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Sample records for rocket exhaust plumes

  1. Particle Characterization in Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Callen, E. Eugene, Jr.; Fisher, J. Scott

    1997-01-01

    A method to characterize particles in rocket exhaust plumes is developed. The particle velocity, size, and material composition are determined from crater characteristics resulting from impacts into aluminum and copper targets passed through the plume. The targets are mounted on a steel arm approximately 21 inches (53 cm) long which is rotated through the plume at sufficient velocity to prevent material failure resulting from thermal effects. A Scanning Electron Microscope (SEM) with secondary x-ray detectors is used to determine the particle material, and a standard optical measurement microscope is used to determine the crater diameter and depth. The crater diameter and depth are used in turn, as inputs to a ballistics computer code to estimate the velocity and size of the particle. The target has a safe residence time in the plume of approximately 50 ms before reaching an unacceptably high temperature. The = must mach a velocity of 104 ft/s (32 m/s) before entering the plume to produce the design residence time of 20 ms. The arm is actuated by a torsion spring with a 5-inch (13 cm) outer diameter, 0.625-inch (16 mm wire diameter, and 11 coils. A prototype of the entire rocket exhaust particle impact characterization system (PICS) was constructed and statically tested.

  2. Prediction of thermal radiation from a rocket's exhaust plume

    NASA Technical Reports Server (NTRS)

    Ludwig, C. B.

    1969-01-01

    Data from absorption coefficients and fine-structure parameters measured for water vapor have been incorporated in an analytic program useful in evaluating heating by radiation from the exhaust plume of a large rocket.

  3. 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.

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

    NASA Technical Reports Server (NTRS)

    Hwang, B.; Pergament, H. S.

    1976-01-01

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

  5. High-speed schlieren imaging of rocket exhaust plumes

    NASA Astrophysics Data System (ADS)

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

    2016-11-01

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

  6. Spectral radiance measurements of exhaust plumes from scale model rocket engines.

    PubMed

    McCaa, D J

    1968-05-01

    A short duration experimental technique for investigating radiative properties of rocket exhaust plumes at high altitudes is described. Experimental measurements of the spectral radiance of two interacting exhaust plumes generated by 1/45 scale F-l engines burning gaseous ethylene and oxygen are presented. In addition, the spectral radiance characteristics of several Saturn auxiliary solid propellant rocket motors have also been measured and these results are included. The measurements were obtained with a rapid scanning ir spectrometer.

  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. Rocket engine exhaust plume diagnostics and health monitoring/management during ground testing

    NASA Technical Reports Server (NTRS)

    Chenevert, D. J.; Meeks, G. R.; Woods, E. G.; Huseonica, H. F.

    1992-01-01

    The current status of a rocket exhaust plume diagnostics program sponsored by NASA is reviewed. The near-term objective of the program is to enhance test operation efficiency and to provide for safe cutoff of rocket engines prior to incipient failure, thereby avoiding the destruction of the engine and the test complex and preventing delays in the national space program. NASA programs that will benefit from the nonintrusive remote sensed rocket plume diagnostics and related vehicle health management and nonintrusive measurement program are Space Shuttle Main Engine, National Launch System, National Aero-Space Plane, Space Exploration Initiative, Advanced Solid Rocket Motor, and Space Station Freedom. The role of emission spectrometry and other types of remote sensing in rocket plume diagnostics is discussed.

  10. Radiation from Rocket Exhaust Plumes. Part 1; Inhomogeneous Radiant Heat Transfer from Saturn Type Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    Huffaker, Robert M.; Carlson, Donald J.

    1966-01-01

    A radiant heat transfer computer program has been developed by R-AERO-A to calculate radiation from inhomogeneous gases prevalent in Saturn-type exhaust plumes. The radiating species considered in this computer program are water vapor, carbon dioxide, carbon monoxide and carbon particles. The infrared spectral absorption characteristics of these species have been determined under NASA contract. Band model parameters have been used to represent the infrared spectral absorption coefficients over 25 cm-I increments. A modified Curtis-Goodson approximation is used in the inhomogeneous heat transfer calculation. This has been shown to give satisfactory results over the temperature and pressure range of interest in Saturn exhaust plumes. Results are shown for the Saturn-type engines for specific flow field assumptions. Some comparison with experimental spectroscopic data will also be presented. The effect of wavelength increment, field of view, and distance increment along the line of sight on the heat transfer will be discussed. Computer techniques for minimum computer time in calculating radiation from a three-dimensional flow field will also be outlined.

  11. 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.

  12. 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

  13. 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

  14. 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.

  15. 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.

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

    NASA Astrophysics Data System (ADS)

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

    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.

  17. 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.

  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. 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

  20. A Computer Program for the Prediction of Radiation from Rocket Exhaust Plumes. Volume 1

    DTIC Science & Technology

    1973-05-31

    axisymmetricor three-dimen- sional cylindrical coordinate system . The program can integrate the ra- diation from the gas to a point to providE overall...used to describe the problem consists of two coordinate systems as illustrated in Figure 1. The gas properties are defined at locations in a cylindrical...coordinate system relative to the left-hand X, Y, Z coordinate system . For axisymmetric exhaust plumes, the Z-axis is the center of the cylindrical

  1. Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method

    NASA Technical Reports Server (NTRS)

    Hueser, J. E.; Brock, F. J.; Melfi, L. T., Jr.; Bird, G. A.

    1984-01-01

    A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).

  2. Rocket plume burn hazard.

    PubMed

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

    1980-05-01

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

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Gasdynamic propagation of rocket exhaust products in the upper atmosphere

    NASA Astrophysics Data System (ADS)

    Molchanov, A. G.; Platov, Yu. V.

    2011-12-01

    The dispersion of exhaust products of rocket fuel in the direction perpendicular to the motion of a rocket is investigated in this work. A comparison of the results of numerical calculations with a self-similar approximation of a strong cylindrically symmetric explosion is fulfilled. It is shown that at sufficiently high rocket velocity V ∞, which exceeds the sum of gas exhaust velocity V e from the nozzle and sound speed V s ( V ∞ > V e + V s ), a gasdynamic hole can arise around the rocket trajectory in the upper atmosphere, inside which the total concentration of gas becomes less than the equilibrium concentration of gas at a given altitude. The dynamics of the profiles of density and temperature of the exhaust products inside a rocket plume is calculated.

  9. 3-D adaptive grid Navier-Stokes rocket plume calculations

    NASA Astrophysics Data System (ADS)

    Holcomb, J. Eric

    1991-01-01

    Three-dimensional adaptive-grid full Navier-Stokes calculations performed for the base region and plume of the Minuteman first stage and a simplified version of the Titan first stage are used to demonstrate the applicability of the Navier-Stokes flow solver, EAGLE adaptive grid generator, and k-epsilon turbulence model to rocket plume flowfields. The calculations include realistic exhaust gas thermodynamic properties, with frozen chemistry.

  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. Effect of Soot Particles on Supersonic Rocket Plume Properties

    NASA Astrophysics Data System (ADS)

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

    2012-06-01

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

  12. 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.

  13. Radiation from advanced solid rocket motor plumes

    NASA Astrophysics Data System (ADS)

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

    1994-12-01

    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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. An exact solution of a simplified two-phase plume model. [for solid propellant rocket

    NASA Technical Reports Server (NTRS)

    Wang, S.-Y.; Roberts, B. B.

