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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. 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.

  4. First Passively-illuminated, High-resolution Polarimetric Images of Exhaust Plumes from Flying Rockets

    NASA Astrophysics Data System (ADS)

    Tyler, D.; Phenis, A.; Mudge, J.; Dank, J.; Tietjen, A.; Hand, D.

    We report video-rate imaging polarimetry observations of boosting rocket exhaust plumes obtained with the Lockheed Martin ATC Simultaneous Stokes Imaging Polarimeter (SSIP). The unique design of the SSIP, mounted to a tracking telescope, allowed us to acquire rocket plume imaging polarimetry over a period of minutes for each of three launches, observing the plume phenomenology at a variety of altitudes and aspect angles. Our data includes image polarimetry of solid and liquid rocket exhaust plumes and mixing of the two exhaust flow fields. We also present multi-spectral image data, showing for the first time distinct shock structures in the plumes at two wavelengths.

  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. Stennis Space Center's approach to liquid rocket engine health monitoring using exhaust plume diagnostics

    NASA Technical Reports Server (NTRS)

    Gardner, D. G.; Tejwani, G. D.; Bircher, F. E.; Loboda, J. A.; Van Dyke, D. B.; Chenevert, D. J.

    1991-01-01

    Details are presented of the approach used in a comprehensive program to utilize exhaust plume diagnostics for rocket engine health-and-condition monitoring and assessing SSME component wear and degradation. This approach incorporates both spectral and video monitoring of the exhaust plume. Video monitoring provides qualitative data for certain types of component wear while spectral monitoring allows both quantitative and qualitative information. Consideration is given to spectral identification of SSME materials and baseline plume emissions.

  7. 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.

  8. Rocket exhaust plume impingement on the Voyager spacecraft

    NASA Technical Reports Server (NTRS)

    Baerwald, R. K.

    1978-01-01

    In connection with the conduction of the long-duration Voyager missions to the outer planets and the sophisticated propulsion systems required, it was necessary to carry out an investigation to avoid exhaust plume impingement problems. The rarefied gas dynamics literature indicates that, for most engineering surfaces, the assumption of diffuse reemission and complete thermal accommodation is warranted in the free molecular flow regime. This assumption was applied to an analysis of a spacecraft plume impingement problem in the near-free molecular flow regime and yielded results to within a few percent of flight data. The importance of a correct treatment of the surface temperature was also demonstrated. Specular reflection, on the other hand, was shown to yield results which may be unconservative by a factor of 2 or 3. It is pointed out that one of the most difficult portions of an exhaust plume impingement analysis is the simulation of the impinged hardware. The geometry involved must be described as accurately and completely as possible.

  9. Pulse moire interferogram of rocket exhaust plume and its quantitative analysis

    NASA Astrophysics Data System (ADS)

    Miao, Peng C.; Yan, Da P.; Wang, Hai L.; He, An Z.

    1992-01-01

    This paper presented the measurement of the exhaust plume field of solid rocket with a moire interferometer. The principle of the moire interferometer is presented. The moire interferograms of the plume were processed with computer digital image process technique. It was found that the grey distribution of the moire interferograms has a sharp valley, and from the analysis this paper presented a method to thin and trace a deformated moire fringe on one step. This paper presented the result of the quantitative measurement.

  10. 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.

  11. 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

  12. 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

  13. 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.

  14. 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.

  15. In situ measurement of the aerosol size distribution in stratospheric solid rocket motor exhaust plumes

    NASA Astrophysics Data System (ADS)

    Ross, M. N.; Whitefield, P. D.; Hagen, D. E.; Hopkins, A. R.

    The concentration and size distribution of aerosol in the stratospheric exhaust plumes of two Space Shuttle rockets and one Titan IV rocket were measured using a two component aerosol sampling system carried aboard a WB-57F aircraft. Aerosol size distribution in the 0.01 µm to 4 µm diameter size range was measured using a two component sampling system. The measured distributions display a trimodal form with modes near 0.005 µm, 0.09 µm, and 2.03 µm and are used to infer the relative mass fractionation among the three modes. While the smallest mode has been estimated to contain as much as 10% of the total mass of SRM exhaust alumina, we find show that the smallest mode contains less than 0.05% of the alumina mass. This fraction is so small so as to significantly reduce the likelihood that heterogeneous reactions on the SRM alumina surfaces could produce a significant global impact on stratospheric chemistry.

  16. 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.

  17. 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.

  18. 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.

  19. 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

  20. 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

  1. 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.

  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 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.

  3. 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.

  4. Nuclear rocket plume studies

    NASA Astrophysics Data System (ADS)

    Hastings, Daniel

    1993-05-01

    A description and detailed computational analysis of a vortex cleaning system designed to remove radioactive material from the plumes of nuclear rockets is included. The proposed system is designed to remove both particulates and radioactive gaseous material from the plume. A two part computational model is used to examine the system's ability to remove particulates, and the results indicate that under some conditions, the system can remove over 99% of the particles in the flow. Two critical parameters which govern the effectiveness of the system are identified and the information necessary to estimate cleaning efficiencies for particles of known sizes and densities is provided. A simple steady analytical solution is also developed to examine the system's ability to remove gaseous radioactive material. This analysis, while inconclusive, suggests that the swirl rates necessary to achieve useful efficiencies are too high to be achieved in any practical manner. Therefore, this system is probably not suitable for use, with gaseous radioactive material. It was concluded that the system can cause negligible specific impulse losses, though there may be a substantial mass penalty associated with its use.

  5. 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.

  6. 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.

  7. Characterization of the exhaust particulates in the ground cloud and high-altitude plume of large solid-propellant booster rockets

    NASA Technical Reports Server (NTRS)

    Strand, L. D.; Bowyer, J. M.; Varsi, G.; Laue, E. G.; Gauldin, R.

    1980-01-01

    The report is concerned with the characterization of Al2O3 particles in the atmosphere. These particles comprise one of the major combustion products of the rocket propellant employed in the Space Shuttle boosters. A ground cloud and stratospheric plume are considered. It is concluded that the typical residence times in the atmosphere are much longer than earlier estimates have indicated.

  8. Detection of Metallic Compounds in Rocket Plumes

    NASA Astrophysics Data System (ADS)

    Rogers, Chris; Dunn, Dr. Robert

    1998-04-01

    Recent experiments using metal mixed in hydroxyl-terminated polybutadiene (HTPB) fuel grains in small hybrid rocket indicates ion detectors may be effective in detection of metallic compounds in rocket plumes. We wanted to ascertain the extent to which the presence of metallic compounds in rocket plumes could be detected using ion probes and Gaussian rings. Charges that collide with or pass near the intruding probe are detected. Gaussian rings, short insulated cylindrical Gaussian surfaces, enclose the plume without intruding into the plume. Charges in the plume are detected by currents they induce in the cylinder.

  9. Rocket plume temperature measurement by wire welded thermocouples

    NASA Astrophysics Data System (ADS)

    Xu, Qiang

    2006-05-01

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

  10. Rocket Engine Plume Diagnostics at Stennis Space Center

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. Propagation of light through ship exhaust plumes

    NASA Astrophysics Data System (ADS)

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

    2014-10-01

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

  15. Terminology and assessment methods of solid propellant rocket exhaust signatures

    NASA Astrophysics Data System (ADS)

    1993-02-01

    The Propulsion and Energetics Panel's Specialists' Meeting in autumn 1985 on Smokeless Propellants demonstrated that no common standard was available in this field and that the lack of common understanding led to misunderstanding amongst the NATO community. After some preparatory discussion, the Panel, therefore, formed Working Group Number 21 with the objectives of defining methods for the assessment of rocket motor exhaust optical properties in the visible and in the infrared range, and of recommending a terminology based on quantitative criteria. The Working Group discussed the subject in a total of eight sessions and prepared this Advisory Report. Following an Introduction and Summary there are six chapters, commencing with an Overview and continuing with Propellant Smoke Classification, Plume Primary Smoke, Plume Secondary Smoke, Plume Radiation and Plume Microwave Properties. In most cases, the conclusions and recommendations follow the chapters and are not repeated at the end of the report.

  16. DESIGN CRITERIA FOR ROCKET EXHAUST SCRUBBERS

    EPA Science Inventory

    The report gives results of an engineering study and design of methods for scrubbing the exhaust of static-tested solid rockets. Pollutants of major concern were hydrogen chloride and hydrogen fluoride gases. The best process for removing these gases was found to be a gas-atomize...

  17. 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.

  18. Ion recombination in aircraft exhaust plumes

    NASA Astrophysics Data System (ADS)

    Sorokin, A.; Mirabel, P.

    In this article, a model which examines the evolution of ion concentrations in a hot aircraft exhaust plume on the ground is proposed. The model includes plume dilution and ion-ion recombination with coefficients which vary with temperature. A comparison of the model is made with the available ground-based experimental data obtained on the ATTAS research aircraft engines. From this comparison, an ion emission index of the order of 8 1016 ions/kg(fuel) inferred.

  19. 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.

  20. Exhaust Plume Measurements of the VASIMR VX-200

    NASA Astrophysics Data System (ADS)

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

    2008-11-01

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

  1. An expert system for spectroscopic analysis of rocket engine plumes

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  2. 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.

  3. 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.

  4. Solar rocket plume/mirror interactions

    NASA Technical Reports Server (NTRS)

    Yu, Sheng-Tao; Chang, Chau-Lyan; Merkle, Charles 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. 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. PMID:24578655

  10. 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.

  11. Monopropellant thruster exhaust plume contamination measurements

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  12. 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.

  13. 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.

  14. Noninvasive Detection of Metallic Ions in a Hybrid Rocket Plume

    NASA Astrophysics Data System (ADS)

    Mack, John Hunter; Dunn, Robert

    2000-03-01

    This project utilizes a system of Gaussian cylinders to detect charged particles in a hybrid rocket plume. The approach is based on the phenomenon that moving charges will induce electric currents in nearby conductive surfaces. Consequently, charged particles in a rocket plume will induce electrical currents in a surrounding conductive cylindrical surface. Since these currents are proportional to the net charge being transported in a plume at any particular time, their measurement provides a method for determining the net charge ejected from the rocket motor. The Gaussian surfaces used in this project were conducting cylinders coated with an insulated material in order to differentiate contact charges from induced charges. Two configurations were used as noninvasive detectors. Single cylinders located around the plume produced results in particle detection. The second configuration consisted of two Gaussian surfaces, which allowed a differential amplifier to measure the potential difference between the cylinders while simultaneously canceling the common electromagnetic noise.

  15. In-Situ Microphysical Measurements In Rocket Plumes With The Cloud And Aerosol Spectrometer (CAS)

    NASA Astrophysics Data System (ADS)

    Kok, G.; Baumgardner, D.; Avallone, L.; Kalnajs, L.; Herman, R.; Ross, M.; Thompson, T.; Toohey, D.

    2005-12-01

    High resolution, single particle measurements have been made in rocket plumes using an optical particle spectrometer that measures diameters from 0.5 to 44 um. The Cloud and Aerosol Spectrometer (CAS) measures the light scattered in two directions from individual particles that pass through a focused, 680 nm laser beam and we derive the diameter, shape and composition from this information. The CAS was mounted on the NASA WB57-F aircraft as part of the Plume Ultrafast Measurements Acquisition (PUMA) project, an experiment funded by NSF and NASA to study the chemistry and microphysics of rocket plumes. Measurements were first made in a plume generated by an Atlas IIAS rocket in May, 2004 and again in July, 2005 in the plume formed from the exhaust of the solid state boosters used to launch the space shuttle Discovery into orbit. The microstructure of the two plumes and the characteristics of their particles were distinctly different. The two cases had similar maximum concentrations of 300 cm-3, but the space shuttle particles were on average larger and a greater percentage of them were irregular in shape. An analysis of the distance between particles suggests clustering because of the non-Poisson shape of the frequency distribution of inter-arrival times.

  16. 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.

  17. 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.

  18. Rocket plume radiation base heating by reverse Monte Carlo simulation

    NASA Astrophysics Data System (ADS)

    Everson, John; Nelson, H. F.

    1993-10-01

    A reverse Monte Carlo radiative transfer code is developed to predict rocket plume base heating. It is more computationally efficient than the forward Monte Carlo method, because only the radiation that strikes the receiving point is considered. The method easily handles both gas and particle emission and particle scattering. Band models are used for the molecular emission spectra, and the Henyey-Greenstein phase function is used for the scattering. Reverse Monte Carlo predictions are presented for (1) a gas-only model of the Space Shuttle main engine plume; (2) a purescattering plume with the radiation emitted by a hot disk at the nozzle exit; (3) a nonuniform temperature, scattering, emitting and absorbing plume; and (4) a typical solid rocket motor plume. The reverse Monte Carlo method is shown to give good agreement with previous predictions. Typical solid rocket plume results show that (1) CO2 radiation is emitted from near the edge of the plume; (2) H2O gas and Al2O3 particles emit radiation mainly from the center of the plume; and (3) Al2O3 particles emit considerably more radiation than the gases over the 400-17,000 cm(exp -1) spectral interval.

  19. 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

  20. Rocket plume flowfield characterization using laser Rayleigh scattering

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J.; Weiss, Jonathan M.

