980-02-19

    The exhaust gas purification device includes an exhaust manifold , a purification cylinder connected with the exhaust manifold through a first honey-comb shaped catalyst, and a second honeycomb shaped catalyst positioned at the rear portion of the purification cylinder. Each catalyst is supported by steel wool rings including coarse and dense portions of steel wool. The purification device further includes a secondary air supplying arrangement.

  11. Liquid rocket engine injectors

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    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.

  12. Sirius-5 experimental rocket

    NASA Astrophysics Data System (ADS)

    Kerstein, A.; Omersel, P.; Goljuf, L.; Zidaric, M.

    1981-09-01

    After giving a historical account of multistage rocket development in Yugoslavia, a status report is presented for the three-stage Sirius-5 program. The rocket is composed of: (1) a solid-propellant first stage, consisting of a cluster of eight standard motors yielding 220 kN thrust for 1.3 sec; (2) a mixed amines/inhibited red fuming nitric acid, bipropellant second stage generating 50 kN thrust; and (3) a third stage of the same design as the second but with only 62 kg of fuel, by contrast to 168 kg. Among the design principles adhered to are: minimization of the number of components, conservative design margins, and specifications for key subsystems based on demonstration programs. The primary use of this system is in amateur rocketry, being able to carry a 20 kg payload to 150 km.

  13. Laser rocket system analysis

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    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.

  14. ISRO's solid rocket motors

    NASA Astrophysics Data System (ADS)

    Nagappa, R.; Kurup, M. R.; Muthunayagam, A. E.

    1989-08-01

    Solid rocket motors have been the mainstay of ISRO's sounding rockets and the first generation satellite launch vehicles. For the new launch vehicle under development also, the solid rocket motors contribute significantly to the vehicle's total propulsive power. The rocket motors in use and under development have been developed for a variety of applications and range in size from 30 mm dia employing 450 g of solid propellant—employed for providing a spin to the apogee motors—to the giant 2.8 m dia motor employing nearly 130 tonnes of solid propellant. The initial development, undertaken in 1967 was of small calibre motor of 75 mm dia using a double base charge. The development was essentially to understand the technological elements. Extruded aluminium tubes were used as a rocket motor casing. The fore and aft closures were machined from aluminium rods. The grain was a seven-pointed star with an enlargement of the port at the aft end and was charged into the chamber using a polyester resin system. The nozzle was a metallic heat sink type with graphite throat insert. The motor was ignited with a black powder charge and fired for 2.0 s. Subsequent to this, further developmental activities were undertaken using PVC plastisol based propellants. A class of sounding rockets ranging from 125 to 560 mm calibre were realized. These rocket motors employed improved designs and had delivered lsp ranging from 2060 to 2256 Ns/kg. Case bonding could not be adopted due to the higher cure temperatures of the plastisol propellants but improvements were made in the grain charging techniques and in the design of the igniters and the nozzle. Ablative nozzles based on asbestos phenolic and silica phenolic with graphite inserts were used. For the larger calibre rocket motors, the lsp could be improved by metallic additives. In the early 1970s designs were evolved for larger and more efficient motors. A series of 4 motors for the country's first satellite launch vehicle SLV-3 were

  15. Hybrid rocket performance

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1992-01-01

    A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model.

  16. Two-Dimensional Motions of Rockets

    ERIC Educational Resources Information Center

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

  17. Exhaust gas purifying device

    SciTech Connect

    Sakurai, S.; Hamada, S.

    1985-04-23

    An exhaust gas purifying device for use with a diesel engine comprising a filter block disposed in an engine exhaust passage for collecting exhaust gas particulates, and a heater for incinerating the collected exhaust gas particulates. The filter block has parallel channels defined therein and separated from one another by porous partition walls, some of the channels being closed at their inlet ends with blind plugs while the other channels are closed at their outlet ends with blind plugs. The heater is supported by the blind plugs.

  18. Microfabricated Liquid Rocket Motors

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

    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.

  19. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2003-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components and with appropriate adjustment of curing and other additives functionally-required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g. powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf life characteristics.

  20. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2004-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  1. Solid propellant rocket motor

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

  2. Small rocket tornado probe

    SciTech Connect

    Colgate, S.A.

    1982-01-01

    A (less than 1 lb.) paper rock tornado probe was developed and deployed in an attempt to measure the pressure, temperature, ionization, and electric field variations along a trajectory penetrating a tornado funnel. The requirements of weight and materials were set by federal regulations and a one-meter resolution at a penetration velocity of close to Mach 1 was desired. These requirements were achieved by telemetering a strain gage transducer for pressure, micro size thermister and electric field, and ionization sensors via a pulse time telemetry to a receiver on board an aircraft that digitizes a signal and presents it to a Z80 microcomputer for recording on mini-floppy disk. Recording rate was 2 ms for 8 channels of information that also includes telemetry rf field strength, magnetic field for orientation on the rocket, zero reference voltage for the sensor op amps as well as the previously mentioned items also. The absolute pressure was recorded. Tactically, over 120 h were flown in a Cessna 210 in April and May 1981, and one tornado was encountered. Four rockets were fired at this tornado, missed, and there were many equipment problems. The equipment needs to be hardened and engineered to a significant degree, but it is believed that the feasibility of the probe, tactics, and launch platform for future tornado work has been proven. The logistics of thunderstorm chasing from a remote base in New Mexico is a major difficulty and reliability of the equipment another. Over 50 dummy rockets have been fired to prove trajectories, stability, and photographic capability. Over 25 electronically equipped rockets have been fired to prove sensors transmission, breakaway connections, etc. The pressure recovery factor was calibrated in the Air Force Academy blow-down tunnel. There is a need for more refined engineering and more logistic support.

  3. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2008-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  4. Density and optical properties of SPARCS plumes

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

    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.

  5. LAMP Observes the LCROSS Plume

    NASA Video Gallery

    This video shows LAMP’s view of the LCROSS plume. The first half of the animation shows the LAMP viewport scanning across the horizon, passing through the plume, and moving on. The second half of...

  6. Hydrostatic Modeling of Buoyant Plumes

    NASA Astrophysics Data System (ADS)

    Stroman, A.; Dewar, W. K.; Wienders, N.; Deremble, B.

    2014-12-01

    The Deepwater Horizon oil spill in the Gulf of Mexico has led to increased interest in understanding point source convection dynamics. Most of the existing oil plume models use a Lagrangian based approach, which computes integral measures such as plume centerline trajectory and plume radius. However, this approach doesn't account for feedbacks of the buoyant plume on the ambient environment. Instead, we employ an Eulerian based approach to acquire a better understanding of the dynamics of buoyant plumes. We have performed a series of hydrostatic modeling simulations using the MITgcm. Our results show that there is a dynamical response caused by the presence of the buoyant plume, in that there is a modification of the background flow. We find that the buoyant plume becomes baroclinically unstable and sheds eddies at the neutral buoyancy layer. We also explore different scenarios to determine the effect of the buoyancy source and the temperature stratification on the evolution of buoyant plumes.