    1974-01-01

    An exact solution of a simplified two-phase, gas-particle, rocket exhaust plume model is presented. It may be used to make the upper-bound estimation of the heat flux and pressure loads due to particle impingement on the objects existing in the rocket exhaust plume. By including the correction factors to be determined experimentally, the present technique will provide realistic data concerning the heat and aerodynamic loads on these objects for design purposes. Excellent agreement in trend between the best available computer solution and the present exact solution is shown.

  20. 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.

  1. Equations for calculating orbiter surface erosion and breakage rates in IUS and SSUS SRM exhaust plumes

    NASA Technical Reports Server (NTRS)

    Wilson, S. W.

    1978-01-01

    Equations and coefficients for calculating the flux of solid particles in the exhaust plumes of the interim upper stage and SSUS solid rocket motors (SRM) are considered. Modifications required to account for the independent motions of the orbiter and the SRM, such as will result during an on-orbit SRM firing are described.

  2. 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.

  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. Spectrally resolved Rayleigh scattering diagnostic for hydrogen-oxygen rocket plume studies

    NASA Technical Reports Server (NTRS)

    Seasholtz, R. G.; Zupanc, F. J.; Schneider, S. J.

    1991-01-01

    A Rayleigh scattering diagnostic has been developed to measure gas density, temperature, and velocity in the exhaust plume of 100 N thrust class hydrogen-oxygen rockets. The spectrum of argon-ion laser light scattered by the gas molecules in the plume (predominantly water vapor) is measured with a scanning Fabry-Perot interferometer. The gas density is determined from the total scattered power, the gas temperature from the spectral width, and the velocity from the shift in the peak of the spectrum from the frequency of the incident laser light. The diagnostic has been demonstrated in a rocket test cell and a discussion of results is given.

  5. Fabry-Perot interferometer development for rocket engine plume spectroscopy

    NASA Technical Reports Server (NTRS)

    Bickford, R. L.; Madzsar, G.

    1990-01-01

    This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

  6. Non-equilibrium radiation from viscous chemically reacting two-phase exhaust plumes

    NASA Technical Reports Server (NTRS)

    Penny, M. M.; Smith, S. D.; Mikatarian, R. R.; Ring, L. R.; Anderson, P. G.

    1976-01-01

    A knowledge of the structure of the rocket exhaust plumes is necessary to solve problems involving plume signatures, base heating, plume/surface interactions, etc. An algorithm is presented which treats the viscous flow of multiphase chemically reacting fluids in a two-dimensional or axisymmetric supersonic flow field. The gas-particle flow solution is fully coupled with the chemical kinetics calculated using an implicit scheme to calculate chemical production rates. Viscous effects include chemical species diffusion with the viscosity coefficient calculated using a two-equation turbulent kinetic energy model.

  7. 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.

  8. 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.

  9. 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

  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. 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.

  12. Development of a Novel Free Molecule Rocket Plume Model

    DTIC Science & Technology

    2007-11-02

    Development of a Novel Free Molecule Rocket Plume Model Michael Woronowicz Swales Aerospace 5050 Powder Mill Road Beltsville, Maryland 20705, USA...Abstract. This paper describes development of a set of analytic point source transient free molecule equations generated to model behavior ranging from...molecular effusion to rocket plumes. A brief review of model performance for step function mass expulsion will be followed by presentation of physical

  13. 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.

  14. 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.

  15. 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.

  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. 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

  18. Production of Nitrous Oxide in a Rocket Motor Exhaust.

    DTIC Science & Technology

    1980-08-01

    Standard Atmosphere, 1962. NASA , USAF and United States Weather Bureau, 1962. 2 Jensen, D.E. Prediction of Rocket Exhaust Flame Properties. Wilson, A.S...are made of concentrations of N20 produced within the exhaust of a double-base propellant rocet motor. Typical concentrations produced are predicted to

  19. 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.

  20. Rocket plume flowfield characterization using laser Rayleigh scattering

    NASA Astrophysics Data System (ADS)

    Zupanc, Frank J.; Weiss, Jonathan M.

    1992-07-01

    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. 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.

  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. Improvement of Rocket Engine Plume Analysis Techniques

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1982-01-01

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

  4. Monopropellant Thruster Exhaust Plume Contamination Measurements

    DTIC Science & Technology

    1977-09-01

    previously discussed. The optical surfaces of infrared sensors , for example, are particularly susceptiblo to performance degradation through in- creased...gases from a monopropellant rocket engine will not, in most cases, be trapped by the higher temperature surfaces of the spacecraft, the sensor may...K NH3 8.37 x 102 24.2 101°K H20 (b) 0.71 1590K N2H4 40 x 10-6 0.012 165 0 K (Ref. 15) Anilineh)- 0.55 1900K Percent Amonia Dissociation(c) - 65.5

  5. 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

  6. 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.

  7. Effects of Rocket Exhaust on Lunar Soil Reflectance Properties

    NASA Astrophysics Data System (ADS)

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

    2012-12-01

    The Apollo, Surveyor, and Luna spacecraft descent engine plumes affected the regolith at and surrounding their landing sites. Owing to the lack of rapid weathering processes on the Moon, surface alterations are still visible as photometric anomalies in Lunar Reconnaissance Orbiter Camera (LROC) Narrow Angle Camera (NAC) images. These areas are interpreted as disturbance of the regolith by rocket exhaust during descent of the spacecraft, which we refer to as "blast zones" (BZs). The BZs consist of an area of lower reflectance (LR-BZ) compared to the surroundings that extends up to a few meters out from the landers, as well as a broader halo of higher reflectance (HR-BZ) that extends tens to hundreds of meters out from the landers. We use phase-ratio images for each landing site to determine the spatial extent of the disturbed regions and to quantify differences in reflectance and backscattering characteristics within the BZs compared to nearby undisturbed regolith. We also compare the reflectance changes and BZ dimensions at the Apollo sites with those at Luna and Surveyor sites. We seek to determine the effects of rocket exhaust in terms of erosion and particle redistribution, as well as the cause(s) of the reflectance variations, i.e., physical changes at the regolith surface. When approximated as an ellipse, the average Apollo BZ area is ~29,000 m2 (~175 ± 60 m by 200 ± 27 m) which is 10x larger than the average Luna BZ, and over 100x larger than the average Surveyor BZ. Moreover, BZ area scales roughly with lander mass (as a proxy for thrust). The LR-BZs are evident at the Apollo sites, especially where astronaut bioturbation has roughened the soil, leading to a 2-14% reduction in reflectance at ~30° phase. The LR-BZs at the Luna and Surveyor sites are less evident and may be mostly confined to the area below the landers. The average normalized reflectance in the HR-BZs for images with a 30° phase angle is 2-16% higher than in the undisturbed surrounding

  8. Multiple-wavelength transmission measurements in rocket motor plumes

    NASA Astrophysics Data System (ADS)

    Kim, Hong-On

    1991-09-01

    Multiple-wavelength light transmission measurements were used to measure the mean particle size (d(sub 32)), index of refraction (m), and standard deviation of the small particles in the edge of the plume of a small solid propellant rocket motor. The results have shown that the multiple-wavelength light transmission measurement technique can be used to obtain these variables. The technique was shown to be more sensitive to changes in d(sub 32) and standard deviation (sigma) than to m. A GAP/AP/4.7 percent aluminum propellant burned at 25 atm produced particles with d32 = 0.150 +/- 0.006 microns, standard deviation = 1.50 +/- 0.04 and m = 1.63 +/- 0.13. The good correlation of the data indicated that only submicron particles were present in the edge of the plume. In today's budget conscious industry, the solid propellant rocket motor is an ideal propulsion system due to its low cost and simplicity. The major obstacle for solid rocket motors, however, is their limited specific impulse compared to airbreathing motors. One way to help overcome this limitation is to utilize metal fuel additives. Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Aluminum propellants also increase propellant densities and suppress transverse modes of combustion oscillations by damping the oscillations with the aluminum agglomerates in the combustion chamber.