    1992-01-01

    A Doppler-resolved laser Rayleigh scattering diagnostic was applied to a 111 N thrust, regenerative and fuel-film cooled, gaseous hydrogen/gaseous oxygen rocket engine. The axial and radial mean gas velocities were measured from the net Doppler shifts observed for two different scattering angles. Translational temperatures and number densities were estimated from the Doppler widths and scattered intensities, respectively, by assuming that water was the dominant scattering species in the exhaust. The experimental results are compared with theoretical predictions from a full Navier-Stokes code (RD/RPLUS) and the JANNAF Two-Dimensional Kinetics (TDK) and Standardized Plume Flowfield (SPF-II) codes. Discrepancies between the measured and predicted axial velocities, temperatures, and number densities are evident. Radial velocity measurements, however, show excellent agreement with predictions. The discrepancies are attributed primarily to inefficient mixing and combustion caused by the injection of excessive oxidizer along one side of the thrust chamber. Thrust and mass flow rate estimates obtained from the Rayleigh measurements show excellent agreement with the globally measured values.

  1. 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.

  2. Instrumentation for In-Flight SSME Rocket Engine Plume Spectroscopy

    NASA Technical Reports Server (NTRS)

    Madzsar, George C.; Bickford, Randall L.; Duncan, David B.

    1994-01-01

    This paper describes instrumentation that is under development for an in-flight demonstration of a plume spectroscopy system on the space shuttle main engine. The instrumentation consists of a nozzle mounted optical probe for observation of the plume, and a spectrometer for identification and quantification of plume content. This instrumentation, which is intended for use as a diagnostic tool to detect wear and incipient failure in rocket engines, will be validated by a hardware demonstration on the Technology Test Bed engine at the Marshall Space Flight Center.

  3. Instrumentation for in-flight SSME rocket engine plume spectroscopy

    NASA Astrophysics Data System (ADS)

    Madzsar, George C.; Bickford, Randall L.; Duncan, David B.

    1994-06-01

    This paper describes instrumentation that is under development for an in-flight demonstration of a plume spectroscopy system on the space shuttle main engine. The instrumentation consists of a nozzle mounted optical probe for observation of the plume, and a spectrometer for identification and quantification of plume content. This instrumentation, which is intended for use as a diagnostic tool to detect wear and incipient failure in rocket engines, will be validated by a hardware demonstration on the Technology Test Bed engine at the Marshall Space Flight Center.

  4. ASRM plume deflector analysis program. [advanced solid rocket motor

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  5. 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.

  6. 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.

  7. Effect of water to ablative performance under solid rocket exhaust environment

    NASA Astrophysics Data System (ADS)

    Miller, M. J.; Koo, J. H.; Sickler, F. M.; Lecureux, F.; Dash, S. M.

    1993-06-01

    The local environment during a missile firing is particularly hostile. Thermal protection of the missile launcher structure is often achieved with ablatives. Ablatives erode when subjected to high-temperature rocket exhaust, but the backside temperature of the protected structure remains relatively cool due to the insulative nature of ablatives. Multiple missile firings can completely erode the ablative, exposing the launching system components to an extremely high temperature. This investigation addresses the concept of injecting water into the missile plume to reduce the amount of ablative erosion per missile firing. This concept also reduces the amount of ablative materials needed in missile launching systems. Injecting water into the exhaust plume in a controlled laboratory environment was performed. Heat flux and material erosion measurements were compared in this study.

  8. Measurements of Unexpected Ozone Loss in a Nighttime Space Shuttle Exhaust Plume: Implications for Geo-Engineering Projects

    NASA Astrophysics Data System (ADS)

    Avallone, L. M.; Kalnajs, L. E.; Toohey, D. W.; Ross, M. N.

    2008-12-01

    Measurements of ozone, carbon dioxide and particulate water were made in the nighttime exhaust plume of the Space Shuttle (STS-116) on 9 December 2006 as part of the PUMA/WAVE campaign (Plume Ultrafast Measurements Acquisition/WB-57F Ascent Video Experiment). The launch took place from Kennedy Space Center at 8:47 pm (local time) on a moonless night and the WB-57F aircraft penetrated the shuttle plume approximately 25 minutes after launch in the lowermost stratosphere. Ozone loss is not predicted to occur in a nighttime Space Shuttle plume since it has long been assumed that the main ozone loss mechanism associated with rocket emissions requires solar photolysis to drive several chlorine-based catalytic cycles. However, the nighttime in situ observations show an unexpected loss of ozone of approximately 250 ppb in the evolving exhaust plume, inconsistent with model predictions. We will present the observations of the shuttle exhaust plume composition and the results of photochemical models of the Space Shuttle plume. We will show that models constrained by known rocket emission kinetics, including afterburning, and reasonable plume dispersion rates, based on the CO2 observations, cannot explain the observed ozone loss. We will propose potential explanations for the lack of agreement between models and the observations, and will discuss the implications of these explanations for our understanding of the composition of rocket emissions. We will describe the potential consequences of the observed ozone loss for long-term damage to the stratospheric ozone layer should geo-engineering projects based on rocket launches be employed.

  9. Rocket Exhaust Cratering: A Significant Challenge for Space Exploration

    NASA Technical Reports Server (NTRS)

    Metzger, Philip

    2005-01-01

    During the Apollo and Viking programs, NASA needed to know how the rocket exhaust would affect the soil on the Moon and Mars. A number of studies were done during the 50's through 70's, but little or no work has been done since then. Existing models are crude and significant theoretical questions still exist

  10. Predicting engine parameters using the optic spectrum of the space shuttle main engine exhaust plume

    NASA Astrophysics Data System (ADS)

    Srivastava, Ashok N.; Buntine, Wray

    The Optical Plume Anomaly Detection (OPAD) system is under development to predict engine anomalies and engine parameters of the Space Shuttle's Main Engine (SSME). The anomaly detection is based on abnormal metal concentrations in the optical spectrum of the rocket plume. Such abnormalities could be indicative of engine corrosion or other malfunctions. Here, we focus on the second task of the OPAD system, namely the prediction of engine parameters such as rated power level (RPL) and mixture ratio (MR). Because of the high dimensionality of the spectrum, we developed a linear algorithm to resolve the optical spectrum of the exhaust plume into a number of separate components, each with a different physical interpretation. These components are used to predict the metal concentrations and engine parameters for online support of ground-level testing of the SSME. Currently, these predictions are labor intensive and cannot be done online. We predict RPL using neural networks and give preliminary results.

  11. 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.

  12. The measurement of electron density in a rocket motor plume

    NASA Astrophysics Data System (ADS)

    Cooper, David A.; Frederick, Robert A.

    1993-06-01

    This paper discusses the development of a diagnostic technique to measure the electron density in a rocket motor plume in order to characterize and rank solid rocket propellants based on their propensity to attenuate the communication signal to a missile. Three techniques were originally investigated as possible low-cost approaches that could be used for plume comparisons as a function of propellant. These approaches consisted of Langmuir probes, electromagnetic coils, and focused microwave probes. The focused microwave probe concept was considered the most appropriate technique to implement for the research to be conducted. The complete design and analysis of a focused microwave probe system operating at 17 GHz was conducted and the selection to determine this operating frequency discussed. Initial estimates of general uncertainty analysis suggest very good results are obtainable using a F-4 lens system and horn diameter of 17 in. for the 17 GHz frequency.

  13. 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

  14. Simulation of UV atomic radiation for application in exhaust plume spectrometry

    NASA Technical Reports Server (NTRS)

    Wallace, T. L.; Powers, W. T.; Cooper, A. E.

    1993-01-01

    Quantitative analysis of exhaust plume spectral data has long been a goal of developers of advanced engine health monitoring systems which incorporate optical measurements of rocket exhaust constituents. Discussed herein is the status of present efforts to model and predict atomic radiation spectra and infer free-atom densities from emission/absorption measurements as part of the Optical Plume Anomaly Detection (OPAD) program at Marshall Space Flight Center (MSFC). A brief examination of the mathematical formalism is provided in the context of predicting radiation from the Mach disk region of the SSME exhaust flow at nominal conditions during ground level testing at MSFC. Computational results are provided for Chromium and Copper at selected transitions which indicate a strong dependence upon broadening parameter values determining the absorption-emission line shape. Representative plots of recent spectral data from the Stennis Space Center (SSC) Diagnostic Test Facility (DTF) rocket engine are presented and compared to numerical results from the present self-absorbing model; a comprehensive quantitative analysis will be reported at a later date.

  15. Simulation of UV atomic radiation for application in exhaust plume spectrometry

    NASA Astrophysics Data System (ADS)

    Wallace, T. L.; Powers, W. T.; Cooper, A. E.

    1993-06-01

    Quantitative analysis of exhaust plume spectral data has long been a goal of developers of advanced engine health monitoring systems which incorporate optical measurements of rocket exhaust constituents. Discussed herein is the status of present efforts to model and predict atomic radiation spectra and infer free-atom densities from emission/absorption measurements as part of the Optical Plume Anomaly Detection (OPAD) program at Marshall Space Flight Center (MSFC). A brief examination of the mathematical formalism is provided in the context of predicting radiation from the Mach disk region of the SSME exhaust flow at nominal conditions during ground level testing at MSFC. Computational results are provided for Chromium and Copper at selected transitions which indicate a strong dependence upon broadening parameter values determining the absorption-emission line shape. Representative plots of recent spectral data from the Stennis Space Center (SSC) Diagnostic Test Facility (DTF) rocket engine are presented and compared to numerical results from the present self-absorbing model; a comprehensive quantitative analysis will be reported at a later date.

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

    NASA Astrophysics Data System (ADS)

    Gelbwachs, Jerry A.

    1996-05-01

    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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Radiation/convection coupling in rocket motors and plumes

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  2. ASRM subscale plume deflector testing. [advanced solid rocket motor

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  3. 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.

  4. Three dimensional particle simulation of high altitude rocket plumes

    NASA Astrophysics Data System (ADS)

    Dagum, Leonardo; Zhu, S. H. K.

    1992-07-01

    The interaction of two nozzles exhausting into vacuum generates a complex three-dimensional shock structure. The shock structure and resulting plume flowfield is characterized by the nozzle separation distance. For the appropriate range of penetration Knudsen numbers, the analysis of this shock structure can be suitably accomplished through a Monte Carlo simulation. This paper describes the application of a general three-dimensional Monte Carlo simulation on the Connection Machine CM-2 to the analysis of the plume self-interaction shock in the near field. Results are presented for two cases, corresponding to a small and a large nozzle separation distance. The results correctly reproduce the expected flow features and demonstrate the ability of this method to properly simulate the start of the plume self-interaction shock. This has significance not only for allowing analysis of the self-interacting plume in the near field, but also for allowing the subsequent simulation of the far field flow through the use of a continuation downstream exit boundary.

  5. 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

  6. 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.

  7. Thermal radiation of heterogeneous combustion products in the model rocket engine plume

    NASA Astrophysics Data System (ADS)

    Kuzmin, V. A.; Maratkanova, E. I.; Zagray, I. A.; Rukavishnikova, R. V.

    2015-05-01

    The work presents a method of complex investigation of thermal radiation emitted by heterogeneous combustion products in the model rocket engine plume. Realization of the method has allowed us to obtain full information on the results in all stages of calculations. Dependence of the optical properties (complex refractive index), the radiation characteristics (coefficients and cross sections) and emission characteristics (flux densities, emissivity factors) of the main determining factors and parameters was analyzed. It was found by the method of computational experiment that the presence of the gaseous phase in the combustion products causes a strongly marked selectivity of emission, due to which the use of gray approximation in the calculation of thermal radiation is unnecessary. The influence of the optical properties, mass fraction, the function of particle size distribution, and the temperature of combustion products on thermal radiation in the model rocket engine plume was investigated. The role of "spotlight" effect-increasing the amount of energy of emission exhaust combustion products due to scattering by condensate particles radiation from the combustion chamber-was established quantitatively.

  8. 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.

  9. Hyper-spectral imaging of aircraft exhaust plumes

    NASA Astrophysics Data System (ADS)

    Bowen, Spencer; Bradley, Kenneth; Gross, Kevin; Perram, Glen; Marciniak, Michael

    2008-10-01

    An imaging Fourier-transform spectrometer has been used to determine low spatial resolution temperature and chemical species concentration distributions of aircraft jet engine exhaust plumes. An overview of the imaging Fourier transform spectrometer and the methodology of the project is presented. Results to date are shared and future work is discussed. Exhaust plume data from a Turbine Technologies, LTD, SR-30 turbojet engine at three engine settings was collected using a Telops Field-portable Imaging Radiometric Spectrometer Technology Mid-Wave Extended (FIRST-MWE). Although the plume exhibited high temporal frequency fluctuations, temporal averaging of hyper-spectral data-cubes produced steady-state distributions, which, when co-added and Fourier transformed, produced workable spectra. These spectra were then reduced using a simplified gaseous effluent model to fit forward-modeled spectra obtained from the Line-By-Line Radiative Transfer Model (LBLRTM) and the high-resolution transmission (HITRAN) molecular absorption database to determine approximate temperature and concentration distributions. It is theorized that further development of the physical model will produce better agreement between measured and modeled data.

  10. 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.