  7. The Fluid Dynamics of Plumes

    NASA Astrophysics Data System (ADS)

    Hansen, U.

    Plumes form as instabilities from thermal boundary layers of convecting systems.The shape, the size and the temporal evolution of plumes is strongly influenced by the vis- cosity of the material. Employing a numerical scheme the evolution of plumes in fluids with strong temperature and temperature-pressure dependent viscosity has been stud- ied. The strong dependence of viscosity on temperature leads to a pulse-like evolution of the plumes.Pulses of hot material rise episodically through the pre-established low viscosity channels. In a later stage the plumes generate extended network-like struc- tures in the thermal boundary layers. Pressure dependence of the viscosity leads to a significant cooling of the plumes.Furthe a fragmentation of one plume into several smaller ones is commonly observed. While internal convection takes place within the plume head, only little entrainment of material is observed.

  8. Validation of scramjet exhaust simulation technique at Mach 6

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    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.

  9. Contamination control and plume assessment of low-energy thrusters

    NASA Technical Reports Server (NTRS)

    Scialdone, John J.

    1993-01-01

    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.

  10. Simulation of Low-density Nozzle Plumes in Non-zero Ambient Pressures

    NASA Technical Reports Server (NTRS)

    Chung, Chan-Hong; Dewitt, Kenneth J.; Stubbs, Robert M.; Penko, Paul F.

    1994-01-01

    The direct simulation Monte-Carlo (DSMC) method was applied to the analysis of low-density nitrogen plumes exhausting from a small converging-diverging nozzle into finite ambient pressures. Two cases were considered that simulated actual test conditions in a vacuum facility. The numerical simulations readily captured the complicated flow structure of the overexpanded plumes adjusting to the finite ambient pressures, including Mach disks and barrel shaped shocks. The numerical simulations compared well to experimental data of Rothe.

  11. Enceladus' Water Vapour Plumes

    NASA Technical Reports Server (NTRS)

    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

    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.

  12. Evaluation of Visible Plumes.

    ERIC Educational Resources Information Center

    Brennan, Thomas

    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…

  13. Buoyant plume calculations

    SciTech Connect

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

    1985-01-01

    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.

  14. PLUME and research sotware

    NASA Astrophysics Data System (ADS)

    Baudin, Veronique; Gomez-Diaz, Teresa

    2013-04-01

    The PLUME open platform (https://www.projet-plume.org) has as first goal to share competences and to value the knowledge of software experts within the French higher education and research communities. The project proposes in its platform the access to more than 380 index cards describing useful and economic software for this community, with open access to everybody. The second goal of PLUME focuses on to improve the visibility of software produced by research laboratories within the higher education and research communities. The "development-ESR" index cards briefly describe the main features of the software, including references to research publications associated to it. The platform counts more than 300 cards describing research software, where 89 cards have an English version. In this talk we describe the theme classification and the taxonomy of the index cards and the evolution with new themes added to the project. We will also focus on the organisation of PLUME as an open project and its interests in the promotion of free/open source software from and for research, contributing to the creation of a community of shared knowledge.

  15. Liquid Rocket Engine Testing Overview

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  16. 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. ...

  17. Electric rockets get a boost

    SciTech Connect

    Ashley, S.

    1995-12-01

    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, chemical rocket engines still provide the only means to boost payloads into orbit and beyond. The less glamorous but equally important job of moving vessels around in space, however, may soon be assumed by a fundamentally different rocket engine technology that has been long in development--electric propulsion.

  18. German scientific sounding rocket program

    NASA Astrophysics Data System (ADS)

    Roehrig, O.

    The German scientific sounding rocket program covers four disciplines: astronomy, aeronomy, magnetosphere, material science. In each of these disciplines there are ongoing projects (e.g., INTERZODIAK, STRAFAM, MAP-WINE, CAESAR, TEXUS). The scientific and technical aspects of these projects will be described. Emphasis will be given to some late technical achievements of DFVLR's Mobile Rocket Base (MORABA) giving support to most of the rocket campaigns. DFVLR-PT is authorized to act as management agency in order to perform and to coordinate German space activities of which the sounding rocket program forms a small part. A brief description of the organization will be given.

  19. Duplex tab exhaust nozzle

    NASA Technical Reports Server (NTRS)

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

    2012-01-01

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

  20. Micro-Rockets for the Classroom.

    ERIC Educational Resources Information Center

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

    1999-01-01

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

  1. If Only Newton Had a Rocket.

    ERIC Educational Resources Information Center

    Hammock, Frank M.

    1988-01-01

    Shows how model rocketry can be included in physics curricula. Describes rocket construction, a rocket guide sheet, calculations and launch teams. Discusses the relationships of basic mechanics with rockets. (CW)

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

    NASA Technical Reports Server (NTRS)

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

    2011-01-01

    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.

  3. Exhaust gas recirculation system

    SciTech Connect

    Minoura, M.; Yorioka, K.

    1980-11-18

    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 exhaust gas pressure while the ratio of the exhaust gas flow rate to the air intake into the engine is varied in correspondence to the intake pipe negative pressure. This exhaust gas recirculation system can be adapted to a fuel injection type intake system as well as other intake systems provided with an air valve for regulating air intake or having no venturi constriction such as employed in an su type carburetor.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    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.

  5. Atmospheric Manmade Glowings Phenomena Observed During the Launches of Solid Propellant Rockets

    NASA Astrophysics Data System (ADS)

    Chernouss, S. A.; Platov, V. V.; Upspensky, M. V.; Alpatov, V. V.; Kirillov, A. S.

    2015-09-01

    Exotic types of luminosities observed in the upper atmosphere always take place during the launch and flight of solid-propellant rockets We consider a large-scale geometry and dynamic features of such phenomena also physics of the intense turquoise (blue-green) glow observed in twilight conditions in the region of missile flight. This study has been based on numerous observations of different rocket flights in the atmosphere over Russia and Scandinavia. Formation of the monoxide aluminum clouds observed in the upper atmosphere is a result of interaction of the exhausted propellant products with the atomic oxygen. The sunlight excited the monoxide aluminum EA1O*) resonance emissions in the atmosphere. Careful studies of spectra of the manmade luminosities during rocket launch/flight permit us to know chemical, thermal and mechanical processes in the atmosphere similar as it is doing in experiments with the artificial cloud release from sounding rockets in the high latitude atmosphere.

  6. Catalytic Microtube Rocket Igniter

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Deans, Matthew C.

    2011-01-01

    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

  7. Rocket + Science = Dialogue

    NASA Technical Reports Server (NTRS)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

    2014-11-01

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

  9. A three body dynamic simulation of a seated tractor rocket escape system for the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Ondler, R. M.