  9. 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

  10. Summary of nozzle-exhaust plume flowfield analyses related to space shuttle applications

    NASA Technical Reports Server (NTRS)

    Penny, M. M.

    1975-01-01

    Exhaust plume shape simulation is studied, with the major effort directed toward computer program development and analytical support of various plume related problems associated with the space shuttle. Program development centered on (1) two-phase nozzle-exhaust plume flows, (2) plume impingement, and (3) support of exhaust plume simulation studies. Several studies were also conducted to provide full-scale data for defining exhaust plume simulation criteria. Model nozzles used in launch vehicle test were analyzed and compared to experimental calibration data.

  11. 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.

  12. 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.

  13. Numerical Investigation of Twin-Nozzle Rocket Plume Phenomenology

    DTIC Science & Technology

    1997-01-01

    calculations indicate that the viscous stress model, kinetic chemistry , and nozzle exit profile are significant parameters that should be considered in analyses and interpretation of the calculations....been completed to assess the effects of three-dimensionality, turbulent viscosity, afterburning, near-field shock structure, finite-rate kinetic ... chemistry , intranozzle geometric spacing, and initial nozzle exit plane profile effects on the subsequent plume exhaust flow field. The results of these

  14. Prediction of rocket plume radiative heating using backward Monte-Carlo method

    NASA Technical Reports Server (NTRS)

    Wang, K. C.

    1993-01-01

    A backward Monte-Carlo plume radiation code has been developed to predict rocket plume radiative heating to the rocket base region. This paper provides a description of this code and provides sample results. The code was used to predict radiative heating to various locations during test firings of 48-inch solid rocket motors at NASA Marshall Space Flight Center. Comparisons with test measurements are provided. Predictions of full scale sea level Redesigned Solid Rocket Motor (RSRM) and Advanced Solid Rocket Motor (ASRM) plume radiative heating to the Space Shuttle external tank (ET) dome center were also made. A comparison with the Development Flight Instrumentation (DFI) measurements is also provided.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Plume spectrometry for liquid rocket engine health monitoring

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

  20. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Experimental Evaluation of Rocket Exhaust Diffusers for Altitude Simulation

    NASA Technical Reports Server (NTRS)

    Sivo, Joseph N.; Meyer, Carl L.; Peters, Daniel J.

    1960-01-01

    An experimental investigation of exhaust diffusers has been conducted to evaluate various methods of minimizing the overall pressure ratio (from chamber to ambient pressure) required to establish and maintain full expansion of the nozzle flow (altitude simulation). Exhaust-diffuser configurations investigated were (1) cylindrical diffusers, (2) diffusers with contraction, and (3) diffusers including a right-angle turn. Cylindrical diffusers were evaluated with primary nozzles of various area ratios and types, as well as two clustered configurations; the other diffusers were evaluated with individual nozzles of constant area ratio and varied type. Air was the working fluid, except for two check points obtained with JP-4 fuel and liquid-oxygen rocket engines and cylindrical diffusers. The minimum length-diameter ratio of cylindrical diffusers was about 6 for minimum pressure-ratio requirements. With cylindrical diffusers of adequate length, the pressure-ratio requirements were primarily a function of the ratio of diffuser to nozzle-throat areas and were essentially independent of primary-nozzle type (including two clustered configurations) or area ratio. The two check points obtained with rocket engines indicated the pressure-ratio requirements at given ratios of diffuser to nozzle-throat areas were lowered, as compared with the requirements with air, as a result of the reduced ratio of specific heats. The minimum length-diameter ratio of the contraction throat of convergent-divergent diffusers was also about 6 for minimum pressure-ratio requirements. With adequate contraction-throat length, the pressure-ratio requirements of such diffusers were appreciably below those of comparable cylindrical diffusers when used with conical and cutoff-isentropic nozzles, but not when used with a bell nozzle. Minimum pressure-ratio requirements of a diffuser including a simple long-radius right-angle turn at maximum diffuser area, obtained with the center of radius of the turn a

  20. An Investigation of Instantaneous Plume Rise from Rocket Exhaust

    DTIC Science & Technology

    1996-12-01

    equations will be used in the following sections to derive expressions for stability and the cloud stabilization height involved with hot, buoyant... cloud stabilization height by solving the conservation of momentum equation in the form given in equation 2-19: d2 ()W3) 3 dt2 - (2-29) This is a second...overall trend of lowering the value of y was an increase in the cloud stabilization height when comparing results from the same model. The trend held

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  2. Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation

    NASA Technical Reports Server (NTRS)

    Sivo, Joseph N.; Peters, Daniel J.

    1959-01-01

    A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.

  3. 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.

  4. 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.

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

    NASA Technical Reports Server (NTRS)

    Everson, John; Nelson, H. F.

    1993-01-01

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

  6. 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.

  7. Effects of rocket exhaust on lunar soil reflectance properties

    NASA Astrophysics Data System (ADS)

    Clegg, Ryan N.; Jolliff, Bradley L.; Robinson, Mark S.; Hapke, Bruce W.; Plescia, Jeffrey B.

    2014-01-01

    High-resolution images of the Surveyor, Luna, and Apollo landing sites obtained by the Lunar Reconnaissance Orbiter Camera (LROC) Narrow Angle Camera (NAC) show regions around the landers where reflectivity of the surface was modified. We interpret the change in reflectance properties of these regions mainly as disturbance of the regolith by rocket exhaust during descent of the spacecraft and we refer to these areas herein as "blast zones" (BZs). The BZs consist of an area of lower reflectance (LR-BZ) compared to the surroundings that extends up to a few meters out from the landers, as well as a broader halo of higher reflectance (HR-BZ) that extends tens to hundreds of meters away from the landers. When approximated as an ellipse, the average Apollo BZ area is ˜29,000 m2 (˜175 ± 60 m by 200 ± 27 m) which is 10× larger than the average Luna BZ, and over 100× larger than the average Surveyor BZ. The LR-BZs are most evident at the Apollo sites, especially where astronaut activity disturbed the soil, leading to a 15-30% (relative to background undisturbed areas) reduction in reflectance at ˜30° phase angle. The LR-BZs at the Surveyor and Luna sites are less evident and are unresolvable with NAC images. The average reflectance in the HR-BZs as determined for 30° phase angle is 3-12% higher than in the undisturbed surrounding areas; this magnitude is the same, within uncertainty, for all sites, indicating a common process or combination of processes causing differences in reflectance properties of the regolith. Phase-ratio images and photometric data collected over a range of illumination geometries show that a greater separation in reflectance occurs between the HR-BZs and undisturbed areas at phase angles between 0° and 70° and indicates that the HR-BZs are less backscattering than undisturbed areas. The LR-BZs are affected by macroscopic disruption of the surface and astronaut activity (at the Apollo sites). For the HR-BZ areas, reflectance has likely been

  8. 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.