  11. Radiation/convection coupling in rocket motor and plume analysis

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    A method for describing radiation/convection coupling to a flow field analysis was developed for rocket motors and plumes. The three commonly used propellant systems (H2/O2, RP-1/O2, and solid propellants) radiate primarily as: molecular emitters, non-scattering small particles (soot), and scattering larger particles (Al2O3), respectively. For the required solution, the divergence of the radiation heat flux was included in the energy equation, and the local, volume averaged intensity was determined by a solution to the radiative transfer equation. A rigorous solution to this problem is intractable, therefore, solution methods which use the ordinary and improved differential approximation were developed. This radiation model was being incorporated into the FDNS code, a Navier-Stokes flowfield solver for multiphase, turbulent combusting flows.

  12. Dual-beam multiple wavelength light transmittance measurement for particle sizing in rocket motor plumes

    NASA Astrophysics Data System (ADS)

    Taylor, Kevin B.

    1993-06-01

    A multiple-wavelength light transmittance measurement system previously used in a laboratory environment to study particles in solid rocket propellant exhaust plumes was modified for use in the field, where high levels of vibration can degrade the accuracy of data. The system was converted from a single light beam configuration to a dual beam configuration which was capable of obtaining a complete set of 1024 reference and scene measurements in 10.0 ms. Modifications included designing, building and testing a new analog-to-digital data converter trigger circuit, and a rotating-wheel light chopper. Optical components including beam splitters, lenses, and a fiber optic cable were installed, and existing data collection system software was modified. The new system was tested by measuring soot from an oxyacetylene torch to prove the design concept. Test results and system performance were documented. Recommendations for further modifications, improvements and applications are presented.

  13. 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.

  14. 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.

  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. Exhaust gas treatment in testing nuclear rocket engines

    NASA Astrophysics Data System (ADS)

    Zweig, Herbert R.; Fischler, Stanley; Wagner, William R.

    1993-01-01

    With the exception of the last test series of the Rover program, Nuclear Furnace 1, test-reactor and rocket engine hydrogen gas exhaust generated during the Rover/NERVA program was released directly to the atmosphere, without removal of the associated fission products and other radioactive debris. Current rules for nuclear facilities (DOE Order 5480.6) are far more protective of the general environment; even with the remoteness of the Nevada Test Site, introduction of potentially hazardous quantities of radioactive waste into the atmosphere must be scrupulously avoided. The Rocketdyne treatment concept features a diffuser to provide altitude simulation and pressure recovery, a series of heat exchangers to gradually cool the exhaust gas stream to 100 K, and an activated charcoal bed for adsorption of inert gases. A hydrogen-gas fed ejector provides auxiliary pumping for startup and shutdown of the engine. Supplemental filtration to remove particulates and condensed phases may be added at appropriate locations in the system. The clean hydrogen may be exhausted to the atmosphere and flared, or the gas may be condensed and stored for reuse in testing. The latter approach totally isolates the working gas from the environment.

  17. Exhaust gas treatment in testing nuclear rocket engines

    SciTech Connect

    Zweig, H.R.; Fischler, S.; Wagner, W.R. )

    1993-01-15

    With the exception of the last test series of the Rover program, Nuclear Furnace 1, test-reactor and rocket engine hydrogen gas exhaust generated during the Rover/NERVA program was released directly to the atmosphere, without removal of the associated fission products and other radioactive debris. Current rules for nuclear facilities (DOE Order 5480.6) are far more protective of the general environment; even with the remoteness of the Nevada Test Site, introduction of potentially hazardous quantities of radioactive waste into the atmosphere must be scrupulously avoided. The Rocketdyne treatment concept features a diffuser to provide altitude simulation and pressure recovery, a series of heat exchangers to gradually cool the exhaust gas stream to 100 K, and an activated charcoal bed for adsorption of inert gases. A hydrogen-gas fed ejector provides auxiliary pumping for startup and shutdown of the engine. Supplemental filtration to remove particulates and condensed phases may be added at appropriate locations in the system. The clean hydrogen may be exhausted to the atmosphere and flared, or the gas may be condensed and stored for reuse in testing. The latter approach totally isolates the working gas from the environment.

  18. 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.

  19. 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.

  20. 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

  1. 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.

  2. 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.

  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. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Computation of underexpanded solid rocket plume and its effects on the mated Shuttle configuration

    NASA Astrophysics Data System (ADS)

    Chen, C. L.; Ramakrishnan, S.; Ota, D. K.; Rajagopal, K.; Wisneski, J.

    A numerical study of underexpanded solid rocket booster (SRB) plume and its effects on the mated Space Shuttle configuration are presented. The overall structure of some underexpanded plumes has been computed. The characteristics of the first shock cell such as barrel shock, Mach disk, plume slipstream, Mach disk slip stream and plume induced shock have been well captured. The plume structure depends on the freestream velocity, real gas effects and viscous diffusion. Solutions for perfect gas plume and for two species flow with different species in the plume and the external flow with different physical properties, especially different values for the ratio of specific heats (alpha), are discussed. Comparisons have been made between SRB plume-off and plume-on cases. The results indicate that the plume upstream effects occur for this low supersonic free-stream flow case (freestream Mach number = 1.25, alpha = 5.1 deg) because of a large local subsonic pocket upstream of the plume region. The two-species model results in larger wing load increments than the single-species model when compared with the plumeoff case.

  12. 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.

  13. Effect of bipropellant plume exhaust effluents on spaceborne optical instruments

    NASA Astrophysics Data System (ADS)

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

    1980-01-01

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

  14. 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.

  15. User's manual 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 REEDM computer program predicts concentrations, dosages, and depositions downwind from normal and abnormal launches of rocket vehicles at NASA's Kennedy Space Center. The atmospheric dispersion models, cloud-rise models, and other formulas used in the REEDM model are described mathematically Vehicle and source parameters, other pertinent physical properties of the rocket exhaust cloud, and meteorological layering techniques are presented as well as user's instructions for REEDM. Worked example problems are included.

  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. 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.

  18. Visualization of impingement field of real-rocket-exhausted jets by using moire deflectometry

    NASA Astrophysics Data System (ADS)

    He, An-Zhi; Yan, Da-Peng; Miao, Peng C.; Wang, Hai-Ling

    1991-12-01

    The experimental methods and results of an impingement field of a real rocket exhausted jet are presented. By using large aperture and long path moire deflector, the moire deflectograms of rocket free jet and rocket jet impingement field are obtained. From these moire deflectograms, the location of the Mach disk, the oblique shock wave, and the value of boundary are calculated quantitatively and compared with the results of numerical simulating and wind tunnel simulating. At the same time, we have found five new shock wave structures from the experiments.

  19. Analyzing rocket plume spectral data with neural networks

    NASA Astrophysics Data System (ADS)

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

    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.

  20. 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.

  1. 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

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

    NASA Technical Reports Server (NTRS)

    Kenny, Robert Jeremy

    2009-01-01

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

  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. 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).

  5. Incoherent scatter from space shuttle and rocket engine plumes in the ionosphere

    NASA Astrophysics Data System (ADS)

    Bernhardt, P. A.; Huba, J. D.; Swartz, W. E.; Kelley, M. C.

    1998-02-01

    Enhanced echoes from the 430 MHz radar at Arecibo were observed during burns of the space shuttle orbital maneuver subsystem (OMS) engines near 317 km altitude. Similar radar signatures of enhanced backscatter were also obtained by the Millstone Hill radar observing the plume of a Centaur engine burning in the ionosphere. A theoretical model of incoherent scatter is presented to explain the radar backscatter observations. The theory considers molecular ion beams generated in the exhaust plume as a result of charge exchange between the ambient O+ ions and the high-speed exhaust molecules (primarily H2O). The field-aligned gyromotion of the pickup ions affects the radio wave scattering from the random thermal fluctuations of electron density. Numerical calculations are carried out for plasmas modified by the space shuttle or Centaur engines, and reasonable agreement with observations is found for the total scattered power. Incoherent backscatter spectra respond to characteristics of the exhaust plume such as vector flow velocity, temperature, and composition. The nonequilibrium velocity distributions for the ions in the pickup ion plume are similar to the distributions found in strongly convecting auroral region ionospheres. The incoherent scatter from the plume ions can be used to validate techniques used to study naturally disturbed plasmas. The predictions of our radar scatter calculations will be tested in future experiments using the space shuttle OMS engines over incoherent scatter radars located at equatorial latitudes and midlatitudes.

  6. 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

  7. 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.

  8. 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.

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

    NASA Astrophysics Data System (ADS)

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

    1993-06-01

    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.

  10. 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.

  11. Dynamics of aircraft exhaust plumes in the jet-regime

    NASA Astrophysics Data System (ADS)

    Kärcher, B.; Fabian, P.

    1994-10-01

    A computational model describing the two-dimensional, turbulent mixing of a single jet of exhaust gas from aircraft engines with the ambient atmosphere is presented. The underlying assumptions and governing equations are examined and supplemented by a discussion of analytical solutions. As an application, the jet dynamics of a B747-400 aircraft engine in cruise and its dependence on key parameters is investigated in detail. The computer code for this dynamical model is computationally fast and can easily be coupled to complex chemical and microphysical models in order to perform comprehensive studies of atmospheric effects from aircraft exhaust emissions in the jet regime.

  12. 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.

  13. Chance encounter with a stratospheric kerosene rocket plume from Russia over California

    NASA Astrophysics Data System (ADS)

    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.

    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 rocket vehicle, most likely 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 one 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-fueled rocket.

  14. 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

  15. Inexpensive photodiode arrays for use in rocket plume and hot source monitoring and diagnostics

    NASA Astrophysics Data System (ADS)

    Snider, Dallas; Shanks, Robert; Cole, Reagan; Hudson, M. Keith

    2003-09-01

    The spectroscopic analysis of plume emissions is a non-intrusive method which has been used to check for fatigue and possible damage throughout the pumps and other mechanisms in a rocket motor or engine. These components are made of various alloys. Knowing the composition of the alloys and for which parts they are used, one can potentially determine from the emissions in the plume which component is failing. Currently, optical multichannel analyser systems are being used which utilize charge coupled devices, cost tens of thousands of dollars, are somewhat delicate, and usually require cooling. We have developed two rugged instruments using less expensive linear photodiode arrays as detectors. A high-resolution system was used to detect atomic emission lines while a low-resolution system was used to detect molecular emission bands. We have also written data acquisition software and built electronic circuits to control the arrays and collect data. While the National Aeronautics and Space Administration has used similar systems for characterization of the space shuttle main engine, the emissions from other rocket systems have not been surveyed so well. The two instruments described will be utilized to study hybrid rocket emissions at the University of Arkansas-Little Rock hybrid rocket facility.

  16. 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.

  17. 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.

  18. Cooled Ceramic Matrix Composite Panel Successfully Tested in Rocket Exhaust

    NASA Technical Reports Server (NTRS)

    Eckel, Andrew J.

    2001-01-01

    Actively cooled ceramic matrix composite (CMC) components are enabling or enhancing for a broad range of hypersonic and reusable launch vehicle propulsion systems. Teaming with other NASA centers, the Air Force, and industry, the Glenn Ceramics Branch has successfully tested multiple cooled CMC panel concepts in high-heat-flux, high-pressure, flowing rocket engine combustion gas environments. Sub-element components survived multiple cycles and the severe thermal gradients imposed by combustion gas temperatures in excess of 5500 F and cryogenic hydrogen or ambient temperature water internal coolants. These demonstrations are critical for the continued development of this class of materials, and the research is expected to continue with additional concepts and increasingly larger and more complex geometries being fabricated and tested in a broad range of engine operating conditions.

  19. 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.

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

    NASA Technical Reports Server (NTRS)

    Anderson, Bruce

    2004-01-01

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

  1. 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. PMID:20134660

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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

  15. Electrets used to measure exhaust cloud effluents from Solid Rocket Motor (SRM) during demonstration model (DM-2) static test firing

    NASA Technical Reports Server (NTRS)

    Susko, M.

    1978-01-01

    Electrets were compared with fixed flow samplers during static test firing. The measurement of the rocket exhaust effluents by samplers and electrets indicated that the Solid Rocket Motor had no significant effect on the air quality in the area sampled. The results show that the electrets (a passive device which needs no power) can be used effectively alongside existing measuring devices (which need power). By placing electrets in areas where no power is available, measurements may be obtained. Consequently, it is a valuable complementary instrument in measuring rocket exhaust effluents in areas where other measuring devices may not be able to assess the contaminants.

  16. 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

  17. An overview of in-flight plume diagnostics for rocket engines

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  18. 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.

  19. 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.