    1989-01-01

    In the tractor-rocket seated-extraction candidate system for Space Shuttle Orbiter crew escape, the crewmember is pulled from his seat and away from the Orbiter via an elastic pendant, using a system of rails to guide the extraction trajectory through an opening on the window frame for flight deck crew and through the side hatch for the middeck crew. A three-body simulation has been developed to model the flight-mechanics aspects of the concept, where the three bodies are the astronaut (six DOF), the tractor rocket (six DOF), and the Shuttle Orbiter (three DOF); attention is given to crewmembers' clearance of the Orbiter structure and engine plumes.

  10. ION ROCKET ENGINE

    DOEpatents

    Ehlers, K.W.; Voelker, F. III

    1961-12-19

    A thrust generating engine utilizing cesium vapor as the propellant fuel is designed. The cesium is vaporized by heat and is passed through a heated porous tungsten electrode whereby each cesium atom is fonized. Upon emergfng from the tungsten electrode, the ions are accelerated rearwardly from the rocket through an electric field between the tungsten electrode and an adjacent accelerating electrode grid structure. To avoid creating a large negative charge on the space craft as a result of the expulsion of the positive ions, a source of electrons is disposed adjacent the ion stream to neutralize the cesium atoms following acceleration thereof. (AEC)

  11. Integral rocket ramjets

    NASA Astrophysics Data System (ADS)

    Calzone, R. F.

    1994-03-01

    A rough overview of the important aspects and problem areas associated with the development of Integral Rocket Ramjet (IRR) technology is given in this report. The IRR is a supersonic air-breathing concept in which the gas generator produces fuel-rich gases. These fuel-rich gases are burnt in the secondary combustion chamber with ambient air captured and decelerated in the inlet. During the boost phase, a solid propellant booster provides the thrust necessary to achieve the velocity at which the ramjet may be operated (about M = 2). The booster is integrated in the secondary combustion chamber.

  12. General purpose rocket furnace

    NASA Technical Reports Server (NTRS)

    Aldrich, B. R.; Whitt, W. D. (Inventor)

    1979-01-01

    A multipurpose furnace for space vehicles used for material processing experiments in an outer space environment is described. The furnace contains three separate cavities designed to process samples of the widest possible range of materials and thermal requirements. Each cavity contains three heating elements capable of independent function under the direction of an automatic and programmable control system. A heat removable mechanism is also provided for each cavity which operates in conjunction with the control system for establishing an isothermally heated cavity or a wide range of thermal gradients and cool down rates. A monitoring system compatible with the rocket telemetry provides furnace performance and sample growth rate data throughout the processing cycle.

  13. An Injector for the Variable Specific Impulse Magnetoplasma Rocket

    NASA Astrophysics Data System (ADS)

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

    1997-11-01

    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.

  14. Soda-Bottle Water Rockets.

    ERIC Educational Resources Information Center

    Kagan, David; And Others

    1995-01-01

    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)

  15. Otrag rocket experiments in Africa

    NASA Technical Reports Server (NTRS)

    1978-01-01

    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.

  16. Coal-Fired Rocket Engine

    NASA Technical Reports Server (NTRS)

    Anderson, Floyd A.

    1987-01-01

    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.

  17. Chemical plume source localization.

    PubMed

    Pang, Shuo; Farrell, Jay A

    2006-10-01

    This paper addresses the problem of estimating a likelihood map for the location of the source of a chemical plume using an autonomous vehicle as a sensor probe in a fluid flow. The fluid flow is assumed to have a high Reynolds number. Therefore, the dispersion of the chemical is dominated by turbulence, resulting in an intermittent chemical signal. The vehicle is capable of detecting above-threshold chemical concentration and sensing the fluid flow velocity at the vehicle location. This paper reviews instances of biological plume tracing and reviews previous strategies for a vehicle-based plume tracing. The main contribution is a new source-likelihood mapping approach based on Bayesian inference methods. Using this Bayesian methodology, the source-likelihood map is propagated through time and updated in response to both detection and nondetection events. Examples are included that use data from in-water testing to compare the mapping approach derived herein with the map derived using a previously existing technique. PMID:17036813

  18. Reusable sounding-rocket design

    NASA Astrophysics Data System (ADS)

    Woo, Dick L. Y.; Martin, James A.

    1995-03-01

    As a result of the reduction of budgets for flights, the ideas of reusability and cost-effectiveness in launch vehicles are becoming more and more important. One class of rockets, in particular the sounding rockets operating in a less demanding environment, has many potentials for many more flights. By augmenting the basic rocket configuration with wings, landing gear, flight controls and guidance systems, the vehicle can be made to glide and land back at the launch site or at a specific recovery site. In this paper, the design of such a reusable rocket is presented. This design can be extended and adapted to larger vehicles, thus attaining higher altitudes required in some of the applications of sounding rockets.

  19. Modeling the chemical effects of ship exhaust in the cloud-free marine boundary layer

    NASA Astrophysics Data System (ADS)

    von Glasow, R.; Lawrence, M. G.; Sander, R.; Crutzen, P. J.

    2003-02-01

    The chemical evolution of the exhaust plumes of ocean-going ships in the cloud-free marine boundary layer is examined with a box model. Dilution of the ship plume via entrainment of background air was treated as in studies of aircraft emissions and was found to be a very important process that significantly alters model results. We estimated the chemical lifetime (defined as the time when differences between plume and background air are reduced to 5% or less) of the exhaust plume of a single ship to be 2 days. Increased concentrations of NOx (= NO + NO2) in the plume air lead to higher catalytical photochemical production rates of O3 and also of OH. Due to increased OH concentrations in the plume, the lifetime of many species (especially NOx) is significantly reduced in plume air. The chemistry on background aerosols has a significant effect on gas phase chemistry in the ship plume, while partly soluble ship-produced aerosols are computed to only have a very small effect. Soot particles emitted by ships showed no effect on gas phase chemistry. Halogen species that are released from sea salt aerosols are slightly increased in plume air. In the early plume stages, however, the mixing ratio of BrO is decreased because it reacts rapidly with NO. To study the global effects of ship emissions we used a simple upscaling approach which suggested that the parameterization of ship emissions in global chemistry models as a constant source at the sea surface leads to an overestimation of the effects of ship emissions on O3 of about 50% and on OH of roughly a factor of 2. The differences are mainly caused by a strongly reduced lifetime (compared to background air) of NOxin the early stages of a ship plume.

  20. Modeling the chemical effects of ship exhaust in the cloud-free marine boundary layer

    NASA Astrophysics Data System (ADS)

    von Glasow, R.; Lawrence, M. G.; Sander, R.; Crutzen, P. J.