  9. Model of lidar range-Doppler signatures of solid rocket fuel plumes

    NASA Astrophysics Data System (ADS)

    Bankman, Isaac N.; Giles, John W.; Chan, Stephen C.; Reed, Robert A.

    2004-09-01

    The analysis of particles produced by solid rocket motor fuels relates to two types of studies: the effect of these particles on the Earth's ozone layer, and the dynamic flight behavior of solid fuel boosters used by the NASA Space Shuttle. Since laser backscatter depends on the particle size and concentration, a lidar system can be used to analyze the particle distributions inside a solid rocket plume in flight. We present an analytical model that simulates the lidar returns from solid rocket plumes including effects of beam profile, spot size, polarization and sensing geometry. The backscatter and extinction coefficients of alumina particles are computed with the T-matrix method that can address non-spherical particles. The outputs of the model include time-resolved return pulses and range-Doppler signatures. Presented examples illustrate the effects of sensing geometry.

  10. 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.

  11. 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.

  12. 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

  13. 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.

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

    NASA Technical Reports Server (NTRS)

    Susko, M.

    1979-01-01

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

  15. 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.

  16. Shuttle primary reaction control system engine exhaust plume contamination effects

    NASA Technical Reports Server (NTRS)

    Koontz, Steve; Ehlers, Horst; Pedley, Mike; Cross, John; Hakes, Charles

    1993-01-01

    Space Shuttle proximity operations constitute an important part of the SSF induced external environment. The impingement of primary reaction control system (PRCS) engine plumes on SSF functional surfaces during docking or berthing and separation leads to concerns about molecular contamination and high speed particle impact. The Shuttle Plume Impingement flight Experiment (SPIE) was designed to provide a direct measure of both the molecular contamination and particle impact rates produced by Shuttle PRCS engines in the LEO environment. The measured permanent deposition produced by PRCS engine firings was less than that assumed in current SSF programatic assessments. Only two to three possible high velocity particle impact pits were observed on the RMS end effector hardware.

  17. 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.

  18. 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.

  19. 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,

  20. 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.

  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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

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

    SciTech Connect

    Yu, Shengtao.

    1989-01-01

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

  10. 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

  11. 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.

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

    NASA Astrophysics Data System (ADS)

    Vincent, R. Anthony; Hawks, Michael R.

    2011-05-01

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

  13. Computational models for the viscous/inviscid analysis of jet aircraft exhaust plumes

    NASA Astrophysics Data System (ADS)

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

    1980-05-01

    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.

  14. 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.

  15. 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.

  16. 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

  17. 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

    2013-03-05

    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.

  18. 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

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. On-road measurement of particle emission in the exhaust plume of a diesel passenger car.

    PubMed

    Vogt, Rainer; Scheer, Volker; Casati, Roberto; Benter, Thorsten

    2003-09-15

    Particle size distributions were measured under real world dilution conditions in the exhaust plume of a diesel passenger car closely followed by a mobile laboratory on a high speed test track. Under carefully controlled conditions the exhaust plume was continuously sampled and analyzed inside the mobile laboratory. Exhaust particle size distribution data were recorded together with exhaust gas concentrations, i.e., CO, CO2, and NO(x), and compared to data obtained from the same vehicle tested on a chassis dynamometer. Good agreement was found for the soot mode particles which occurred at a geometric mean diameter of approximately 50 nm and a total particle emission rate of 10(14) particles km(-1). Using 350 ppm high sulfur fuel and the standard oxidation catalyst a bimodal size distribution with a nucleation mode at 10 nm was observed at car velocities of 100 km h(-1) and 120 km h(-1), respectively. Nucleation mode particles were only present if high sulfur fuel was used with the oxidation catalyst installed. This is in agreement with prior work that these particles are of semivolatile nature and originate from the nucleation of sulfates formed inside the catalyst. Temporal effects of the occurrence of nucleation mode particles during steady-state cruising and the dynamical behavior during acceleration and deceleration were investigated.

  6. 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.

  7. Laser-Induced Fluorescence Measurements of OMS and LR91 Amine Rocket Exhausts

    DTIC Science & Technology

    1997-06-01

    interpretation of the LIF signal to yield radial and temporal profiles of radial flow velocity and static temperature. CFD modeling of the plumes was used to...provide baseline estimates of the exhaust flow properties. It was found that the CFD model does a fairly good job of predicting the gross behavior...collection and interpretation of the LIF signal to yield radial and temporal profiles of radial flow velocity and static temperature. CFD modeling of the

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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

  13. 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.

  14. Numerical and experimental capabilities for studying rocket plume-regolith interactions

    NASA Astrophysics Data System (ADS)

    White, C.; Scanlon, T. J.; Merrifield, J. A.; Kontis, K.; Langener, T.; Alves, J.

    2016-11-01

    Soft landings on extra-terrestrial airless bodies will be required for future sample return missions, such as the Phobos Sample Return (PhSR). PhSR is a candidate mission of ESA's Mars Robotic Exploration Preparation (MREP-2) Programme. Its main objective is to acquire and return a sample from the Martian moon Phobos, after a scientific characterisation phase of the moon and of the landing site. If a rocket is used to slow down the spacecraft to a vertical descent velocity that it will be able to free-fall from, care has to be taken to ensure that the rocket exhaust does not contaminate the surface regolith that is to be collected, and that the rocket does not cause unacceptable levels of erosion to the surface, which could jeopardise the mission. In addition to the work being done in the scope of PhSR, the European Space Agency is funding an experimental facility for investigating these nozzle expansion problems; the current progress of this is described. To support this work, an uncoupled hybrid computational fluid dynamics-direct simulation Monte Carlo method is developed and used to simulate the exhaust of a mono-propellant rocket above the surface of an airless body. The pressure, shear stress, and heat flux at the surface are compared to an analytical free-molecul solution to determine the altitude above which the free-molecular solution is suffcient for predicting these properties. The pressures match well as low as 15 m above the surface, but the heat flux and shear stress are not in agreement until an altitude of 40 m. A new adsorption/desorption boundary condition for the direct simulation Monte Carlo code has also been developed for future use in in-depth contamination studies.

  15. 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.

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

    NASA Astrophysics Data System (ADS)

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

    Sulfuric acid (SA) was for the first time directly detected in the exhaust plume of a jet aircraft in flight. The measurements were made by a novel aircraft-based VACA (Volatile Aerosol Component Analyzer) instrument of MPI-K Heidelberg while the research aircraft Falcon was chasing another research aircraft ATTAS. The VACA measures the total SA in the gas and in volatile submicron aerosol particles. During the chase the engines of the ATTAS alternatively burned sulfur-poor and sulfur-rich fuel. In the sulfur-rich plume very marked enhancements of total SA were observed of up to 1300 pptv which were closely correlated with ΔCO2 and ΔT and were far above the local ambient atmospheric background-level of typically 15-50 pptv. Our observations indicate a lower limit for the efficiency ɛ for fuel-sulfur conversion to SA of 0.34 %.