  20. Space shuttle exhaust plumes in the lower thermosphere: Advective transport and diffusive spreading

    NASA Astrophysics Data System (ADS)

    Stevens, Michael H.; Lossow, Stefan; Siskind, David E.; Meier, R. R.; Randall, Cora E.; Russell, James M.; Urban, Jo; Murtagh, Donal

    2014-02-01

    The space shuttle main engine plume deposited between 100 and 115 km altitude is a valuable tracer for global-scale dynamical processes. Several studies have shown that this plume can reach the Arctic or Antarctic to form bursts of polar mesospheric clouds (PMCs) within a few days. The rapid transport of the shuttle plume is currently not reproduced by general circulation models and is not well understood. To help delineate the issues, we present the complete satellite datasets of shuttle plume observations by the Sounding of the Atmosphere using Broadband Emission Radiometry instrument and the Sub-Millimeter Radiometer instrument. From 2002 to 2011 these two instruments observed 27 shuttle plumes in over 600 limb scans of water vapor emission, from which we derive both advective meridional transport and diffusive spreading. Each plume is deposited at virtually the same place off the United States east coast so our results are relevant to northern mid-latitudes. We find that the advective transport for the first 6-18 h following deposition depends on the local time (LT) of launch: shuttle plumes deposited later in the day (~13-22 LT) typically move south whereas they otherwise typically move north. For these younger plumes rapid transport is most favorable for launches at 6 and 18 LT, when the displacement is 10° in latitude corresponding to an average wind speed of 30 m/s. For plumes between 18 and 30 h old some show average sustained meridional speeds of 30 m/s. For plumes between 30 and 54 h old the observations suggest a seasonal dependence to the meridional transport, peaking near the beginning of year at 24 m/s. The diffusive spreading of the plume superimposed on the transport is on average 23 m/s in 24 h. The plume observations show large variations in both meridional transport and diffusive spreading so that accurate modeling requires knowledge of the winds specific to each case. The combination of transport and spreading from the STS-118 plume in August

  1. 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.

  2. A ROCKET EXHAUST DEPLETION (RED) OF IONOSPHERIC TOTAL ELECTRON CONTENT DETECTED BY MULTIPLE DIAGNOSTICS

    NASA Astrophysics Data System (ADS)

    Baumgardner, J. L.; Smith, S. M.; Mendillo, M.; Martinis, C. R.; Coster, A. J.; Erickson, P. J.

    2009-12-01

    On April 30, 2005 one of the last Titan IIIC rockets was launched from Cape Canaveral, Fl. The trajectory was north eastward close to the eastern seaboard of the U.S. The second stage engine deposited over 30,000kg of exhaust gases over an altitude range from 273 km to 494 km, spanning a down range distance of ~1400km . The resulting Rocket Exhaust Depletion (RED) of ionospheric plasma was measured by multiple ground-based and space-based diagnostic techniques. The optical signature was imaged by the Boston Univ. all-sky imaging system at Millstone Hill Observatory, Westford, MA as a 3.5kR increase in 6300A airglow. The Total Electron Content (TEC), as measured by GPS receivers, was observed to decrease to 75% of pre-launch values within 30 minutes of the passage of the rocket. The Ionosonde at Wallops Is., ~600km from the trajectory, recorded a 70% decrease of NmF2 ~40 minus after the launch. A passage of DMSP (alt=840 km), ~ 1 hour after the launch, and ~700km to the west recorded a 20% drop in Total Ion Density. The ionospheric effects of RED’s are similar in magnitude to those from Storm Enhanced Densities (SED’s) and can have similar consequences for GPS navigation systems.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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

  12. Plume radiation

    NASA Astrophysics Data System (ADS)

    Dirscherl, R.

    1993-06-01

    The electromagnetic radiation originating from the exhaust plume of tactical missile motors is of outstanding importance for military system designers. Both missile- and countermeasure engineer rely on the knowledge of plume radiation properties, be it for guidance/interference control or for passive detection of adversary missiles. To allow access to plume radiation properties, they are characterized with respect to the radiation producing mechanisms like afterburning, its chemical constituents, and reactions as well as particle radiation. A classification of plume spectral emissivity regions is given due to the constraints imposed by available sensor technology and atmospheric propagation windows. Additionally assessment methods are presented that allow a common and general grouping of rocket motor properties into various categories. These methods describe state of the art experimental evaluation techniques as well as calculation codes that are most commonly used by developers of NATO countries. Dominant aspects influencing plume radiation are discussed and a standardized test technique is proposed for the assessment of plume radiation properties that include prediction procedures. These recommendations on terminology and assessment methods should be common to all employers of plume radiation. Special emphasis is put on the omnipresent need for self-protection by the passive detection of plume radiation in the ultraviolet (UV) and infrared (IR) spectral band.

  13. 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.

  14. 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.

  15. Exhaust cloud rise and diffusion in the atmosphere

    NASA Technical Reports Server (NTRS)

    Chandler, M. W.; Chu, R. T.; Thayer, S. D.

    1971-01-01

    Analytical approach develops physical-mathematical model of rocket engine exhaust cloud rise, growth, and diffusion. Analytic derivations and resultant model apply to hot exhaust cloud study or industrial stack plumes, making work results applicable to air pollution. Model formulations apply to all exhaust cloud types and various atmospheric conditions.

  16. 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.

  17. 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.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  20. Approach to SSME health monitoring. II - Exhaust plume emission spectroscopy at the DTF

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.; Loboda, John A.; Wheatley, Joseph S.; Chenevert, Donald J.

    1990-01-01

    The Diagnostics Testbed Facility (DTF) located at the Stennis Space Center (SSC) is used for obtaining extensive sets of H2O2 exhaust plume emission spectral data for the SSME critical components related elements and materials. The SSME related elements and materials are simulated by mixing appropriate amounts of compounds of their respective constituent elements in an aqueous solution which is injected into the combustion chamber of the DTFT. Five of the most critical components of the SSME which have experienced very severe wear and tear problems in the past are analyzed. These are high pressure turbopump (HPTP) turbine blades, HPTP turbine disks, HPTP bearing, main injector LOX posts, and the main combustion chamber structural shell. The alloys used in the manufacturing of these components are MAR-M 246 + Hf, Waspaloy X, AISI 440C, Haynes 188, and Inconel 718, respectively. The experimental setup and procedures at the DTF are described; stratospheric data for the five alloys are presented; and strategies for the material identification in the SSME exhaust plume are discussed.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. A feasibility study and mission analysis for the Hybrid Plume Plasma Rocket

    NASA Technical Reports Server (NTRS)

    Sullivan, Daniel J.; Micci, Michael M.

    1990-01-01

    The Hybrid Plume Plasma Rocket (HPPR) is a high power electric propulsion concept which is being developed at the MIT Plasma Fusion Center. This paper presents a theoretical overview of the concept as well as the results and conclusions of an independent study which has been conducted to identify and categorize those technologies which require significant development before the HPPR can be considered a viable electric propulsion device. It has been determined that the technologies which require the most development are high power radio-frequency and microwave generation for space applications and the associated power processing units, low mass superconducting magnets, a reliable, long duration, multi-megawatt space nuclear power source, and long term storage of liquid hydrogen propellant. In addition to this, a mission analysis of a one-way transfer from low earth orbit (LEO) to Mars indicates that a constant acceleration thrust profile, which can be obtained using the HPPR, results in faster trip times and greater payload capacities than those afforded by more conventional constant thrust profiles.

  11. 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.

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

    NASA Technical Reports Server (NTRS)

    Park, C.

    1976-01-01

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

  13. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

  15. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  17. 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.

  18. Crew radiation dose from the plume of a high impulse gas-core nuclear rocket during a Mars mission.

    NASA Technical Reports Server (NTRS)

    Masser, C. C.

    1971-01-01

    Analytical calculations are performed to determine the radiation dose rate and total dose to the crew of a gas-core nuclear rocket from the fission fragments located throughout the plume volume. The radiation dose from the plume fission fragments to two crew locations of 100 and 200 meters from the nozzle exit are calculated. It is found that, in the case of the most probable fission fragment retention time of 100 seconds, the crew must be protected from the radiation dose. Five centimeters of lead shielding would reduce the radiation dose by two orders of magnitude thereby protecting the crew. The increase in vehicle weight would be insignificant (7150 kg to a vehicle gross weight of 0.94 million kg).

  19. 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.

  20. 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.

  1. Correlative Observations with Space-Borne Direct Doppler Wind Instruments of the Rapid Transport of Shuttle Exhaust Plumes (Invited)

    NASA Astrophysics Data System (ADS)

    Niciejewski, R.; Meier, R. R.; Stevens, M. H.; Skinner, W. R.; Cooper, M.; Marshall, A.; Ortland, D. A.; Wu, Q.

    2010-12-01

    The Upper Atmosphere Research Satellite (UARS) was launched by Space Shuttle STS-48 on 12 September 1991 and included a direct Doppler experiment, the High Resolution Doppler Imager, HRDI. Ten years later, the TIMED Doppler Interferometer, TIDI, joined HRDI in direct neutral wind observations of the mesosphere and lower thermosphere (MLT). The removal of instrumental artifacts from the raw spectra, complicated by the loss of good attitude knowledge for HRDI and unexpected signal contamination for TIDI has matured to a level where excellent agreement exists for common volume measurements between them. The two experiments were able to perform overlapping measurements of tidal and planetary wave fields for three years permitting unprecedented clarity in the description of the cyclical behaviour of the MLT. The exhaust plume left in the wake of the launch of STS-107 (16 January 2003) provided a stringent test between TIDI, HRDI, and independent imagery, the latter of which showed rapid transport across the equator to the Antarctic. Though TIDI and HRDI observed the atmosphere at the plume’s location at different local solar times, all correlative observations supported the hypothesis indicated by once-a-day images of the plume - rapid southern transport over thousands of kilometers. A simple spectral analysis of simultaneous observations of the neutral winds by HRDI and TIDI indicates that a classical two-day wave (longitudinal wavenumber = 3) exists in the southern hemisphere during the ~80-hour transit time coinciding with the transport of the plume exhaust from launch to the Antarctic. A least-squares fit of the wave in the meridional wind indicates maximum amplitude in the MLT of ~80 m/s southwards. Other shuttle launches have also been accompanied by evidence that implies rapid transport of exhaust plumes to Arctic latitudes. This paper will summarize correlative HRDI and/or TIDI wind observations of these events and associated spectral analysis of the

  2. Characterizing and overcoming spectral artifacts in imaging Fourier-transform spectroscopy of turbulent exhaust plumes

    NASA Astrophysics Data System (ADS)

    Moore, Elizabeth A.; Gross, Kevin C.; Bowen, Spencer J.; Perram, Glen P.; Chamberland, Martin; Farley, Vincent; Gagnon, Jean-Philippe; Lagueux, Philippe; Villemaire, André

    2009-05-01

    The midwave and shortwave infrared regions of the electromagnetic spectrum contain rich information enabling the characterization of hot, rapid events such as explosions, engine plumes, flares and other combustion events. High-speed sensors are required to analyze the content of such rapidly evolving targets. Cameras with high frame rates and non-imaging spectrometers with high data rates are typically used; however the information from these two types of instruments must be later fused to enable characterization of the transient targets. Imaging spectrometers have recently become commercially available for general scientific use, thus enabling simultaneous capture of both spatial and spectral information without co-registration issues. However, their use against rapidly-varying sources has traditionally been considered problematic, for even at moderate spatial and spectral resolutions the time to acquire a single spectrum can be long compared to the timescales associated with combustion events. This paper demonstrates that imaging Fourier-transform spectroscopy (IFTS) can successfully characterize the turbulent combustion exhaust from a turbojet engine. A Telops Hyper-Cam IFTS collected hyperspectral video from a Turbine Technologies SR-30 turbojet engine with a spectral resolution of δν = 1/cm-1 on a 200×64 pixel sub-window at a rate of 0.3 Hz. Scene-change artifacts (SCAs) are present in the spectra; however, the stochastic fluctuations in source intensity translate into high-frequency "noise." Temporal averaging affords a significant reduction of the noise associated with SCAs. Emission from CO and CO2 are clearly recognized in the averaged spectra, and information about their temperature and relative concentrations is evident.

  3. 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.

  4. 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.

  5. OPAD data analysis. [Optical Plumes Anomaly Detection

    NASA Technical Reports Server (NTRS)

    Buntine, Wray L.; Kraft, Richard; Whitaker, Kevin; Cooper, Anita E.; Powers, W. T.; Wallace, Tim L.

    1993-01-01

    Data obtained in the framework of an Optical Plume Anomaly Detection (OPAD) program intended to create a rocket engine health monitor based on spectrometric detections of anomalous atomic and molecular species in the exhaust plume are analyzed. The major results include techniques for handling data noise, methods for registration of spectra to wavelength, and a simple automatic process for estimating the metallic component of a spectrum.

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

    NASA Technical Reports Server (NTRS)

    Powers, W. T.

    1989-01-01

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

  7. In situ measurements of carbon dioxide, 0.37-4.0 μm particles, and water vapor in the stratospheric plumes of small rockets

    NASA Astrophysics Data System (ADS)

    Gates, Amelia M.; Avallone, Linnea M.; Toohey, Darin W.; Rutter, Andrew P.; Whitefield, Philip D.; Hagen, Donald E.; Hopkins, A. Raymond; Ross, Martin N.; Zittel, Paul F.; Thompson, Thomas L.; Herman, Robert L.; Friedl, Randall R.