    2002-06-01

    The chemical evolution of the exhaust plumes of ocean-going ships in the cloud-free marine boundary layer is examined with a box model. Dilution of the ship plume via entrainment of background air was treated as in studies of aircraft emissions and was found to be a very important process that significantly alters model results. We estimated the chemical lifetime (defined as the time when differences between plume and background air are reduced to 5% or less) of the exhaust plume of a single ship to be 2 days. Increased concentrations of NOx in the plume air lead to higher catalytical photochemical production rates of O3 and also of OH. Due to increased OH concentrations in the plume, the lifetime of many species (especially NOx) is reduced in plume air. The chemistry on background aerosols has a significant effect on gas phase chemistry in the ship plume, while partly soluble ship-produced aerosols are computed to only have a very small effect. Soot particles emitted by ships showed no effect on gas phase chemistry. Halogen species that are released from sea salt aerosols are slightly increased in plume air. In the early plume stages, however, the mixing ratio of BrO is decreased because it reacts rapidly with NO. To study the global effects of ship emissions we used a simple upscaling approach which suggested that the parameterization of ship emissions in global chemistry models as a constant source at the sea surface leads to an overestimation of the effects of ship emissions of roughly a factor of 2. The differences are caused by a strongly reduced lifetime (compared to background air) of NOx in the early stages of a ship plume.

  1. Rhenium Rocket Manufacturing Technology

    NASA Technical Reports Server (NTRS)

    1997-01-01

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

  2. Mars Rocket Propulsion System

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Harber, Dan; Nabors, Sammy

    2008-01-01

    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.

  3. Flow field for an underexpanded, supersonic nozzle exhausting into an expansive launch tube

    NASA Technical Reports Server (NTRS)

    Morris, R. R.; Bertin, J. J.; Batson, J. L.

    1976-01-01

    Static pressure distributions along the launcher wall and pitot pressure measurements from the annular region between the rocket and the launcher were made as an underexpanded supersonic nozzle exhausted into an expansive launch tube. The flow remained supersonic along the entire length of the launcher for all nozzle locations studied.

  4. Impact of aircraft plume dynamics on airport local air quality

    NASA Astrophysics Data System (ADS)

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

    2013-08-01

    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.

  5. 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....

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

    PubMed

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

    2010-08-01

    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.

  7. Subscale Validation of the Subsurface Active Filtration of Exhaust (SAFE) Approach to the NTP Ground Testing

    NASA Technical Reports Server (NTRS)

    Marshall, William M.; Borowski, Stanley K.; Bulman, Mel; Joyner, Russell; Martin, Charles R.

    2015-01-01

    Nuclear thermal propulsion (NTP) has been recognized as an enabling technology for missions to Mars and beyond. However, one of the key challenges of developing a nuclear thermal rocket is conducting verification and development tests on the ground. A number of ground test options are presented, with the Sub-surface Active Filtration of Exhaust (SAFE) method identified as a preferred path forward for the NTP program. The SAFE concept utilizes the natural soil characteristics present at the Nevada National Security Site to provide a natural filter for nuclear rocket exhaust during ground testing. A validation method of the SAFE concept is presented, utilizing a non-nuclear sub-scale hydrogen/oxygen rocket seeded with detectible radioisotopes. Additionally, some alternative ground test concepts, based upon the SAFE concept, are presented. Finally, an overview of the ongoing discussions of developing a ground test campaign are presented.

  8. Image Analysis Based Estimates of Regolith Erosion Due to Plume Impingement Effects

    NASA Technical Reports Server (NTRS)

    Lane, John E.; Metzger, Philip T.

    2014-01-01

    Characterizing dust plumes on the moon's surface during a rocket landing is imperative to the success of future operations on the moon or any other celestial body with a dusty or soil surface (including cold surfaces covered by frozen gas ice crystals, such as the moons of the outer planets). The most practical method of characterizing the dust clouds is to analyze video or still camera images of the dust illuminated by the sun or on-board light sources (such as lasers). The method described below was used to characterize the dust plumes from the Apollo 12 landing.

  9. Computational studies of hard-body and 3-D effects in plume flows

    NASA Technical Reports Server (NTRS)

    Venkatapathy, Ethiraj; Feiereisen, William J.; Obayashi, Shigeru

    1989-01-01

    Axisymmetric and three-dimensional, multi-nozzle plume flows around generic rocket geometries are investigated with a three-dimensional Navier-Stokes solver to study the interactive effects between hard body and the plume. Time-asymptotic, laminar, ideal-gas solutions obtained with a two-factor, flux-split scheme and a diagonal, upwind scheme are presented. Computed solutions to three-dimensional, multi-nozzle problems and single-nozzle, axisymmetric problems demonstrate flow field features including three-dimensionality and hard-body effects. Geometry and three-dimensional effects are shown to be significant in multi-nozzle flows.

  10. Development of high performance hybrid rocket fuels

    NASA Astrophysics Data System (ADS)

    Zaseck, Christopher R.

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

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

    NASA Astrophysics Data System (ADS)

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

    1992-02-01

    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

  12. Analysis of rocket beacon transmissions for computerized reconstruction of ionospheric densities

    NASA Technical Reports Server (NTRS)

    Bernhardt, P. A.; Huba, J. D.; Chaturvedi, P. K.; Fulford, J. A.; Forsyth, P. A.; Anderson, D. N.; Zalesak, S. T.

    1993-01-01

    Three methods are described to obtain ionospheric electron densities from transionospheric, rocket-beacon TEC data. First, when the line-of-sight from a ground receiver to the rocket beacon is tangent to the flight trajectory, the electron concentration can be obtained by differentiating the TEC with respect to the distance to the rocket. A similar method may be used to obtain the electron-density profile if the layer is horizontally stratified. Second, TEC data obtained during chemical release experiments may be interpreted with the aid of physical models of the disturbed ionosphere to yield spatial maps of the modified regions. Third, computerized tomography (CT) can be used to analyze TEC data obtained along a chain of ground-based receivers aligned along the plane of the rocket trajectory. CT analysis of TEC data is used to reconstruct a 2D image of a simulated equatorial plume. TEC data is computed for a linear chain of nine receivers with adjacent spacings of either 100 or 200 km. The simulation data are analyzed to provide an F region reconstruction on a grid with 15 x 15 km pixels. Ionospheric rocket tomography may also be applied to rocket-assisted measurements of amplitude and phase scintillations and airglow intensities.

  13. Upwelling relaxation and estuarine plumes

    NASA Astrophysics Data System (ADS)

    Rao, Shivanesh; Pringle, James; Austin, Jay

    2011-09-01

    After coastal upwelling, the water properties in the nearshore coastal region close to estuaries is determined by the race between the new estuarine plume traveling along the coast and the upwelled front (a marker for the old upwelled plume and the coastal pycnocline) returning to the coast under downwelling winds. Away from an estuary, downwelling winds can return the upwelled front to the coast bringing less dense water nearshore. Near the estuary, the estuarine plume can arrive along the coast and return less dense water to the nearshore region before the upwelled front returns to the coast. Where the plume brings less dense water to the coast first, the plume keeps the upwelled front from returning to the coast. In this region, only the plume and the anthropogenic input and larvae associated with the plume waters influence the nearshore after upwelling. We quantify the extent of the region where the plume is responsible for bringing less dense water to the nearshore and keeping the upwelled front from returning to the coast after upwelling. We successfully tested our predictions against numerical experiments and field observations of the Chesapeake plume near Duck, North Carolina. We argue that this alongshore region exists for other estuaries where the time-integrated upwelling and downwelling wind stresses are comparable.