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

    NASA Technical Reports Server (NTRS)

    Ryason, P. R.

    1975-01-01

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

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Optical Emissions from the High Speed Rocket Exhaust Interaction with the Ionosphere

    NASA Astrophysics Data System (ADS)

    Bhatt, A.; Bernhardt, P. A.; Erickson, P. J.; Lind, F. D.; Varney, R. H.; Kelley, M. C.

    2009-12-01

    An increasing number of space shuttle and rocket launches have inspired an investigation into the effects of vehicle exhaust on the earth's upper atmosphere. A controlled Charged Aerosol Release Experiment (CARE) will be carried out in September 2009 from Wallops Island, Virginia. The high speed exhaust from the Nihka motor on the rocket contains primarily Al2O3, H2O, CO, HCl, N2, CO2 and H2, would have an exit velocity of 2.8 km/s, and the exhaust would last for 18 seconds. The heavy particles are expected to form a charged dust layer in the lower thermosphere and the ionospheric E region. Sunlight scattered from the particulates will produce a bright optical display that can be observed with a ground spectrograph. In addition, the interactions of these molecular ions with the background ionospheric electrons are expected to also produce optical emissions in the range of 400-1000 nm. Observations from the CARE campaign were obtained using a ground based Visible/NIR spectrograph and the Millstone Hill ionospheric incoherent scatter radar located in Massachusetts. The chemical processes behind the expected emissions are compared with the observed optical emissions for both temporal and spatial scales. The spectral emissions observed with the spectrograph from the Wallops site are also compared to those observed at other ground based optical diagnostics sites. The temporal evolution of the emissions is correlated with that of the formation of the ionospheric layers as observed with the Millstone Hill radar. Finally, the significance of these results and future plans for more and enhanced observation techniques will be discussed.

  5. Solid rocket motor plume particle size measurements using multiple optical techniques in a probe

    NASA Astrophysics Data System (ADS)

    Manser, John R.

    1995-03-01

    An experimental investigation to measure particle size distributions in the plume of sub-scale solid rocket motors was conducted. A phase-Doppler particle analyzer (pDPA) in conjunction with three-wavelength extinction measurements were used in a specially designed particle collection probe in an attempt to determine the entire plume particle size distribution. In addition, a laser ensemble particle sizer was used for comparative data. The PDPA and Malvem distributions agreed in the observed modes near 1 and 4.5 micron diameter (d). Scanning electron microscope (SEM) pictures of collected particles were in good agreement with the measured Malvem Sauter mean diameter (d(sub 32)) of 2.59 micron. Data analysis indicates that less than 3% of the total mass of the particles was contained in particles with diameter d dess than 0.5 micron. Therefore, the PDPA, which can typically measure particles down to a minimum diameter of 0.5 micron with a dynamic range (d(sub max):d(sub min)) of 50:1, can be used by itself to determine the particle size distribution. Multiple wavelength measurements were found to be very sensitive to inaccuracies in the measured transmittances.

  6. 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.

  7. 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.

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

    NASA Astrophysics Data System (ADS)

    Wang, Yan Jason; Zhang, K. Max

    2012-11-01

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

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Verification of large-scale rapid transport in the lower thermosphere: Tracking the exhaust plume of STS-107 from launch to the Antarctic

    NASA Astrophysics Data System (ADS)

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

    2011-05-01

    New analysis of the Doppler shift of O2 airglow spectra recorded by the TIMED Doppler Interferometer (TIDI) and the High Resolution Doppler Imager (HRDI) have provided conclusive evidence that the shuttle main engine exhaust plume generated in the lower thermosphere by the launch of STS-107 and imaged by the Global Ultraviolet Imager (GUVI) instrument on TIMED was transported to the Antarctic in ˜80 h, supporting a key inference from the initial study by Stevens et al. (2005). These new results were aided by improved knowledge of the effects of instrumental and satellite artifacts imposed on the Doppler spectra. STS-107 launched on 16 January 2003, and the neutral wind near its launch trajectory and nearby volume was sampled within minutes by TIDI. These initial observations suggested that the northernmost end of the shuttle's exhaust plume would move northeast and that the southern end would move southeast, motions that were identified in imagery acquired during the next orbit of TIMED. The direction and magnitude of plume motion inferred from GUVI images obtained 12, 26, and 50 h after launch were again confirmed by TIDI and HRDI. The appearance of the plume over the Antarctic ˜80 h after launch, inferred from earlier work by the appearance of iron ablated from the shuttle's main engines, was consistent with neutral winds measured by the satellite Doppler instruments over the Antarctic. The transport of the plume from the coast of Florida to the Antarctic was aided by the favorable phase and strong amplitude of a 2 day planetary wave of wave number three in the southern hemisphere on 18 January 2003. The existence of the 2 day wave was deduced from zonally averaged and combined TIDI and HRDI neutral wind observations. We conclude that the existence of strong and sustained winds in the MLT, significantly greater than expected from empirical and theoretical models, is indisputable and provides compelling evidence supporting the global-scale nature of

  14. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1970-01-01

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

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

    SciTech Connect

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

    1980-10-01

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

  17. 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. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2011-02-01

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

  1. Ionospheric effects of rocket exhaust products: Skylab and HEAO-C

    SciTech Connect

    Zinn, J.; Sutherland, C.D.; Duncan, L.M.; Stone, S.N.

    1981-01-01

    This paper is about ionospheric F-layer depletions produced by chemical reactions with exhaust gases from large rockets. It describes a 2-dimensional computer model of the ionosphere, and it compares model results with experimental data on the structure and variability of the natural ionosphere, as well as data on ionospheric holes produced by the launches of Skylab (May, 1973) and HEAO-C (September, 1979). It also describes measurements made in conjunction with the HEAO-C launch. The computer model includes an approximate representation of thermospheric tidal winds and E fields in addition to vertical motions associated with diurnal changes in temperature. The computed ionospheric structure is sensitive to all the above. For a small number of cases, results are compared of computations of the normal diurnal variations of ionospheric structure with incoherent scatter and total electron content data. Computations of ionospheric depletions from the Skylab and HEAO-C launches are in satisfactory agreement with the observations. The winds appear to be essential for interpretation of the Skylab results.

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

    NASA Technical Reports Server (NTRS)

    Susko, M.

    1976-01-01

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

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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).

  8. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2002-08-01

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

  10. 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.

  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. Development of Terahertz Rayleigh Scattering Diagnostics for a Solid Rocket Exhaust Plume

    DTIC Science & Technology

    2010-10-28

    Approved for public release. Distribution unlimited. The views expressed in this article are those of the author and do not...the atmospheric properties of other large bodies such as asteroids and comets. These characteristics can lead to improvements in the models of

  13. 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.