    2002-11-01

    Carbon dioxide (CO2) and large particles (0.37-4.0 μm) were measured in the stratospheric plume wakes of three rockets, an Atlas IIAS, a Delta II, and an Athena II. The correlations between CO2 mass and particle number densities in each plume are consistent with the unique combination of solid and liquid engine emissions of each rocket. Measured size distributions indicate a 1.1 μm mode with density of 2 g cm-3, consistent with spherical alumina particles emitted by solid rocket motors. Disagreement between the measured size distributions and the mean sizes inferred from the known alumina and CO2 emission indices and an observed increase in the particle number emission index with altitude are evidence for large particle oversampling effects and the presence of condensed volatile compounds within the particle population. Direct evidence for the latter is a persistent ~0.5-1 part per million (ppm) shortfall of water vapor relative to CO2 measured in the plume of the Athena II rocket based on the expected H2O/CO2 emission ratio. Although pure ice particles would not persist at the conditions of the measurements, a more stable coating of HNO3 (as either nitric acid trihydrate or as a liquid layer) could have reduced the sublimation rate of the underlying ice, thereby increasing the lifetimes of volatile particles within the plume. If confirmed, such a process would have important implications for the radiative and chemical properties of rocket plumes, including global ozone depletion associated with rocket launch activities.

  8. In situ measurements of carbon dioxide, 0.37-4.0 μm particles, and water vapor in the stratospheric plumes of small rockets

    NASA Astrophysics Data System (ADS)

    Gates, Amelia M.; Avallone, Linnea M.; Toohey, Darin W.; Rutter, Andrew P.; Whitefield, Philip D.; Hagen, Donald E.; Hopkins, A. Raymond; Ross, Martin N.; Zittel, Paul F.; Thompson, Thomas L.; Herman, Robert L.; Friedl, Randall R.

    2002-11-01

    Carbon dioxide (CO2) and large particles (0.37-4.0 μm) were measured in the stratospheric plume wakes of three rockets, an Atlas IIAS, a Delta II, and an Athena II. The correlations between CO2 mass and particle number densities in each plume are consistent with the unique combination of solid and liquid engine emissions of each rocket. Measured size distributions indicate a 1.1 μm mode with density of 2 g cm-3, consistent with spherical alumina particles emitted by solid rocket motors. Disagreement between the measured size distributions and the mean sizes inferred from the known alumina and CO2 emission indices and an observed increase in the particle number emission index with altitude are evidence for large particle oversampling effects and the presence of condensed volatile compounds within the particle population. Direct evidence for the latter is a persistent ˜0.5-1 part per million (ppm) shortfall of water vapor relative to CO2 measured in the plume of the Athena II rocket based on the expected H2O/CO2 emission ratio. Although pure ice particles would not persist at the conditions of the measurements, a more stable coating of HNO3 (as either nitric acid trihydrate or as a liquid layer) could have reduced the sublimation rate of the underlying ice, thereby increasing the lifetimes of volatile particles within the plume. If confirmed, such a process would have important implications for the radiative and chemical properties of rocket plumes, including global ozone depletion associated with rocket launch activities.

  9. 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.

  10. The investigation of man-made modifications of the ionosphere. [effects of detonations and rocket exhaust

    NASA Technical Reports Server (NTRS)

    Bernhardt, P. A.; Darosa, A. V.; Price, K. M.

    1980-01-01

    Topics covered include: (1) the application of ionosphere modifications models to the simulation of results obtained when rocket-borne explosives were detonated in the ionosphere; (2) the problem of hypersonic vapor releases from orbiting vehicles; (3) measuring the electron content reduction resulting from the firing of a Centaur rocket in the ionosphere; and (4) the preliminary design of the critical frequency tracker which displays the value of electron concentration at the peak of the F 2 region, in real time.

  11. 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.

  12. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    The magnitude of the effects of an ion thruster plume on S-band signals is measured. Modeling techniques are developed to predict the effects. Results show that the RF signal transmitted through an ion thruster plume is reduced in amplitude and shifted in phase. An increase in noise is also experienced.

  14. Effects of plume-scale versus grid-scale treatment of aircraft exhaust photochemistry

    NASA Astrophysics Data System (ADS)

    Cameron, Mary A.; Jacobson, Mark Z.; Naiman, Alexander D.; Lele, Sanjiva K.

    2013-11-01

    is a study to examine the impact of modeling photochemistry from aircraft emissions in an expanding plume versus at the grid scale in an atmospheric model. Differences in model treatments for a single flight occurred at all altitudes during takeoff, cruise, and landing. After 10 h, the plume treatment decreased grid-scale ozone production by 33%, methane destruction by 30%, and carbon monoxide destruction by 32% at cruise altitude compared with the grid-scale treatment. The plume treatment changed the odd nitrogen partitioning by ~10%. For multiple overlapping flights at cruise altitude, final ozone, methane, and carbon monoxide perturbations decreased by 77, 68, and 74%, respectively, compared with the grid-scale treatment. Enhanced mixing with ambient air reduced the plume-scale and grid-scale differences. The persistent differences in photochemical activity indicate that individual plume treatment should be incorporated into 3-D modeling studies.

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

    NASA Technical Reports Server (NTRS)

    Hawthorne, P. J.

    1976-01-01

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

  16. 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.

  17. 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).

  18. 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).

  19. 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.

  20. 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.

  1. 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.

  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; 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

  3. 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

  4. Study Of Heating Of The Base Region Of A Rocket

    NASA Technical Reports Server (NTRS)

    Ascoli, Edward P.; Heiba, Adel A.; Hsu, Yann-Fu; Lagnado, Ronald R.; Lynch, Edward D.; Ungewitter, Ronald J.

    1994-01-01

    Report describes theoretical study of heating in base region of proposed rocket called "NLS 1.5 stage reference vehicle." Study employed approach based on computational fluid dynamics (CFD). Involved numerical simulations of flow field in base region and in main exhaust plume of cluster of six engines with heat shields.

  5. Multiple dopant injection system for small rocket engines

    NASA Astrophysics Data System (ADS)

    Sakala, G. G.; Raines, N. G.

    1992-07-01

    The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

  6. 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.

  7. 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.

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

    NASA Technical Reports Server (NTRS)

    Goswami, Kisholoy

    2011-01-01

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

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Comparison of the particle size distribution of heavy-duty diesel exhaust using a dilution tailpipe sampler and an in-plume sampler during on-road operation.

    PubMed

    Brown, J E; Clayton, M J; Harris, D B; King, F G

    2000-08-01

    Originally constructed to develop gaseous emission factors for heavy-duty diesel trucks, the U.S. Environmental Protection Agency's (EPA) On-Road Diesel Emissions Characterization Facility has been modified to incorporate particle measurement instrumentation. An electrical low-pressure impactor designed to continuously measure and record size distribution data was used to monitor the particle size distribution of heavy-duty diesel truck exhaust. For this study, which involved a high-mileage (900,000 mi) truck running at full load, samples were collected by two different methods. One sample was obtained directly from the exhaust stack using an adaptation of the University of Minnesota's air-ejector-based mini-dilution sampler. The second sample was pulled from the plume just above the enclosed trailer, at a point approximately 11 m from the exhaust discharge. Typical dilution ratios of about 300:1 were obtained for both the dilution and plume sampling systems. Hundreds of particle size distributions were obtained at each sampling location. These were compared both selectively and cumulatively to evaluate the performance of the dilution system in simulating real-world exhaust plumes. The data show that, in its current residence-time configuration, the dilution system imposes a statistically significant bias toward smaller particles, with substantially more nanoparticles being collected than from the plume sample. PMID:11002602

  14. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2010-10-01

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

  16. Approach to SSME health monitoring. III - Exhaust plume emission spectroscopy: Recent results and detailed analysis

    NASA Astrophysics Data System (ADS)

    Tejwani, Gopal D.; van Dyke, David B.; Bircher, Felix E.

    1993-06-01

    Spectral data for two recent A-1 test firings, 901-717 and 901-718, obtained from an Optical Multichannel Analyzer and an Optical Plume Anomaly Detector, are presented. The spectral data encompasses the database of SSME critical components and materials and the spectral database for the SSME related elements and materials. Relatively strong and continuous emissions from Cr and Fe atomic transitions were observed starting at engine start plus 494 s and persisting until the engine shut off at engine start plus 520 s. These emissions are considered to be emanated from the SSME material AISI 440C, which is traced to high pressure turbopump bearings.

  17. Approach to SSME health monitoring. III - Exhaust plume emission spectroscopy: Recent results and detailed analysis

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.; Van Dyke, David B.; Bircher, Felix E.

    1993-01-01

    Spectral data for two recent A-1 test firings, 901-717 and 901-718, obtained from an Optical Multichannel Analyzer and an Optical Plume Anomaly Detector, are presented. The spectral data encompasses the database of SSME critical components and materials and the spectral database for the SSME related elements and materials. Relatively strong and continuous emissions from Cr and Fe atomic transitions were observed starting at engine start plus 494 s and persisting until the engine shut off at engine start plus 520 s. These emissions are considered to be emanated from the SSME material AISI 440C, which is traced to high pressure turbopump bearings.

  18. 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.

  19. 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

  20. 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

  1. Low altitude plume impingement handbook

    NASA Technical Reports Server (NTRS)

    Smith, Sheldon D.

    1991-01-01

    Plume Impingement modeling is required whenever an object immersed in a rocket exhaust plume must survive or remain undamaged within specified limits, due to thermal and pressure environments induced by the plume. At high altitudes inviscid plume models, Monte Carlo techniques along with the Plume Impingement Program can be used to predict reasonably accurate environments since there are usually no strong flowfield/body interactions or atmospheric effects. However, at low altitudes there is plume-atmospheric mixing and potential large flowfield perturbations due to plume-structure interaction. If the impinged surface is large relative to the flowfield and the flowfield is supersonic, the shock near the surface can stand off the surface several exit radii. This results in an effective total pressure that is higher than that which exists in the free plume at the surface. Additionally, in two phase plumes, there can be strong particle-gas interaction in the flowfield immediately ahead of the surface. To date there have been three levels of sophistication that have been used for low altitude plume induced environment predictions. Level 1 calculations rely on empirical characterizations of the flowfield and relatively simple impingement modeling. An example of this technique is described by Piesik. A Level 2 approach consists of characterizing the viscous plume using the SPF/2 code or RAMP2/LAMP and using the Plume Impingement Program to predict the environments. A Level 3 analysis would consist of using a Navier-Stokes code such as the FDNS code to model the flowfield and structure during a single calculation. To date, Level 1 and Level 2 type analyses have been primarily used to perform environment calculations. The recent advances in CFD modeling and computer resources allow Level 2 type analysis to be used for final design studies. Following some background on low altitude impingement, Level 1, 2, and 3 type analysis will be described.

  2. 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.

  3. 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.

  4. 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.

  5. 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…

  6. 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.

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

    NASA Technical Reports Server (NTRS)

    Capo, M. A.; Mickle, R.

    1975-01-01

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

  8. 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.

  9. 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.

  10. 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.

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

    NASA Technical Reports Server (NTRS)

    Love, D. A.

    1978-01-01

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

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. An analytical analysis of the dispersion predictions for effluents from the Saturn 5 and Scout-Algol 3 rocket exhausts

    NASA Technical Reports Server (NTRS)

    Stephens, J. B.; Susko, M.; Kaufman, J. W.; Hill, C. K.

    1973-01-01

    Predictions of the spatial concentration mapping of the potentially toxic constituents of the exhaust effluents from a launch of a Saturn 5 and of a Scout-Algol 3 vehicle utilizing the NASA/MSFC Multilayer Diffusion Program are provided. In the case of the Saturn 5, special attention was given to the concentration fields of carbon monoxide with a correlation of carbon dioxide concentrations. The Scout-Algol 3 provided an example of the centerline concentrations of hydrogen chloride, carbon monoxide, and alumina under typical meteorological conditions. While these results define the specific environmental impact of these two launches under the meteorological conditions existing during launches, they also provide a basis for the empirical monitoring of the constituents of the exhaust effluents of these vehicles.

  17. Ablative Rocket Deflector Testing and Computational Modeling

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

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

    EPA Science Inventory

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

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. The dominant effect of alumina on nearfield plume radiation

    NASA Astrophysics Data System (ADS)

    Laredo, David; Netzer, David W.

    1993-11-01

    Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Combustion of aluminum produces condensed alumina particles. Besides causing performance losses in the nozzle, the condensed Al2O3 particles are the major source of primary smoke in the exhaust plume. The particulate matter can also have major effects upon the plume i.r. signature. High number densities of particles can block gas-phase radiation from the plume. They can also be the source of radiation, especially the larger particles which exit the nozzle not in thermal equilibrium with the gas. In the past, the expected effects of particle size on the plume i.r. signature have been determined almost exclusively from predictions made with flow and radiation codes. The aim of the present work was to investigate the role of the Al/Al2O3 particles from a highly loaded solid propellant (up to 16% in weight) on the plume radiation of a small rocket motor (5 cm in diameter). The spatial variation of particle size distribution was simultaneously measured with the overall radiation of a portion of the plume in the i.r. band (3.5-5.0 microns). In micro-motors, operating with highly aluminized solid propellant, the condensed particles in the near exhaust plume were the major source of radiation in the 3.5-5 micron wavelength band. Motors with longer residence time and operating at medium chamber pressures produced more particles in the micron sized range. The role of after burning was predominately confined to reheating of the alumina particles to a higher temperature, at which the condensed Al2O3 radiated more than gaseous species. Even with 30% Al2O3 in the plume, the plume of small motors can be considered as approximately conical in shape, with volume distributed radiating sources. Motor conditions producing larger particles in the plume core were thus found to increase plume radiation from that region. The overall apparent emissivity of the plume was between 0

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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

  13. 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.

  14. 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.