  14. Structure of axisymmetric mantle plumes

    NASA Technical Reports Server (NTRS)

    Olson, Peter; Schubert, Gerald; Anderson, Charles

    1993-01-01

    The structure of axisymmetric subsolidus thermal plumes in the earth's lower mantle is inferred from calculations of axisymmetric thermal plumes in an infinite Prandtl number fluid with thermally activated viscosity. The velocity and temperature distribution is determined for axisymmetric convection above a heated disk in an incompressible fluid cylinder 2,400 km in height and 1,200 km in diameter. Several calculations of plumes with heat transport in the range 100-400 GW, similar to the advective heat transport at the Hawaiian hotspot, are presented. Hotspot formation by plumes originating at the base of the mantle requires both large viscosity variations and a minimum heat transport.

  15. Exergy Analysis of Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, Andrew; Mesmer, Bryan; Watson, Michael D.

    2015-01-01

    Exergy is defined as the useful work available from a system in a specified environment. Exergy analysis allows for comparison between different system designs, and allows for comparison of subsystem efficiencies within system designs. The proposed paper explores the relationship between the fundamental rocket equation and an exergy balance equation. A previously derived exergy equation related to rocket systems is investigated, and a higher fidelity analysis will be derived. The exergy assessments will enable informed, value-based decision making when comparing alternative rocket system designs, and will allow the most efficient configuration among candidate configurations to be determined.

  16. Dynamic characterization of solid rockets

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The structural dynamics of solid rockets in-general was studied. A review is given of the modes of vibration and bending that can exist for a solid propellant rocket, and a NASTRAN computer model is included. Also studied were the dynamic properties of a solid propellant, polybutadiene-acrylic acid-acrylonitrile terpolymer, which may be used in the space shuttle rocket booster. The theory of viscoelastic materials (i.e, Poisson's ratio) was employed in describing the dynamic properties of the propellant. These studies were performed for an eventual booster stage development program for the space shuttle.

  17. Liquid Rocket Boosters for Shuttle

    NASA Astrophysics Data System (ADS)

    Hughes, James E.

    The Liquid Rocket Booster study was initiated by NASA to define an alternative to the Solid Rocket Boosters used on the STS. These studies have involved MSFC, JSC and KSC and their contractors. The prime study contractors, Martin Marietta Corporation and General Dynamics Space Systems, have identified Liquid Booster configurations which would replace the SRB's in the Shuttle stack. The Liquid Rocket Booster increases Shuttle performance to 70 K LBS, provides improved reliability, hold down and verification prior to vehicle release, engine out and improved abort capability, and is phased into the STS launch operations without adversely affecting flight rate.

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

    NASA Technical Reports Server (NTRS)

    Stephens, J. B. (Editor)

    1974-01-01

    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.

  19. Dr. Robert H. Goddard and His Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    Goddard rocket with four rocket motors. This rocket attained an altitude of 200 feet in a flight, November 1936, at Roswell, New Mexico. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

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

    NASA Astrophysics Data System (ADS)

    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

    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.

  1. Monte-Carlo particle dynamics in a variable specific impulse magnetoplasma rocket

    SciTech Connect

    Ilin, A.V.; Diaz, F.R.C.; Squire, J.P.

    1999-01-01

    The self-consistent mathematical model in a Variable Specific Impulse Magnetoplasma Rocket (VASIMR) is examined. Of particular importance is the effect of a magnetic nozzle in enhancing the axial momentum of the exhaust. Also, different geometries and rocket symmetries are considered. The magnetic configuration is modeled with an adaptable mesh, which increases accuracy without compromising the speed of the simulation. The single particle trajectories are integrated with an adaptive time-scheme, which can quickly solve extensive Monte-Carlo simulations for systems of hundred thousands of particles in a reasonable time (1--2 hours) and without the need for a powerful supercomputer.

  2. Exhaust gas recirculator

    SciTech Connect

    Suda, K.

    1983-01-04

    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.

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

    NASA Technical Reports Server (NTRS)

    1972-01-01

    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.

  4. 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...

  5. 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...

  6. Solid rocket motor internal insulation

    NASA Technical Reports Server (NTRS)

    Twichell, S. E. (Editor); Keller, R. B., Jr.

    1976-01-01

    Internal insulation in a solid rocket motor is defined as a layer of heat barrier material placed between the internal surface of the case propellant. The primary purpose is to prevent the case from reaching temperatures that endanger its structural integrity. Secondary functions of the insulation are listed and guidelines for avoiding critical problems in the development of internal insulation for rocket motors are presented.

  7. Solid Rocket Motor Acoustic Testing

    SciTech Connect

    Rogers, J.D.

    1999-03-31

    Acoustic data are often required for the determination of launch and powered flight loads for rocket systems and payloads. Such data are usually acquired during test firings of the solid rocket motors. In the current work, these data were obtained for two tests at a remote test facility where we were visitors. This paper describes the data acquisition and the requirements for working at a remote site, interfacing with the test hosts.

  8. Atmospheric chemistry in volcanic plumes

    PubMed Central

    von Glasow, Roland

    2010-01-01

    Recent field observations have shown that the atmospheric plumes of quiescently degassing volcanoes are chemically very active, pointing to the role of chemical cycles involving halogen species and heterogeneous reactions on aerosol particles that have previously been unexplored for this type of volcanic plumes. Key features of these measurements can be reproduced by numerical models such as the one employed in this study. The model shows sustained high levels of reactive bromine in the plume, leading to extensive ozone destruction, that, depending on plume dispersal, can be maintained for several days. The very high concentrations of sulfur dioxide in the volcanic plume reduces the lifetime of the OH radical drastically, so that it is virtually absent in the volcanic plume. This would imply an increased lifetime of methane in volcanic plumes, unless reactive chlorine chemistry in the plume is strong enough to offset the lack of OH chemistry. A further effect of bromine chemistry in addition to ozone destruction shown by the model studies presented here, is the oxidation of mercury. This relates to mercury that has been coemitted with bromine from the volcano but also to background atmospheric mercury. The rapid oxidation of mercury implies a drastically reduced atmospheric lifetime of mercury so that the contribution of volcanic mercury to the atmospheric background might be less than previously thought. However, the implications, especially health and environmental effects due to deposition, might be substantial and warrant further studies, especially field measurements to test this hypothesis. PMID:20368458