  14. Passive Ranging of Dynamic Rocket Plumes Using Infrared and Visible Oxygen Attenuation

    DTIC Science & Technology

    2011-03-01

    405 m range and a full -scale solid rocket motor at 900 m range. The NIR O2 band provided range estimates accurate to within 3 % for both tests, while...and zenith angles. Data sets for passive ranging were taken of both a halogen lamp and a full scale stationary solid rocket motor at various ranges up...combustion was targeted as the emission source. Two static tests were conducted with a full scale solid rocket motor and a shoulder-launched model SAM

  15. 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.

  16. Plume primary smoke

    NASA Astrophysics Data System (ADS)

    Chastenet, J. C.

    1993-06-01

    The exhaust from a solid propellant rocket motor usually contains condensed species. These particles, also called 'Primary Smoke', are often prejudicial to missile detectability and to the guidance system. To avoid operational problems it is necessary to know and quantify the effects of particles on all aspects of missile deployment. A brief description of the origin of the primary smoke is given. It continues with details of the interaction between particles and light as function of both particles and light properties (nature, size, wavelength, etc). The effects of particles on plume visibility, attenuation of an optical beam propagated through the plume and the contribution of particles on optical signatures of the plume are also described. Finally, various methods used in NATO countries to quantify the primary smoke effects are discussed.

  17. 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.

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

    NASA Technical Reports Server (NTRS)

    Hilst, G. R.

    1974-01-01

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

  19. 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.

  20. Rocket noise - A review

    NASA Astrophysics Data System (ADS)

    McInerny, S. A.

    1990-10-01

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

  1. Overestimation of on-road air quality surveying data measured with a mobile laboratory caused by exhaust plumes of a vehicle ahead in dense traffic areas.

    PubMed

    Woo, Sang-Hee; Kwak, Kyung-Hwan; Bae, Gwi-Nam; Kim, Kyung Hwan; Kim, Chang Hyeok; Yook, Se-Jin; Jeon, Sangzin; Kwon, Sangil; Kim, Jeongsoo; Lee, Seung-Bok

    2016-11-01

    The unintended influence of exhaust plumes emitted from a vehicle ahead to on-road air quality surveying data measured with a mobile laboratory (ML) at 20-40 km h(-1) in dense traffic areas was investigated by experiment and life-sized computational fluidic dynamics (CFD) simulation. The ML equipped with variable sampling inlets of five columns by four rows was used to measure the spatial distribution of CO2 and NOx concentrations when following 5-20 m behind a sport utility vehicle (SUV) as an emitter vehicle equipped with a portable emission monitoring system (PEMS). The PEMS measured exhaust gases at the tailpipe for input data of the CFD simulations. After the CFD method was verified with experimental results of the SUV, dispersion of exhaust plumes emitted from a bus and a sedan was numerically analyzed. More dilution of the exhaust plume was observed at higher vehicle speeds, probably because of eddy diffusion that was proportional to turbulent kinetic energy and vehicle speed. The CO2 and NOx concentrations behind the emitter vehicle showed less overestimation as both the distance between the two vehicles and their background concentrations increased. If the height of the ML inlet is lower than 2 m and the ML travels within 20 m behind a SUV and a sedan ahead at 20 km h(-1), the overestimation should be considered by as much as 200 ppb in NOx and 80 ppm in CO2. Following a bus should be avoided if possible, because effect of exhaust plumes from a bus ahead could not be negligible even when the distance between the bus and the ML with the inlet height of 2 m, was more than 40 m. Recommendations are provided to avoid the unintended influence of exhaust plumes from vehicles ahead of the ML during on-road measurement in urban dense traffic conditions.

  2. Measurements of Cl, ClO, and CO2 in the Exhaust Plume of the Space Shuttle

    NASA Astrophysics Data System (ADS)

    Toohey, D. W.; Thornton, B. F.; Avallone, L. M.; Ross, M. N.; Richard, E.; Kelly, K.

    2001-05-01

    Measurements of chlorine atoms (Cl), chlorine oxide (ClO), and carbon dioxide (CO2) were carried out in the exhaust plume of the Space Shuttle on September 8, 2000 (STS-106). Cl and ClO were detected at abundances exceeding 20 parts per billion (ppb) and 50 ppb, respectively at a sampling rate of 25 Hz by the technique of resonance fluorescence on the NASA WB-57 aircraft. These results provide a horizontal resolution of about 10 meters, and extreme variability is observed over these scales for over an hour. CO2 was measured by non-dispersive infrared with a sampling frequency of 1 second, with enhancements of tens of parts per million (ppm) on most passes. We examine the evolution of the ratio of Cl to ClO, referenced to CO2 as a conservative tracer of mixing, examine the correlations with ozone and methane, and discuss the implications for ozone loss.

  3. Optical properties of mercury ion thruster exhaust plumes Significance for candidate SEP science instruments. [Solar Electric Propulsion

    NASA Technical Reports Server (NTRS)

    Goldstein, R.; Monahan, K. M.

    1975-01-01

    Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume emission rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter was considered incompatible with simultaneous thruster operation.

  4. Characterizing and Overcoming Spectral Artifacts in Imaging Fourier-Transform Spectroscopy of Turbulent Exhaust Plumes

    DTIC Science & Technology

    2009-10-01

    refer to Figure 1). Within the Spectro Mode, the non-uniformity corrected broadband IR image is displayed in real-time. Moreover, the raw spectrum or...simplifications to help understand the problem, the first being that the plume is in local thermodynamic equilibrium . Inelastic collision rates are...Pi can be written as (2) where the gj denotes the degeneracy of the upper state, h is Planck’s constant , c is the speed of light, and kB is

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

    NASA Technical Reports Server (NTRS)

    Castner, Raynold S.

    2010-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Castner, Raymond S.

    2009-01-01

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

  7. 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.

  8. 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.

  9. Characterizing the Exhaust Plume of the Three-Electrode Micro Pulsed Plasma Thrusters

    DTIC Science & Technology

    2009-03-01

    Three-electrode micro -PPTs are a new version of two electrode micro -PPTs devices. Performance predictions and contamination expectations are...and velocity of the exhaust particles is the emphasis of this work. The results show that the three-electrode micro -PPT is more reliable than two...electrode micro -PPT and the operating energy range from 2 to 4.5 Joule is proper value to operate it. This research also shows that the angle

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

    NASA Technical Reports Server (NTRS)

    Vonglahn, U. H.; Goodykoontz, J. H.; Wasserbauer, C.

    1984-01-01

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

  11. Model of Chlorocarbon (CFC-12) Chemisorption on Solid Rocket Motor Alumina Exhaust Particles.

    DTIC Science & Technology

    1995-12-01

    propellant for solid rocket motors currently used in the United States is a mixture commonly known as ammonium per chlorate, or AP. The three main...chamber, the sodium ions 23 scavenge most of the chloride ions , forming NaCL thus preventing the chloride ions from forming HC1. Laboratory...an accurate quantification of the acid reducing capability of the scavenged propellants during an open ah test has not yet been made.33 The sodium

  12. Multiple dopant injection system for small rocket engines

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

    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.

  13. 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.

  14. 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

  15. 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.

  16. 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.