  15. Measured particulate behavior in a subscale solid propellant rocket motor

    NASA Astrophysics Data System (ADS)

    Brennan, W. D.; Hovland, D. L.; Netzer, D. W.

    1992-10-01

    Particulate matter are sized in the exhaust nozzle and plume of small rocket motors of varying geometry to assess the effects of the expansion process on particle size. Both converging and converging-diverging nozzles are considered, and particle sizing is accomplished at pressures of up to 4.36 MPa with aluminum loadings of 2.0 and 4.7 percent. An instrument based on Fraunhofer diffraction is used to measure the particle-size distributions showing that: (1) high burning rates reduce particle agglomeration and increase C* efficiency; (2) high pressures lead to small and monomodal D32 entering the nozzle; and (3) D32 sizes increase appreciably at the tailoff. Some variations in plume signature are theorized to be caused by the tailoff phenomenon, and particle collisions and/or surface effects in the nozzle convergence are suggested by the reduced number of larger particles at the nozzle convergence.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

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

    NASA Astrophysics Data System (ADS)

    Nakashima, Yuki; Heki, Kosuke

    2014-07-01

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

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. 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)

  9. A tandem mirror plasma source for a hybrid plume plasma propulsion concept

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

    This paper describes a tandem mirror magnetic plasma confinement device to be considered as a hot plasma source for the hybrid plume rocket concept. The hot plasma from this device is injected into an exhaust duct, which will interact with an annular layer of hypersonic neutral gas. Such a device can be used to study the dynamics of the hybrid plume and to experimentally verify the numerical predictions obtained with computer codes. The basic system design is also geared toward being lightweight and compact, as well as having high power density (i.e., several kW/sq cm) at the exhaust. This feature is aimed toward the feasibility of 'space testing'. The plasma is heated by microwaves. A 50 percent heating efficiency can be obtained by using two half-circle antennas. The preliminary Monte Carlo modeling of test particles result reported here indicates that interaction does take place in the exhaust duct. Neutrals gain energy from the ion, which confirms the hybrid plume concept.

  10. 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.

  11. Ionospheric hole behind an ascending rocket observed with a dense GPS array

    NASA Astrophysics Data System (ADS)

    Furuya, T.; Heki, K.

    2008-03-01

    An ascending liquid-fuel rocket is known to make a hole in the ionosphere, or localized electron depletion, by leaving behind large amounts of neutral molecules (e.g. water) in the exhaust plume. Such a hole was made by the January 24, 2006 launch of an H-IIA rocket from Tanegashima, Southwestern Japan, and here we report its observation with a dense array of Global Positioning System receivers as a sudden and temporary decrease of total electron content. The observed disturbances have been compared with a simple numerical model incorporating the water diffusion and chemical reactions in the ionosphere. The substantial vanishing of the ionosphere lasted more than one hour, suggesting its application as a window for ground-based radio astronomical observations at low frequencies.

  12. Supersonic Rocket Thruster Flow Predicted by Numerical Simulation

    NASA Technical Reports Server (NTRS)

    Davoudzadeh, Farhad

    2004-01-01

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

  13. 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

  14. ASRM plume deflector analysis program

    NASA Astrophysics Data System (ADS)

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

    1992-07-01

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

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Air admixture to exhaust jets

    NASA Technical Reports Server (NTRS)

    Sanger, Eugen

    1953-01-01

    The problem of thrust increase by air admixture to exhaust jets of rockets, turbojet, ram- and pulse-jet engines is investigated theoretically. The optimum ratio of mixing chamber pressure to ambient pressure and speed range for thrust increase due to air admixture is determined for each type of jet engine.

  20. 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.

  1. 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.

  2. 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.

  3. Helping HAN for hybrid rockets

    NASA Astrophysics Data System (ADS)

    Ramohalli, Kumar; Dowler, Warren

    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.

  4. Predicted rocket and shuttle effects on stratospheric ozone

    NASA Technical Reports Server (NTRS)

    Harwood, Robert S.; Karol, Igor L.; Jackman, Charles H.; Qiu, Lian-Xiong; Prather, Michael J.; Pyle, John A.

    1991-01-01

    The major chemical effluents of either solid- or liquid-fueled rockets that can potentially perturb stratospheric ozone include chlorine compounds (HCl), nitrogen compounds (NO(x)), and hydrogen compounds (H2 and H2O). Radicals (Cl, ClO, H, OH, HO2, NO, and NO2) formed directly or indirectly from rocket exhaust can cause the catalytic destruction of ozone. Other exhaust compounds that could presumably lead to ozone destruction either by direct reaction with ozone or by providing a surface for heterogeneous processes include the particulates Al2O3, ice, and soot. These topics are discussed in terms of the possible effects of rocket exhausts on stratospheric ozone.

  5. 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)

  6. 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.

  7. 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.

  8. 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)

  9. 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

  10. 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.

  11. 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.

  12. Preliminary investigations on improving air-augmented rocket performance

    NASA Astrophysics Data System (ADS)

    Anil, K. N.; Damodaran, K. A.

    1994-05-01

    Use of the Petal nozzle instead of the conventional conical nozzle as the primary stream representing fuel-rich gases exiting from a rocket nozzle has demonstrated considerable improvement in the performance of an air-augmented rocket. This can be attributed to the improved mixing of the hot, exhaust gases containing unburnt fuel with the surrounding airstream, and subsequent heat release.

  13. 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%.

  14. Four-nozzle benchmark wind tunnel model USA code solutions for simulation of multiple rocket base flow recirculation at 145,000 feet altitude

    NASA Astrophysics Data System (ADS)

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

    1993-07-01

    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.

  15. 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.

  16. 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

  17. 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

  18. 9. Building 202 test stand B and exhaust scrubber stack, ...

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

    9. Building 202 test stand B and exhaust scrubber stack, looking southwest from concrete apron north of Building 202 test cell. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  19. 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.

  20. Simulation of the Flow Field Associated with a Rocket Thruster Having an Attached Panel

    NASA Technical Reports Server (NTRS)

    Davoudzadeh, Farhad; Liu, Nan-Suey

    2003-01-01

    Two-dimensional inviscid and viscous numerical simulations are performed to predict the flow field induced by a H2-O2 rocket thruster and to provide insight into the heat load on the articles placed in the hot gas exhaust of the thruster under a variety of operating conditions, using the National Combustion Code (NCC). The simulations have captured physical details of the flow field, such as the plume formation and expansion, formation of the shock waves and their effects on the temperature and pressure distributions on the walls of the apparatus and the flat panel. Comparison between the computed results for 2-D and adiabatic walls and the related experimental measurements for 3-D and cooled walls shows that the results of the simulations are consistent with those obtained from the related rig tests.

  1. 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.

  2. 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.

  3. Supersonic-combustion rocket

    NASA Technical Reports Server (NTRS)

    Weber, R. J.; Franciscus, L. C. (Inventor)

    1973-01-01

    A supersonic combustion rocket is provided in which a small rocket motor is substituted for heavy turbo pumps in a conventional rocket engine. The substitution results in a substantial reduction in rocket engine weight. The flame emanating from the small rocket motor can act to ignite non-hypergolic fuels.

  4. Plasma Diagnostics Development for Advanced Rocket Engines

    NASA Astrophysics Data System (ADS)

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

    2000-10-01

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

  5. Devising rocket power for smaller engines

    SciTech Connect

    Burruss, R.

    1996-04-01

    Compact, high-power engines that burn fuel and oxygen could be made by winding copper tubing in a helix around boiler sections. With more than 1,000 horsepower per pound of engine weight, liquid-fueled rockets have the highest specific power of any engines designed for sustained operation. Yet those engines generally run for about only 1,000 seconds--nowhere near the sustained operation time for lower-power automotive and aircraft engines of more than 1,000 hours. In theory, at least, a fuel/oxygen rocket can be built that combines the best of both classes: high specific power (from perhaps two to 10 times that of a gas turbine) and a 1,000-hour service life. Such an engine would almost certainly be possible if the rocket`s exhaust gases could be simultaneously cooled and expanded by mixing water with the rocket`s exhaust and boiling it before it reaches the turbine. The technology itself is not new. variations of these rocket-turbine-type engines, for example, powered torpedoes during World War I. Some 30 years later, German V-2 rockets used fuel pumps, driven by the reaction of hydrogen peroxide with hydrocarbon fuels, to produce high-pressure steam that was directed against a turbine. Alternatively, fuel/oxygen combustion could produce steam to drive a piston engine. Either way, the challenge remains to construct a compact, long-service-life, high-specific-power boiler that burns fuel and oxygen. The new type of engine could be derived from recent research on electric vehicles (EVs).

  6. 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.

  7. Radiative forcing caused by rocket engine emissions

    NASA Astrophysics Data System (ADS)

    Ross, Martin N.; Sheaffer, Patti M.

    2014-04-01

    Space transportation plays an important and growing role in Earth's economic system. Rockets uniquely emit gases and particles directly into the middle and upper atmosphere where exhaust from hundreds of launches accumulates, changing atmospheric radiation patterns. The instantaneous radiative forcing (RF) caused by major rocket engine emissions CO2, H2O, black carbon (BC), and Al2O3 (alumina) is estimated. Rocket CO2 and H2O emissions do not produce significant RF. BC and alumina emissions, under some scenarios, have the potential to produce significant RF. Absorption of solar flux by BC is likely the main RF source from rocket launches. In a new finding, alumina particles, previously thought to cool the Earth by scattering solar flux back to space, absorb outgoing terrestrial longwave radiation, resulting in net positive RF. With the caveat that BC and alumina microphysics are poorly constrained, we find that the present-day RF from rocket launches equals 16 ± 8 mW m-2. The relative contributions from BC, alumina, and H2O are 70%, 28%, and 2%. respectively. The pace of rocket launches is predicted to grow and space transport RF could become comparable to global aviation RF in coming decades. Improved understanding of rocket emission RF requires more sophisticated modeling and improved data describing particle microphysics.

  8. ASRM subscale plume deflector testing

    NASA Astrophysics Data System (ADS)

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

    1992-07-01

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

  9. Dispersion of turbojet engine exhaust in flight

    NASA Technical Reports Server (NTRS)

    Holdeman, J. D.

    1973-01-01

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

  10. Plume detachment from a magnetic nozzle

    SciTech Connect

    Deline, Christopher A.; Bengtson, Roger D.; Breizman, Boris N.; Tushentsov, Mikhail R.; Jones, Jonathan E.; Chavers, D. Greg; Dobson, Chris C.; Schuettpelz, Branwen M.

    2009-03-15

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

  11. 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.

  12. 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...

  13. 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...

  14. 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...

  15. 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...

  16. 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...

  17. IUS solid rocket motor contamination prediction methods

    NASA Technical Reports Server (NTRS)

    Mullen, C. R.; Kearnes, J. H.

    1980-01-01

    A series of computer codes were developed to predict solid rocket motor produced contamination to spacecraft sensitive surfaces. Subscale and flight test data have confirmed some of the analytical results. Application of the analysis tools to a typical spacecraft has provided early identification of potential spacecraft contamination problems and provided insight into their solution; e.g., flight plan modifications, plume or outgassing shields and/or contamination covers.

  18. 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.

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

    NASA Astrophysics Data System (ADS)

    Lehman, Matthew Kurt

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

  20. 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.

  1. 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.

  2. 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...

  3. 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.

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

    NASA Astrophysics Data System (ADS)

    Ross, M.

    2009-12-01

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

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Solar Thermal Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Sercel, J. C.

    1986-01-01

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

  12. American Rocket Society

    NASA Technical Reports Server (NTRS)

    2004-01-01

    In addition to Dr. Robert Goddard's pioneering work, American experimentation in rocketry prior to World War II grew, primarily in technical societies. This is an early rocket motor designed and developed by the American Rocket Society in 1932.

  13. Dragonfly directional sensor versus rocket-propelled grenades

    NASA Astrophysics Data System (ADS)

    Geary, Joseph; Blackwell, Lisa

    2015-02-01

    The Dragonfly directional sensor was deployed at the Army's Yuma Proving Grounds for preliminary field tests against rocket-propelled grenades. This wide-field (nonimaging) sensor's purpose was to angularly locate the latter's launch plume. These tests successfully demonstrated proof-of-concept.

  14. Hybrid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    1994-01-01

    A 10,000-pound thrust hybrid rocket motor is tested at Stennis Space Center's E-1 test facility. A hybrid rocket motor is a cross between a solid rocket and a liquid-fueled engine. It uses environmentally safe solid fuel and liquid oxygen.

  15. 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.

  16. 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

  17. 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).

  18. 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.

  19. Dust Plumes off Libya

    NASA Technical Reports Server (NTRS)

    2007-01-01

    Two-toned dust plumes blew northward off the coast of Libya on October 26, 2007, as the Moderate Resolution Imaging Spectroradiometer (MODIS) on NASA's Terra satellite took this picture. While plumes in the west are beige, reminiscent of the Sahara's sands, the plumes in the east are distinctly darker. The differences in color can be traced to the plumes's varied origins.

  20. 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.

  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. 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 Astrophysics Data System (ADS)

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

    1992-07-01

    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.

  3. Exhaust emission control apparatus

    SciTech Connect

    Eng, J.W.