  9. Atmospheric chemistry in volcanic plumes.

    PubMed

    von Glasow, Roland

    2010-04-13

    Recent field observations have shown that the atmospheric plumes of quiescently degassing volcanoes are chemically very active, pointing to the role of chemical cycles involving halogen species and heterogeneous reactions on aerosol particles that have previously been unexplored for this type of volcanic plumes. Key features of these measurements can be reproduced by numerical models such as the one employed in this study. The model shows sustained high levels of reactive bromine in the plume, leading to extensive ozone destruction, that, depending on plume dispersal, can be maintained for several days. The very high concentrations of sulfur dioxide in the volcanic plume reduces the lifetime of the OH radical drastically, so that it is virtually absent in the volcanic plume. This would imply an increased lifetime of methane in volcanic plumes, unless reactive chlorine chemistry in the plume is strong enough to offset the lack of OH chemistry. A further effect of bromine chemistry in addition to ozone destruction shown by the model studies presented here, is the oxidation of mercury. This relates to mercury that has been coemitted with bromine from the volcano but also to background atmospheric mercury. The rapid oxidation of mercury implies a drastically reduced atmospheric lifetime of mercury so that the contribution of volcanic mercury to the atmospheric background might be less than previously thought. However, the implications, especially health and environmental effects due to deposition, might be substantial and warrant further studies, especially field measurements to test this hypothesis.

  10. Midwave infrared imaging Fourier transform spectrometry of combustion plumes

    NASA Astrophysics Data System (ADS)

    Bradley, Kenneth C.

    A midwave infrared (MWIR) imaging Fourier transform spectrometer (IFTS) was used to successfully capture and analyze hyperspectral imagery of combustion plumes. Jet engine exhaust data from a small turbojet engine burning diesel fuel at a low rate of 300 cm3/min was collected at 1 cm -1 resolution from a side-plume vantage point on a 200x64 pixel window at a range of 11.2 meters. Spectral features of H2O, CO, and CO2 were present, and showed spatial variability within the plume structure. An array of thermocouple probes was positioned within the plume to aid in temperature analysis. A single-temperature plume model was implemented to obtain spatially-varying temperatures and plume concentrations. Model-fitted temperatures of 811 +/- 1.5 K and 543 +/- 1.6 K were obtained from plume regions in close proximity to thermocouple probes measuring temperatures of 719 K and 522 K, respectively. Industrial smokestack plume data from a coal-burning stack collected at 0.25 cm-1 resolution at a range of 600 meters featured strong emission from NO, CO, CO2, SO 2, and HCl in the spectral region 1800-3000 cm-1. A simplified radiative transfer model was employed to derive temperature and concentrations for clustered regions of the 128x64 pixel scene, with corresponding statistical error bounds. The hottest region (closest to stack centerline) was 401 +/- 0.36 K, compared to an in-stack measurement of 406 K, and model-derived concentration values of NO, CO2, and SO2 were 140 +/- 1 ppmV, 110,400 +/- 950 ppmV, and 382 +/- 4 ppmV compared to in-stack measurements of 120 ppmV (NOx), 94,000 ppmV, and 382 ppmV, respectively. In-stack measurements of CO and HCl were not provided by the stack operator, but model-derived values of 19 +/- 0.2 ppmV and 111 +/- 1 ppmV are reported near stack centerline. A deployment to Dugway Proving Grounds, UT to collect hyperspectral imagery of chemical and biological threat agent simulants resulted in weak spectral signatures from several species. Plume

  11. Design of a Novel Gaseous Hydrogen-Oxygen Rocket Injector Element

    NASA Technical Reports Server (NTRS)

    Glenn, Dennis E.

    1999-01-01

    NASA and Aerojet are developing a Rocket-Based Combined Cycle (RBCC) engine under the Advanced Reusable Technology program. The rocket application requires that the combustion process be stable, complete, and take place in as short a distance as possible without compromising the structural integrity of the injector itself. A novel gaseous hydrogen-oxygen rocket injector element design was arrived at through an iterative design process making extensive use of CFD simulations, which resulted in a design that is meeting design goals. Sub-scale versions of the injector have been built and tested in a unique test-rig and in a sub-scale RBCC engine. The Aerojet RBCC concept integrates small rocket thrusters into the rearfacing base area of struts placed in the flowpath of a scramjet (Supersonic Combusting Ramjet) engine. In one mode of operation, at vehicle takeoff, the rockets provide the primary thrust with additional thrust coming from an ejector effect as air is drawn into the engine inlet, entrained, and accelerated by the rocket exhaust.

  12. An experimental investigation of an arcjet thruster exhaust using Langmuir probes. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Carney, Lynnette M.

    1988-01-01

    Electrostatic (Langmuir) probes of both spherical and cylindrical geometry have been used to obtain electron number density and temperature in the exhaust of a laboratory arcjet. The arcjet thruster operated on nitrogen and hydrogen mixtures to simulate fully decomposed hydrazine in a vacuum environment with background pressures less than 5 x 10 to the -2 Pa. The exhaust appears to be only slightly ionized (less than 1 percent) with local plasma potentials near facility ground. The current-voltage characteristics of the probes indicate a Maxwellian temperature distribution. Plume data are presented as a function of arcjet operating condition and also position in the exhaust.

  13. Apollo Video Photogrammetry Estimation Of Plume Impingement Effects

    NASA Technical Reports Server (NTRS)

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

    2008-01-01

    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.

  14. Measurement and analysis of a small nozzle plume in vacuum

    NASA Technical Reports Server (NTRS)

    Penko, P. F.; Boyd, I. D.; Meissner, D. L.; Dewitt, K. J.

    1993-01-01

    Pitot pressures and flow angles are measured in the plume of a nozzle flowing nitrogen and exhausting to a vacuum. Total pressures are measured with Pitot tubes sized for specific regions of the plume and flow angles measured with a conical probe. The measurement area for total pressure extends 480 mm (16 exit diameters) downstream of the nozzle exit plane and radially to 60 mm (1.9 exit diameters) off the plume axis. The measurement area for flow angle extends to 160 mm (5 exit diameters) downstream and radially to 60 mm. The measurements are compared to results from a numerical simulation of the flow that is based on kinetic theory and uses the direct-simulation Monte Carlo (DSMC) method. Comparisons of computed results from the DSMC method with measurements of flow angle display good agreement in the far-field of the plume and improve with increasing distance from the exit plane. Pitot pressures computed from the DSMC method are in reasonably good agreement with experimental results over the entire measurement area.