  17. Development of a Fabry-Perot interferometer for rocket engine plume monitoring

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    The development of a lightweight, compact, high-resolution Fabry-Perot interferometer (FPI) based spectrometer capable of detecting the spectral signatures of eroding engine components during test and/or flight operations is discussed. FPI based spectrometers will be designed to be smaller and lighter than grating or prism type devices and to provide greater wavelength resolving capability. The FPI system seeks to combine the features of high line resolution, active background discrimination, smart digital signal processing, and rocket engine flight capability. The design, fabrication, and test of a breadboard FPI spectrometer have been completed and the breadboard instrument has clearly demonstrated the viability of the approach. The breadboard instrument design and test results are also presented.

  18. 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.

  19. 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

  20. Plasma-gas interactions studies in a hybrid plume plasma rocket

    NASA Technical Reports Server (NTRS)

    Chang, F. R.; Krueger, W. A.; Yang, T. F.; Fisher, J. L.

    1985-01-01

    Plasma-gas interaction was investigated and the basic mechanisms for energy and particle transport. The solution approach assumes cylindrical geometry and includes a multiplicity of atomic reactions, and the presence of a strong magnetic field is described. The principal reactions are electron and ion impact ionization, and charge exchange between hot ions and cold neutrals. Radial particle and energy transport is mainly by diffusion. A modified Bohm diffusion model for plasma in the core of the plume, and classical neutral particle diffusion in the cooler regions of the flow are presented. Neutrals are allowed to free stream in the low density regions, where the collision mean-free-path becomes comparable, or even larger than the characteristic dimensions of the system.

  1. Investigation of the effects of solid rocket motor propellant composition on plume signature

    NASA Astrophysics Data System (ADS)

    Snaza, Clay J.

    1994-06-01

    Three propellants with aluminum/silicon weight percentages of 18/0%, 13.5/4.5%, and 12/6% were fired in a subscale motor to determine if the plume infrared signature could be reduced without a significant loss in specific impulse. Spectral measurements from 2.5 to 5.5 micrometers and thermal measurements from 3.5 to 5.0 micrometers were made. Plume particle size measurements showed that only particles with small diameters (less than 1.93 micrometers) were present with any significant volume. Replacing a portion of the aluminum in a highly metallized solid propellant with silicon was found to eliminate the Al2O3 in favor of SiO2 and Al6SiO13, without any change in particulate mass concentration or any large change in particle size distribution. These particulates were found to have significantly lower absorptivity than Al2O3. An additional investigation was conducted to determine the particle size distribution at the nozzle entrance. Malvern ensemble scattering, phase-Doppler single particle scattering, and laser transmittance measurements made through windows in the combustion chamber at the nozzle entrance indicated that large particles were present (to 250 micrometers). However, most of the mass of the particles was contained in particles with diameters smaller than 5 micrometers. Approximate calculations made with the measured data showed that if 100 micrometer particles are present with the smoke (particles with diameters less than 2 micrometers) they could account for only approximately 10% of the article volume.

  2. Radiometric observations of the 752.033-GHz rotational absorption line of H2O from a laboratory jet. [simulation of rocket plumes

    NASA Technical Reports Server (NTRS)

    Dionne, G. F.; Fitzgerald, J. F.; Chang, T.-S.; Fetterman, H. R.; Litvak, M. M.

    1980-01-01

    With the aid of a high-resolution two-stage heterodyne radiometer, spectral absorption measurements of the 752.033 GHz line of water vapor were carried out, using a blackbody continuum as a background radiation source for investigating the absorptive properties of the H2O content of high altitude rocket plumes. To simulate this physical situation in a laboratory environment, a small steam jet was operated within a large high-vacuum chamber, with the H2O jet plume traversing the radiometer line of sight. The experiments verified that this rotational line is optically thick, with excitation temperatures below 100 K, in the downstream part of the plume, as predicted by theoretical modelling.

  3. 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.

  4. Some environmental considerations relating to the interaction of the solid rocket motor exhaust with the atmosphere: Predicted chemical composition of exhaust species and predicted conditions for the formation of HCl aerosol

    NASA Technical Reports Server (NTRS)

    Rhein, R. A.

    1973-01-01

    The exhaust products of a solid rocket motor using as propellant 14% binder, 16% aluminum, and 70% (wt) ammonium perchlorate consist of hydrogen chloride, water, alumina, and other compounds. The equilibrium and some frozen compositions of the chemical species upon interaction with the atmosphere were computed. The conditions under which hydrogen chloride interacts with the water vapor in humid air to form an aerosol containing hydrochloric acid were computed for various weight ratios of air/exhaust products. These computations were also performed for the case of a combined SRM and hydrogen-oxygen rocket engine. Regimes of temperature and relative humidity where this aerosol is expected were identified. Within these regimes, the concentration of HCL in the aerosol and weight fraction of aerosol to gas phase were plotted. Hydrochloric acid aerosol formation was found to be particularly likely in cool humid weather.

  5. 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.

  6. Plume base flow simulation technology

    NASA Technical Reports Server (NTRS)

    Roberts, B. B.; Wallace, R. O.; Sims, J. L.

    1983-01-01

    A combined analytical/empirical approach was studied in an effort to define the plume simulation parameters for base flow. For design purposes, rocket exhaust simulation (i.e., plume simulation) is determined by wind tunnel testing. Cold gas testing was concluded to be a cost and schedule effective data base of substantial scope. The results fell short of the target, although work conducted was conclusive and advanced the state of the art. Comparisons of wind tunnel predictions with Space Transportation System (STS) flight data showed considerable differences. However, a review of the technology program data base has yielded an additional parameter that may correlate flight and cold gas test data. Data from the plume technology program and the NASA test flights are presented to substantiate the proposed simulation parameters.

  7. 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

  8. Exhaust Trail of STS-51-L

    NASA Technical Reports Server (NTRS)

    1986-01-01

    On January 28, 1986, the Space Shuttle Challenger and her seven-member crew were lost when a ruptured O-ring in the right Solid Rocket Booster caused an explosion soon after launch. This photograph, taken a few seconds after the accident, shows the Space Shuttle Main Engines and Solid Rocket Booster exhaust plumes entwined around a ball of gas from the External Tank. Because shuttle launches had become almost routine after fifty successful missions, those watching the shuttle launch in person and on television found the sight of the explosion especially shocking and difficult to believe until NASA confirmed the accident.

  9. Modeling a VASIMR rocket at UVSC

    NASA Astrophysics Data System (ADS)

    Matheson, Phil; Gray, William; Page, Leland

    2004-10-01

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

  10. 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.

  11. Investigations of Rocket Engine Combustion Emissions During ACCENT

    NASA Astrophysics Data System (ADS)

    Ross, M. N.; Friedl, R. R.

    2001-12-01

    The composition of rocket combustion emissions and the atmospheric processes that determine their stratospheric impacts are poorly understood. While present day rocket emissions do not significantly affect stratospheric chemistry, the potential for vigorous growth of the space transportation industry in coming decades suggests that rocket emissions and their stratospheric impacts should be better understood. A variety of in-situ measurements and modeling results were obtained during the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) effort that will be used to evaluate the role of rocket exhaust in perturbing ozone chemistry in plume wakes and in the global stratosphere. We present a review of the ACCENT rocket emissions science objectives, summarize data obtained during the WB-57F plume wake sorties, and briefly discuss how the data will help resolve several outstanding questions regarding the impact of rocket emissions on the stratosphere. These include measurement of the emission indices for several important rocket engine combustion products and validation of plume wake chemistry models.