    1991-09-24

    This patent describes an exhaust control apparatus for muffling noise and treating odors and pollutants, including solid particulate and gases in the exhaust of an internal combustion engine. It comprises an exhaust inlet tube for receiving the exhaust generated by an internal combustion engine; a cyclone barrier concentrically surrounding the exhaust inlet tube, a ring cavity between the cyclone tube and exhaust inlet tube defining a cyclone chamber in which the exhaust is treated; means for directing the exhaust from the exhaust inlet tube into the cyclone chamber; electrode means having small openings through which the exhaust passes to enter the cyclone chamber, the electrode means generating electrostatic forces which charge the solid particulate in the exhaust, ionize air and generate ozone in the cyclone chamber near the electrode; means for injecting air into the cyclone chamber causing centrifugal flow of the air and the exhausted within the cyclone chamber and increasing a dwell time of the exhaust within the cyclone chamber.

  4. 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.

  5. 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.

  6. 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.

  7. Prediction of fluctuating pressure environments associated with plume-induced separated flow fields

    NASA Technical Reports Server (NTRS)

    Plotkin, K. J.

    1973-01-01

    The separated flow environment induced by underexpanded rocket plumes during boost phase of rocket vehicles has been investigated. A simple semi-empirical model for predicting the extent of separation was developed. This model offers considerable computational economy as compared to other schemes reported in the literature, and has been shown to be in good agreement with limited flight data. The unsteady pressure field in plume-induced separated regions was investigated. It was found that fluctuations differed from those for a rigid flare only at low frequencies. The major difference between plume-induced separation and flare-induced separation was shown to be an increase in shock oscillation distance for the plume case. The prediction schemes were applied to PRR shuttle launch configuration. It was found that fluctuating pressures from plume-induced separation are not as severe as for other fluctuating environments at the critical flight condition of maximum dynamic pressure.

  8. 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.

  9. 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.

  10. 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.

  11. Spacecraft plume interactions with the magnetosphere plasma environment in geostationary Earth orbit

    NASA Astrophysics Data System (ADS)

    Stephani, K. A.; Boyd, I. D.

    2016-02-01

    Particle-based kinetic simulations of steady and unsteady hydrazine chemical rocket plumes are presented in a study of plume interactions with the ambient magnetosphere in geostationary Earth orbit. The hydrazine chemical rocket plume expands into a near-vacuum plasma environment, requiring the use of a combined direct simulation Monte Carlo/particle-in-cell methodology for the rarefied plasma conditions. Detailed total and differential cross sections are employed to characterize the charge exchange reactions between the neutral hydrazine plume mixture and the ambient hydrogen ions, and ion production is also modeled for photoionization processes. These ionization processes lead to an increase in local plasma density surrounding the spacecraft owing to a partial ionization of the relatively high-density hydrazine plume. Results from the steady plume simulations indicate that the formation of the hydrazine ion plume are driven by several competing mechanisms, including (1) local depletion and (2) replenishing of ambient H+ ions by charge exchange and thermal motion of 1 keV H+ from the ambient reservoir, respectively, and (3) photoionization processes. The self-consistent electrostatic field forces and the geostationary magnetic field have only a small influence on the dynamics of the ion plume. The unsteady plume simulations show a variation in neutral and ion plume dissipation times consistent with the variation in relative diffusion rates of the chemical species, with full H2 dissipation (below the ambient number density levels) approximately 33 s after a 2 s thruster burn.

  12. A plume spectroscopy system for flight applications

    NASA Astrophysics Data System (ADS)

    Makel, D. B.; Petersen, T. V.; Duncan, D. B.; Madzsar, G. C.

    1993-06-01

    An operational plume spectroscopy system will be an important element of any rocket engine health management system (HMS). The flight capable FPI spectrometer will enable prognosis and response to incipient rocket engine failures as well as diagnosis of wear and degradation for on-condition maintenance. Spectrometer application to development programs, such as the Space Lifter, NASP, and SSTO, will reduce program risks, allow better adherence to schedules and save money by reducing or eliminating redesign and test costs. The diagnostic capability of a proven, calibrated spectrometer will enhance post-burn certification of high value, reusable engines, such as the Space Shuttle Main Engine (SSME), where life and reliability are key cost drivers. This paper describes a prototype FPI spectrometer for demonstration and validation testing on NASA's Technology Test Bed Engine (TTBE) at Marshall Space Flight Center. The TTBE test unit is designed with flight prototype optics and a commercial off-the-shelf data processing system.

  13. 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

  14. 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

  15. 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

  16. 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

  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. Postal Rocket Stamps

    NASA Technical Reports Server (NTRS)

    2004-01-01

    In the 19th Century, experiments in America, Europe, and elsewhere attempted to build postal rockets to deliver mail from one location to another. The idea was more novel than successful. Many stamps used in these early postal rockets have become collector's items.

  20. 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)

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. The large motor plume/material impingement test program at AEDC utilizing the four selected propellants from the small motor tests at MSFC (July, August, and September 1974), section 3

    NASA Technical Reports Server (NTRS)

    1976-01-01

    Efforts made to determine the vulnerability of Orbiter and ET materials located at various positions within exhaust plumes from test SSRM's using four different propellant formulations are discussed. Data also cover the effect on TPS materials from a single SSRM plume and dual SSRM plumes, and definitions of test SSRM plume environment at material specimen locations.

  7. 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.

  8. 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.

  9. Dust Plume off Mauritania

    NASA Technical Reports Server (NTRS)

    2007-01-01

    A thick plume of dust blew off the coast of Mauritania in western Africa on October 2, 2007. The Moderate Resolution Imaging Spectroradiometer (MODIS) on NASA's Aqua satellite observed the dust plume as it headed toward the southwest over the Atlantic Ocean. In this image, the dust varies in color from nearly white to medium tan. The dust plume is easier to see over the dark background of the ocean, but the plume stretches across the land surface to the east, as well. The dust plume's structure is clearest along the coastline, where relatively clear air pockets separate distinct puffs of dust. West of that, individual pillows of dust push together to form a more homogeneous plume. Near its southwest tip, the plume takes on yet another shape, with stripes of pale dust fanning out toward the northwest. Occasional tiny white clouds dot the sky overhead, but skies are otherwise clear.

  10. 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.

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

    NASA Astrophysics Data System (ADS)

    Waung, Timothy S.

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

  12. 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.

  13. Solid Rocket Motor Backflow Analysis For CONTOUR Mishap Investigation

    NASA Astrophysics Data System (ADS)

    Woronowicz, Michael

    2005-05-01

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

  14. 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).

  15. 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.

  16. 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

  17. 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.

  18. 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.

  19. ASTRID rocket flight test

    SciTech Connect

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

    1994-07-01

    On February 4, 1994, we successfully flight tested the ASTRID rocket from Vandenberg Air Force Base. The technology for this rocket originated in the Brilliant Pebbles program and represents a five-year development effort. This rocket demonstrated how our new pumped-propulsion technology-which reduced the total effective engine mass by more than one half and cut the tank mass to one fifth previous requirements-would perform in atmospheric flight. This demonstration paves the way for potential cost-effective uses of the new propulsion system in commercial aerospace vehicles, exploration of the planets, and defense applications.

  20. 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

  1. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

  2. 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.

  3. Hybrid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    1994-01-01

    Stennis Space Center conducts a test on a hybrid rocket motor fed by a liquid oxygen turbopump. The test occurred at the E-1 test facility. The test was believed to be the first of its kind in the world.

  4. 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...

  5. Robust Rocket Engine Concept

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.

    1995-01-01

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

  6. 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)

  7. 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.

  8. 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

  9. 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.

  10. Investigation of the Rocket Induced Flow Field in a Rectangular Duct

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

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

  11. 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.

  12. Rocket Engine Oscillation Diagnostics

    NASA Technical Reports Server (NTRS)

    Nesman, Tom; Turner, James E. (Technical Monitor)

    2002-01-01

    Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

  13. 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 ...

  14. 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.

  15. 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

  16. 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.

  17. COOLING TOWER PLUME MODEL

    EPA Science Inventory

    A review of recently reported cooling tower plume models yields none that is universally accepted. The entrainment and drag mechanisms and the effect of moisture on the plume trajectory are phenomena which are treated differently by various investigators. In order to better under...

  18. 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.

  19. 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.

  20. 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.

  1. Two stage turbine for rockets

    NASA Astrophysics Data System (ADS)

    Veres, Joseph P.

    1993-11-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

  2. Two stage turbine for rockets

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.

    1993-01-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

  3. 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)

  4. 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

  5. Three Dimensional Monitoring of Stack Plume Dynamics by a Scanning Mie Lidar System as a Plume Watchdog Station

    NASA Astrophysics Data System (ADS)

    Saito, Yasunori; Kurata, Hidehumi; Hara, Yuta; Kobayashi, Fumitoshi; Kawahara, Takuya; Nomura, Akio

    A scanning lidar system was developed to watch the nighttime diffusion process of plume from a smokestack of a large incinerator located around 3 km from the system. Observed data sets were visualized as three dimensional images in which could be seen the diffusion pattern from any direction, and this made it easy to investigate the exhaust dynamics. Observation results showed that the original plume extended at least 1.6 km, where there was a residential area, from the smokestack increasing in diameter to about 500 m. High density aerosols originating from the smokestack were measured in that area even at midnight. The lidar system performance as a plume watchdog station was discussed from the standpoint of publichealth-related plume monitoring.

  6. 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.

  7. Rocket motor aeroacoustics

    NASA Astrophysics Data System (ADS)

    Hegde, U. G.; Strahle, W. C.

    1983-10-01

    Vibration problems in solid propellant rocket motors are investigated. A class of interior flows modelled to simulate flow conditions inside rocket motor cavities is considered. Turbulence generated pressure fluctuations are shown to consist of two components - acoustic and hydrodynamics. The Bernoulli enthalpy theory of aeroacoustics is employed to extract acoustic pressure spectra from experimentally obtained turbulence data and acoustic impedance values at flow boundaries. The effects of turbulence intensities, sidewall acoustic impedance, axial mass blowing distribution, length to diameter ratio of the cavity and different mass flux on the acoustic pressure level are investigated. Typical pressure levels, under rocket motor conditions, are calculated using the A/B model of propellant response. Estimates of the hydrodynamic component of the pressure fluctuation are provided for the case of fully developed turbulent pipe flow terminated by a choked nozzle.

  8. 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

  9. 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.

  10. Advanced liquid rockets

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

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

  11. 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.

  12. 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…

  13. 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.

  14. 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).

  15. 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...

  16. 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.

  17. 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

  18. 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

  19. 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.

  20. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust. The test was the first test ever anywhere outside Russia of a Russian designed and built engine.

  1. Hybrid rocket instability

    NASA Technical Reports Server (NTRS)

    Greiner, B.; Frederick, R. A., Jr.

    1993-01-01

    The paper provides a brief review of theoretical and experimental studies concerned with hybrid rocket instability. The instabilities discussed include atomization and mixing instabilities, chuffing instabilities, pressure coupled combustion instabilities, and vortex shedding. It is emphasized that the future use of hybrid motor systems as viable design alternatives will depend on a better understanding of hybrid instability.

  2. Hybrid rocket instability

    NASA Astrophysics Data System (ADS)

    Greiner, B.; Frederick, R. A., Jr.

    1993-06-01

    The paper provides a brief review of theoretical and experimental studies concerned with hybrid rocket instability. The instabilities discussed include atomization and mixing instabilities, chuffing instabilities, pressure coupled combustion instabilities, and vortex shedding. It is emphasized that the future use of hybrid motor systems as viable design alternatives will depend on a better understanding of hybrid instability.

  3. 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.

  4. 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)

  5. 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.

  6. 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…

  7. 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.

  8. 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.

  9. 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…

  10. 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.

  11. 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.

  12. Electrification of volcanic plumes

    NASA Astrophysics Data System (ADS)

    Mather, T. A.; Harrison, R. G.

    2006-07-01

    Volcanic lightning, perhaps the most spectacular consequence of the electrification of volcanic plumes, has been implicated in the origin of life on Earth, and may also exist in other planetary atmospheres. Recent years have seen volcanic lightning detection used as part of a portfolio of developing techniques to monitor volcanic eruptions. Remote sensing measurement techniques have been used to monitor volcanic lightning, but surface observations of the atmospheric electric Potential Gradient (PG) and the charge carried on volcanic ash also show that many volcanic plumes, whilst not sufficiently electrified to produce lightning, have detectable electrification exceeding that of their surrounding environment. Electrification has only been observed associated with ash-rich explosive plumes, but there is little evidence that the composition of the ash is critical to its occurrence. Different conceptual theories for charge generation and separation in volcanic plumes have been developed to explain the disparate observations obtained, but the ash fragmentation mechanism appears to be a key parameter. It is unclear which mechanisms or combinations of electrification mechanisms dominate in different circumstances. Electrostatic forces play an important role in modulating the dry fall-out of ash from a volcanic plume. Beyond the local electrification of plumes, the higher stratospheric particle concentrations following a large explosive eruption may affect the global atmospheric electrical circuit. It is possible that this might present another, if minor, way by which large volcanic eruptions affect global climate. The direct hazard of volcanic lightning to communities is generally low compared to other aspects of volcanic activity.

  13. 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.

  14. 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.

  15. 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.