  15. Numerical investigations in the backflow region of a vacuum plume

    NASA Technical Reports Server (NTRS)

    Liaw, Goang-Shin

    1995-01-01

    Four tasks were completed in this period and results were published in AIAA papers. First, a Boltzmann-2D code, was developed and applied to compute MSFC-A2 nozzle/plume flow field. It solved the two-dimensional Boltzmann-BGK equation using the Finite Difference Discrete Ordinate (FDDO) numerical technique. The code was validated by experimental data for one-dimensional shock structure predictions, paper 95-2056. Successful results for nozzle/plume flow simulation using the developed Boltzmann-2D code were presented at the 1995 AIAA Aerospace Science Conference, paper 95-0627. Second, a computer code solving two-dimensional Burnett equations was developed and applied to low-density nozzle flow field calculation. Results were also published at the 1994 AIAA Thermophysics Conference, paper 94-2055. Third, the developed two-dimensional Burnett code was extended to compute axisymmetric flow field inside MSFC-A2 nozzle, paper 95-2008. The computed nozzle exit conditions are used as input data for Direct Simulation Monte Carlo (DSMC) plume calculation. Fourth, a DSMC code was modified to compute the exhausted plume near the nozzle exit and in the backflow region.

  16. Numerical investigations in the backflow region of a vacuum plume

    NASA Astrophysics Data System (ADS)

    Liaw, Goang-Shin

    1995-08-01

    Four tasks were completed in this period and results were published in AIAA papers. First, a Boltzmann-2D code, was developed and applied to compute MSFC-A2 nozzle/plume flow field. It solved the two-dimensional Boltzmann-BGK equation using the Finite Difference Discrete Ordinate (FDDO) numerical technique. The code was validated by experimental data for one-dimensional shock structure predictions, paper 95-2056. Successful results for nozzle/plume flow simulation using the developed Boltzmann-2D code were presented at the 1995 AIAA Aerospace Science Conference, paper 95-0627. Second, a computer code solving two-dimensional Burnett equations was developed and applied to low-density nozzle flow field calculation. Results were also published at the 1994 AIAA Thermophysics Conference, paper 94-2055. Third, the developed two-dimensional Burnett code was extended to compute axisymmetric flow field inside MSFC-A2 nozzle, paper 95-2008. The computed nozzle exit conditions are used as input data for Direct Simulation Monte Carlo (DSMC) plume calculation. Fourth, a DSMC code was modified to compute the exhausted plume near the nozzle exit and in the backflow region.

  17. Scanning thermal plumes

    NASA Technical Reports Server (NTRS)

    Scarpace, F. L.; Madding, R. P.; Green, T., III

    1975-01-01

    Over a three-year period 800 thermal line scans of power plant plumes were made by an airborne scanner, with ground truth measured concurrently at the plants. Computations using centered finite differences in the thermal scanning imagery show a lower bound in the horizontal temperature gradient in excess of 1.6 C/m. Gradients persist to 3 m below the surface. Vector plots of the velocity of thermal fronts are constructed by tracing the front motion in successive thermal images. A procedure is outlined for the two-point ground calibration of a thermal scanner from an equation describing the scanner signal and the voltage for two known temperatures. The modulation transfer function is then calculated by input of a thermal step function and application of digital time analysis techniques using Fast Fourier Transforms to the voltage output. Field calibration tests are discussed. Data accuracy is limited by the level of ground truth effort chosen.

  18. Improved hybrid rocket fuel

    NASA Technical Reports Server (NTRS)

    Dean, David L.

    1995-01-01

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

  19. EUVS Sounding Rocket Payload

    NASA Technical Reports Server (NTRS)

    Stern, Alan S.

    1996-01-01

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

  20. Exhaust bypass flow control for exhaust heat recovery

    SciTech Connect

    Reynolds, Michael G.

    2015-09-22

    An exhaust system for an engine comprises an exhaust heat recovery apparatus configured to receive exhaust gas from the engine and comprises a first flow passage in fluid communication with the exhaust gas and a second flow passage in fluid communication with the exhaust gas. A heat exchanger/energy recovery unit is disposed in the second flow passage and has a working fluid circulating therethrough for exchange of heat from the exhaust gas to the working fluid. A control valve is disposed downstream of the first and the second flow passages in a low temperature region of the exhaust heat recovery apparatus to direct exhaust gas through the first flow passage or the second flow passage.

  1. Diesel engine exhaust

    Integrated Risk Information System (IRIS)

    Diesel engine exhaust ; CASRN N.A . Human health assessment information on a chemical substance is included in the IRIS database only after a comprehensive review of toxicity data , as outlined in the IRIS assessment development process . Sections I ( Health Hazard Assessments for Noncarcinogenic Ef

  2. Hybrid Exhaust Component

    NASA Technical Reports Server (NTRS)

    Pelletier, Gerard D. (Inventor); Logan, Charles P. (Inventor); McEnerney, Bryan William (Inventor); Haynes, Jeffrey D. (Inventor)

    2015-01-01

    An exhaust includes a wall that has a first composite material having a first coefficient of thermal expansion and a second composite material having a second coefficient of the thermal expansion that is less than the first coefficient of thermal expansion.

  3. Characteristics of an electron-beam rocket pellet accelerator

    SciTech Connect

    Tsai, C.C.; Foster, C.A.; Milora, S.L.; Schechter, D.E.

    1991-01-01

    A proof-of-principle (POP) electron-beam pellet accelerator has been developed and used for accelerating hydrogen and deuterium pellets. An intact hydrogen pellet was accelerated to a speed of 460 m/s by an electron beam of 13.5 keV. 0.3 A, and 2 ms. The maximum speed is limited by the acceleration path length (0.4 m) and pellet integrity. Experimental data have been collected for several hundred hydrogen pellets, which were accelerated by electron beams with parameters of voltage up to 16 kV, current up to 0.4 A, and pulse length up to 10 ms. Preliminary results reveal that the measured burn velocity increases roughly with the square of the beam voltage, as the theoretical model predicts. The final pellet velocity is proportional to the exhaust velocity, which increases with the beam power. To reach the high exhaust velocity needed for accelerating pellets to >1000 m/s, a new electron gun, with its cathode indirectly heated by a graphite heater and an electron beam, is being developed to increase beam current and power. A rocket casing or shell around the pellet has been designed and developed to increase pellet strength and improve the electron-rocket coupling efficiency. We present the characteristics of this pellet accelerator, including new improvements. 13 refs., 6 figs.

  4. Rocket Science 101 Interactive Educational Program

    NASA Technical Reports Server (NTRS)

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

    2007-01-01

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

  5. Flow fields of low pressure vent exhausts

    NASA Technical Reports Server (NTRS)

    Scialdone, John J.

    1989-01-01

    The flow field produced by low pressure gas vents are described based on experimental data obtained from tests in a large vacuum chamber. The gas density, pressure, and flux at any location in the flow field are calculated based on the vent plume description and the knowledge of the flow rate and velocity of the venting gas. The same parameters and the column densities along a specified line of sight traversing the plume are also obtained and shown by a computer-generated graphical representation. The fields obtained with a radially scanning Pitot probe within the exhausting gas are described by a power of the cosine function, the mass rate and the distance from the exit port. The field measurements were made for gas at pressures ranging from 2 to 50 torr venting from pipe fittings with diameters of 3/16 inch to 1-1/2 inches I.D. (4.76 mm to 38.1 mm). The N(2) mass flow rates ranged from 2E-4 to 3.7E-1 g/s.