  16. Comparison of ACCENT 2000 Shuttle Plume Data with SIMPLE Model Predictions

    NASA Astrophysics Data System (ADS)

    Swaminathan, P. K.; Taylor, J. C.; Ross, M. N.; Zittel, P. F.; Lloyd, S. A.

    2001-12-01

    The JHU/APL Stratospheric IMpact of PLume Effluents (SIMPLE)model was employed to analyze the trace species in situ composition data collected during the ACCENT 2000 intercepts of the space shuttle Space Transportation Launch System (STS) rocket plume as a function of time and radial location within the cold plume. The SIMPLE model is initialized using predictions for species depositions calculated using an afterburning model based on standard TDK/SPP nozzle and SPF plume flowfield codes with an expanded chemical kinetic scheme. The time dependent ambient stratospheric chemistry is fully coupled to the plume species evolution whose transport is based on empirically derived diffusion. Model/data comparisons are encouraging through capturing observed local ozone recovery times as well as overall morphology of chlorine chemistry.

  17. Scalable hardbody and plume optical signatures

    NASA Astrophysics Data System (ADS)

    Crow, Dennis R.; Hawes, Fred; Braunstein, Matthew; Coker, Charles F.; Smith, Thomas, Jr.

    2004-08-01

    The Fast Line-of-sight Imagery for Target and Exhaust Signatures (FLITES) is a High Performance Computing (HPC-CHSSI) and Missile Defense Agency (MDA) funded effort that provides a scalable program to compute highly resolved temporal, spatial, and spectral hardbody and plume optical signatures. Distributed processing capabilities are included to allow complex, high fidelity, solutions to be generated quickly generated. The distributed processing logic includes automated load balancing algorithms to facilitate scalability using large numbers of processors. To enhance exhaust plume optical signature capabilities, FLITES employs two different radiance transport algorithms. The first algorithm is the traditional Curtis-Godson bandmodel approach and is provided to support comparisons to historical results and high-frame rate production requirements. The second algorithm is the Quasi Bandmodel Line-by-line (QBL) approach, which uses randomly placed "cloned" spectral lines to yield highly resolved radiation spectra for increased accuracy while maintaining tractable runtimes. This capability will provide a significant advancement over the traditional SPURC/SIRRM radiance transport methodology.

  18. Method of hybrid plume plasma propulsion

    NASA Technical Reports Server (NTRS)

    Chang, Franklin R. (Inventor)

    1990-01-01

    A technique for producing thrust by generating a hybrid plume plasma exhaust is disclosed. A plasma flow is generated and introduced into a nozzle which features one or more inlets positioned to direct a flow of neutral gas about the interior of the nozzle. When such a neutral gas flow is combined with the plasma flow within the nozzle, a hybrid plume is constructed including a flow of hot plasma along the center of the nozzle surrounded by a generally annular flow of neutral gas, with an annular transition region between the pure plasma and the neutral gas. The temperature of the outer gas layer is below that of the pure plasma and generally separates the pure plasma from the interior surfaces of the nozzle. The neutral gas flow both insulates the nozzle walls from the high temperatures of the plasma flow and adds to the mass flow rate of the hybrid exhaust. The rate of flow of neutral gas into the interior of the nozzle may be selectively adjusted to control the thrust and specific impulse of the device.

  19. 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.

  20. 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.

  1. Electrification of volcanic plumes

    NASA Astrophysics Data System (ADS)

    Mather, T. A.; Harrison, R. G.

    We present a review of our current understanding of the electrification of volcanic plumes on Earth and discuss the possible implications both in terms of the volcanic monitoring, early Earth evolution and planetary exploration. Volcanic lightning is perhaps the most spectacular consequence of the electrification of volcanic plumes. Recent years have seen volcanic lightning detection used as part of a portfolio of developing techniques to monitor volcanic eruptions. Remote sensing measurement techniques have been used to monitor volcanic lightning, but surface observations of the atmospheric electric Potential Gradient (PG) and the charge carried on volcanic ash also show that many volcanic plumes, whilst not sufficiently electrified to produce lightning, have detectable electrification exceeding that of their surrounding environment. Electrification has only been observed associated with ash-rich explosive plumes, but there is little evidence that the composition of the ash is critical to its occurrence. Different conceptual theories for charge generation and separation in volcanic plumes have been developed to explain the disparate observations obtained, but the ash fragmentation mechanism appears to be a key parameter. It is unclear which mechanisms or combinations of electrification mechanisms dominate in different circumstances. Electrostatic forces play an important role in modulating the dry fall-out of ash from a volcanic plume. Beyond the local electrification of plumes, the higher stratospheric particle concentrations following a large explosive eruption may affect the global atmospheric electrical circuit. It is possible that this might present another, if minor, way by which large volcanic eruptions affect global climate. Volcanic lightning has been implicated in a number of ways in the origin of life on Earth, and may also exist in other planetary atmospheres where measurements of its occurrence might give clues about the nature of volcanism on other

  2. Immune Exhaustion and Transplantation.

    PubMed

    Sanchez-Fueyo, A; Markmann, J F

    2016-07-01

    Exhaustion of lymphocyte function through chronic exposure to a high load of foreign antigen is well established for chronic viral infection and antitumor immunity and has been found to be associated with a distinct molecular program and characteristic cell surface phenotype. Although exhaustion has most commonly been studied in the context of CD8 viral responses, recent studies indicate that chronic antigen exposure may affect B cells, NK cells and CD4 T cells in a parallel manner. Limited information is available regarding the extent of lymphocyte exhaustion development in the transplant setting and its impact on anti-graft alloreactivity. By analogy to the persistence of a foreign virus, the large mass of alloantigen presented by an allograft in chronic residence could provide an ideal setting for exhausting donor-reactive T cells. The extent of T cell exhaustion occurring with various allografts, the kinetics of its development, whether exhaustion is influenced positively or negatively by different immunosuppressants, and the impact of exhaustion on graft survival and tolerance development remains a fertile area for investigation. Harnessing or encouraging the natural processes of exhaustion may provide a novel means to promote graft survival and transplantation tolerance. PMID:26729653

  3. 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.

  4. Diesel engine exhaust oxidizer

    SciTech Connect

    Kammel, R.A.

    1992-06-16

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

  5. Electric field and radio frequency measurements for rocket engine health monitoring applications

    NASA Technical Reports Server (NTRS)

    Valenti, Elizabeth L.

    1992-01-01

    Electric-field (EF) and radio-frequency (RF) emissions generated in the exhaust plumes of the diagnostic testbed facility thruster (DTFT) and the SSME are examined briefly for potential applications to plume diagnostics and engine health monitoring. Hypothetically, anomalous engine conditions could produce measurable changes in any characteristic EF and RF spectral signatures identifiable with a 'healthly' plumes. Tests to determine the presence of EF and RF emissions in the DTFT and SSME exhaust plumes were conducted. EF and RF emissions were detected using state-of-the-art sensors. Analysis of limited data sets show some apparent consistencies in spectral signatures. Significant emissions increases were detected during controlled tests using dopants injected into the DTFT.

  6. Electric field and radio frequency measurements for rocket engine health monitoring applications

    NASA Astrophysics Data System (ADS)

    Valenti, Elizabeth L.

    1992-10-01

    Electric-field (EF) and radio-frequency (RF) emissions generated in the exhaust plumes of the diagnostic testbed facility thruster (DTFT) and the SSME are examined briefly for potential applications to plume diagnostics and engine health monitoring. Hypothetically, anomalous engine conditions could produce measurable changes in any characteristic EF and RF spectral signatures identifiable with a 'healthly' plumes. Tests to determine the presence of EF and RF emissions in the DTFT and SSME exhaust plumes were conducted. EF and RF emissions were detected using state-of-the-art sensors. Analysis of limited data sets show some apparent consistencies in spectral signatures. Significant emissions increases were detected during controlled tests using dopants injected into the DTFT.

  7. 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.

  8. Exhaust purification apparatus

    SciTech Connect

    Shinzawa, M.; Ushimura, S.

    1987-05-05

    An exhaust purification apparatus is described for use in an internal combustion engine having an exhaust conduit through which exhaust particles are discharged together with exhaust gas to the atmosphere. Included is an outer shell having an inlet connected to the exhaust conduit and an outlet connected to the atmosphere. The outer shell contains a trap element and a regenerative burner located upstream of the trap element, the regenerative burner comprising: a cylindrical hollow member fixed to the liner and extending within a combustion chamber to define an evaporation chamber, a glow plug for igniting the mixture supplied into the evaporated chamber when actuated; and a control unit responsive to a regeneration requirement for actuating the glow plug and supplying an air-fuel mixture into the evaporation chamber through the mixture conduit.

  9. A Brilliant Plume

    NASA Technical Reports Server (NTRS)

    2007-01-01

    The Long Range Reconnaissance Imager (LORRI) on New Horizons captured another dramatic picture of Jupiter's moon Io and its volcanic plumes, 19 hours after the spacecraft's closest approach to Jupiter on Feb. 28, 2007. LORRI took this 75 millisecond exposure at 0035 Universal Time on March 1, 2007, when Io was 2.3 million kilometers (1.4 million miles) from the spacecraft.

    Io's dayside is deliberately overexposed to bring out faint details in the plumes and on the moon's night side. The continuing eruption of the volcano Tvashtar, at the 1 o'clock position, produces an enormous plume roughly 330 kilometers (200 miles) high, which is illuminated both by sunlight and 'Jupiter light.'

    The shadow of Io, cast by the Sun, slices across the plume. The plume is quite asymmetrical and has a complicated wispy texture, for reasons that are still mysterious. At the heart of the eruption incandescent lava, seen here as a brilliant point of light, is reminding scientists of the fire fountains spotted by the Galileo Jupiter orbiter at Tvashtar in 1999.

    The sunlit plume faintly illuminates the surface underneath. 'New Horizons and Io continue to astonish us with these unprecedented views of the solar system's most geologically active body' says John Spencer, deputy leader of the New Horizons Jupiter Encounter Science Team and an Io expert from Southwest Research Institute.

    Because this image shows the side of Io that faces away from Jupiter, the large planet does not illuminate the moon's night side except for an extremely thin crescent outlining the edge of the disk at lower right. Another plume, likely from the volcano Masubi, is illuminated by Jupiter just above this lower right edge. A third and much fainter plume, barely visible at the 2 o'clock position, could be the first plume seen from the volcano Zal Patera.

    As in other New Horizons images of Io, mountains catch the setting Sun just beyond the terminator (the line dividing day and night

  10. Advanced rocket propulsion

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1993-01-01

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

  11. 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.

  12. Hybrid rocket performance

    NASA Astrophysics Data System (ADS)

    Frederick, Robert A., Jr.

    1992-12-01

    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.

  13. 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.

  14. 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.

  15. Advanced rocket propulsion

    NASA Astrophysics Data System (ADS)

    Obrien, Charles J.

    1993-02-01

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

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Excitation by rockets

    NASA Technical Reports Server (NTRS)

    Tammadge, C. E.

    1975-01-01

    Standard methods of excitation are not always practical when a single mode of known frequency requires investigation. This form of investigation is often required on a modified aircraft. A new method of excitation was developed and proved in flight, which consists of firing small rocket charges attached to the aircraft structure. Damping values at gradually increasing airspeeds are obtained, as in Stick Jerk tests, and flutter speeds predicted.

  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. 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.

  3. 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…

  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. 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.

  6. Io Pele plume

    NASA Technical Reports Server (NTRS)

    2000-01-01

    Voyager 1 took this narrow-angle camera image on 5 March 1979 from a distance of 450,000 kilometers. At this geometry, the camera looks straight down through a volcanic plume at one of Io's most active volcanos, Pele. The large heart-shaped feature is the region where Pele's plume falls to the surface. At the center of the 'heart' is the small dark fissure that is the source of the eruption. The Voyager Project is managed by the Jet Propulsion Laboratory for NASA's Office of Space Science.

  7. Evidence for mantle plumes?

    PubMed

    Anderson, Don L; Natland, James H

    2007-11-22

    Geophysical hotspots have been attributed to partially molten asthenosphere, fertile blobs, small-scale convection and upwellings driven by core heat. Most are short-lived or too close together to be deeply seated, and do not have anomalous heat flow or temperature; many are related to tectonic features. Bourdon et al. investigate the dynamics of mantle plumes from uranium-series geochemistry and interpret their results as evidence for thermal plumes. Here we show why alternative mechanisms of upwelling and melting should be considered. PMID:18033248

  8. Exhaust gas afterburner

    SciTech Connect

    Hudson, S.J. Jr.

    1986-12-23

    This patent describes an exhaust gas afterburner device adapted for installation between an exhaust manifold and a corresponding portion of the engine block of an internal combustion engine. The device comprises: a spacer sandwiched between portions of two sheet metal members forming a gasket section of the device, the gasket section surrounding at least one exhaust gas port, a plenum section formed by remaining portions of the members, and wall sections defining passageways extending from the interior of the plenum section to the port and an air supply inlet on the plenum.

  9. 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.

  10. Simulation of low-density nozzle plumes in non-zero ambient pressures

    NASA Astrophysics Data System (ADS)

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

    1994-02-01

    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. 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.

  12. 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.

  13. 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. ...

  14. 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.

  15. 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.

  16. 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...

  17. 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.

  18. 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)

  19. 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)

  20. 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…