  6. Anchoring Atmospheric Density Models Using Observed Shuttle Plume Emissions

    NASA Astrophysics Data System (ADS)

    Dimpfl, W. L.; Bernstien, L. S.

    2010-12-01

    Atmospheric number densities at a given low-earth orbit (LEO) altitude can vary by more than an order of magnitude, depending on such parameters as diurnal variations and solar activity. The MSIS atmospheric model, which includes these dependent variables as input, is reported as being accurate to ±15%. Improvement to such models requires accurate direct atmospheric measurement. Here, a means of anchoring atmospheric models is offered through measuring the size and shape of atomic line or molecular band radiance resulting from the atmospheric interaction from rocket engine plumes or gas releases in LEO. Many discrete line or band emissions, ranging from the infrared to the ultraviolet may be suitable. For this purpose we are focusing on NH(A→X), centered at 316 nm. This emission is seen in the plumes of the Shuttle Orbiter PRCS engines, is expected in the plume of any amine fueled engine, and can be observed from remote sensors in space or on the ground. The atmospheric interaction of gas releases or plumes from spacecraft in LEO are understood by comparison of observed radiance with that predicted by Direct Simulation Monte Carlo (DSMC) models. The recent Extended Variable Hard Sphere (EVHS) improvements in treating hyperthermal collisions has produced exceptional agreement between measured and modeled steady-state Space Shuttle OMS and PRCS 190-250 nm Cameron band plume radiance from CO(a→X), which is understood to result from a combination of two- and three-step mechanisms. Radiance from NH(A→X) in far field plumes is understood to result from a simpler single-step process of the reaction of a minor plume species with atomic oxygen, making it more suitable for use in determining atmospheric density. It is recommended that direct retrofire burns of amine fueled engines be imaged in a narrow band from remote sensors to reveal atmospheric number density. In principal the simple measurement of the distance between the engine exit and the peak in the steady

  7. Analysis of Nozzle Jet Plume Effects on Sonic Boom Signature

    NASA Technical Reports Server (NTRS)

    Bui, Trong

    2010-01-01

    An axisymmetric full Navier-Stokes computational fluid dynamics (CFD) study was conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock (LaNCETS) research airplane, was considered. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature, similar to what was observed in the LaNCETS flight data, is observed for the highly underexpanded nozzle flow. The CFD results provide a detailed description of the nozzle flow physics involved in the LaNCETS nozzle at different nozzle expansion conditions and help in interpreting LaNCETS flight data as well as in the eventual CFD analysis of a full LaNCETS aircraft. The current study also provided important information on proper modeling of the LaNCETS aircraft nozzle. The primary objective of the current CFD research effort was to support the LaNCETS flight research data analysis effort by studying the detailed nozzle exhaust jet plume s imperfect expansion effects on the sonic boom signature of a supersonic aircraft. Figure 1 illustrates the primary flow physics present in the interaction between the exhaust jet plume shock and the sonic boom coming off of an axisymmetric body in supersonic flight. The steeper tail shock from highly expanded jet plume reduces the dip of the sonic boom N-wave signature. A structured finite-volume compressible full Navier-Stokes CFD code was used in the current study. This approach is not limited by the simplifying assumptions inherent in previous sonic boom analysis efforts. Also, this study was the first known jet plume sonic boom CFD study in which the full viscous nozzle flow field was modeled, without

  8. Thrust Performance Improvement for a Water/Liquid Nitrogen Rocket Engine

    NASA Astrophysics Data System (ADS)

    Watanabe, Rikio; Mikami, Ryo

    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.

  9. Marshall Team Recreates Goddard Rocket

    NASA Technical Reports Server (NTRS)

    2003-01-01

    In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

  10. A Matter of Energy Efficiency: From Modeling Falling Raindrops to Controlling Rocket Thrust

    NASA Astrophysics Data System (ADS)

    Harding, Eric

    2014-03-01

    A power-law accretion model is used to investigate the energy dynamics of a falling raindrop in a Newtonian gravitational field where air resistance is included in the analysis. This model is seen to be related to the rate of ejection of exhaust gases for the rocket problem. Energy analysis of the falling raindrop will be presented for the motion of very small droplets, those of diameter less than 0.003 inches, which are falling at relatively slow speeds, of less than 0.188 m/s. The deviation from self-similar accretion, and other relevant model parameters will be re-interpreted as related to control parameters in the rocket problem. Efficiency of natural energy transfer for the falling raindrop will be compared with the power transfer model for the rocket. Acknowledge support from Dept. of Physics, Engineering and Geoscience, Noyce TPOD-STEM Grant, GT STEP Grant.

  11. Nuclear thermal rockets using indigenous extraterrestrial propellants

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert M.

    1990-01-01

    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.

  12. 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 ...

  13. Rocket Combustion Chambers Resist Thermal Fatigue

    NASA Technical Reports Server (NTRS)

    Kazaroff, John M.; Jankovsky, Robert S.; Pavli, Albert J.

    1995-01-01

    Improved design concept developed for combustion chambers for rocket engines, described in three reports. Provides compliance allowing unrestrained thermal expansion in circumferential direction. Compliance lengthens life of rocket engine by reducing amount of thermal deformation caused by repeated firings.

  14. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1986-01-01

    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. The author has attempted to make this a self-contained report. 57 refs., 10 figs.

  15. Rocket Science at the Nanoscale.

    PubMed

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale.

  16. Emergency egress fixed rocket package

    NASA Technical Reports Server (NTRS)

    Allen, Margaret A. (Inventor)

    1989-01-01

    A method of effecting the in-flight departure of an astronaut from a shuttle craft, and apparatus is presented. A plurality of removeable compartment covers are provided, behind which rocket assemblies are stowed. To actuate the system, the astronaut pulls off a tab from one of the compartments which exposes a cannister having a lanyard with a hook. The lanyard extends around a spring biased sleeve with a safety lever preventing rocket ignition until the hook is moved by the astronaut. Upward movement of the hook allows the trigger mechanism to actuate the system resulting in the rods projecting out of the hatch. When the lanyard becomes taut, a lanyard elongation detector transmits a signal to the firing mechanisms to fire the rocket.

  17. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1987-03-01

    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. The author has attempted to make this a self-contained report.

  18. Rocket Science at the Nanoscale.

    PubMed

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale. PMID:27219742

  19. Automated Rocket Propulsion Test Management

    NASA Technical Reports Server (NTRS)

    Walters, Ian; Nelson, Cheryl; Jones, Helene

    2007-01-01

    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.

  20. Saving Lives With Rocket Power

    NASA Technical Reports Server (NTRS)

    2000-01-01

    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.