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

    NASA Astrophysics Data System (ADS)

    Glover, Timothy Ward

    2002-01-01

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

  13. 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.

  14. 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.

  15. 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.

  16. A tandem mirror plasma source for hybrid plume plasma studies

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  17. Analysis of the AlCl Absorption Feature and the Searchlight Emission Effect Observed in Solid-Propellant Rocket Plumes

    DTIC Science & Technology

    1992-09-01

    phase function was approximated by the Henyey - Greenstein function ...searchlight emission are supported by Monte Carlo radiative transfer calculations using the measured plume extinction profile. It is concluded that...Press, New York, 1980. 25. Kamiuto, K. "Study of the Henyey - Greenstein Approximation to Scattering Phase Functions ." Journal of

  18. Analysis of the AICI Absorption Feature and the Searchlight Emission Effect Observed in Solid-Propellant Rocket Plumes

    DTIC Science & Technology

    1992-09-01

    Henyey - Greenstein Approximation to Scattering Phase Functions ." Journal of Quantitative Spectroscopy and Radiative Transfer , Vol. 37, "To. 4... phase function . The ray trace is continued until the photon is either absorbed or escapes from the plume boundary. Monte Carlo calculations for the... Greenstein function (Refs. 25 and 26). This is an analytical approximation to the Mie scattering phase function . It lacks

  19. 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.

  20. Characterization of Rocket Propellant Combustion Products. Chemical Characterization and Computer Modeling of the Exhaust Products from Four Propellant Formulations

    DTIC Science & Technology

    1990-12-31

    Fractions as a Function of Exit/Throat Area Ratios Com position Q ............................................... 43 20 Effect of ± 5% Shift in Heat of...Constituents in ASCF Chamber .................................. 49 23 Effect of Choice Gaseous Equation of State on Computed Mole Fractions for Composition...health hazards from weapons combustion products, to include rockets and missiles, became evident, Research to elucidate significant health effects of

  1. 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.

  2. 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.

  3. 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.

  4. 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…

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

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

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  12. 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.

  13. 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.

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

    NASA Astrophysics Data System (ADS)

    O'Brien, Brian

    2009-05-01

    Dust is the Number 1 environmental hazard on the Moon, yet its movements and adhesive properties are little understood. Matchbox-sized, 270-gram Dust Detector Experiments (DDEs) measured contrasting effects triggered by rocket exhausts of Lunar Modules (LM) after deployment 17 m and 130 m from Apollo 11 and 12 LMs. Apollo 11 Lunar Seismometer was contaminated, overheated and terminated after 21 days operation. Apollo 12 hardware was splashed with collateral lunar dust during deployment. DDE horizontal solar cell was cleansed of nominally 0.3 mg cm-2 dust by 80% promptly at LM ascent and totally within 7 minutes. A vertical cell facing East was half-cleaned promptly then totally over hundreds of hours. Each cell cooled slightly. For the first time lunar electrostatic adhesive forces on smooth silicon were directly measured by comparison with lunar gravity. Analyses imply this adhesive force weakens as solar angle of incidence decreases. If valid, future lunar astronauts may have greater problems with dust adhesion in the middle half of the day than faced by Apollo missions in early morning. A sunproof shed may provide dust-free working environments on the Moon. Low-cost laboratory tests with DDEs and simulated lunar dust can use DDE benchmark lunar data quickly, optimising theoretical modelling and planning of future lunar expeditions, human and robotic.

  15. Reusable rocket engine optical condition monitoring

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  17. Composition of Individual Particles in the Wakes of an Athena II Rocket and the Space Shuttle

    NASA Astrophysics Data System (ADS)

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

    2002-12-01

    The NOAA Particle Analysis by Laser Mass Spectrometry (PALMS) instrument was used to obtain the first in situ measurements of the composition of particles in the exhaust wakes of launch vehicles powered by solid rocket motors (SRMs). PALMS, mounted in the nose of a NASA WB-57F research aircraft, acquired mass spectra of over 2300 individual exhaust particles during stratospheric encounters with the plumes of an Athena II rocket and the Space Shuttle. The majority of positive mass spectra indicated the presence of Al, Fe, Ca, Na, and K, all primary or trace components of the aluminum fuel or the combustion catalyst. Organic material, presumably from combustion of binding and curing agents, was another common feature. Negative mass spectra showed Cl from the oxidizer, ammonium perchlorate, as well as aluminum oxide produced during combustion. Nitrate and phosphate fragments and water complexes were common features of spectra acquired during the Space Shuttle but not the Athena II plume intercepts. Elemental carbon (EC) was a significant particle type observed in the Athena II plume but not the Space Shuttle. The data show that the composition of particles emitted by SRMs are more diverse, more varied from rocket to rocket, and possibly more reactive than previously considered.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    1975-01-01

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

  20. Scaled Rocket Testing in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Li, Yi; Reimann, Bodo; Eggers, Thino

    2016-11-01

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

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

    SciTech Connect

    Youngblood, Stewart

    2015-08-01

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

  4. Delta 2 Explosion Plume Analysis Report

    NASA Technical Reports Server (NTRS)

    Evans, Randolph J.

    2000-01-01

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

  5. Composition of individual particles in the wakes of an Athena II rocket and the space shuttle

    NASA Astrophysics Data System (ADS)

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

    2002-11-01

    The Particle Analysis by Laser Mass Spectrometry (PALMS) instrument was used to obtain the first in situ measurements of the composition of particles in the wakes of solid rocket motor (SRMs) launch vehicles. PALMS acquired mass spectra of over 2300 exhaust particles within the plumes of an Athena II rocket and the Space Shuttle. The majority of positive spectra indicated the presence of primary and trace components of the aluminum fuel and the combustion catalyst. Negative spectra showed chlorine from the oxidizer. Nitrate and phosphate fragments and water were common features of spectra acquired during the Space Shuttle encounters. Elemental carbon (EC) was a significant particle type observed in the Athena II plume. The data show that particles emitted by SRMs are more diverse and probably more reactive than previously considered.

  6. 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.

  7. 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.

  8. Liquid propellant rockets.

    NASA Technical Reports Server (NTRS)

    Dipprey, D. F.

    1972-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Toohey, D. W.; Ross, M.

    2009-12-01

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

  10. 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.

  11. 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.

  12. 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.

  13. 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

  14. 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.

  15. 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.

  16. 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.

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

    PubMed

    Griggs, M; Harshbarger, F C

    1966-02-01

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

  18. 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.

  19. Chemical rocket propulsion and the environment

    SciTech Connect

    Mcdonald, A.J. )

    1992-03-01

    Results are presented from the examination by the Chemical Rocket Propulsion and the Environment Workshop conducted by AIAA in June 1991 of the impact of rocket launches and ground testing on the earth's environment. The major conclusions of this workshop were: (1) at projected rocket launch rates, neither the liquid- nor the solid-rocket motors will significantly impact stratospheric ozone; (2) there is no global acid rain problem associated with rocket exhaust; and (3) the local launch site and static test site acidification is a minor problem and can be managed.

  20. 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.

  1. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  3. 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

  4. 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.

  5. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  6. 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)

  7. 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.

  8. Ionospheric modification by rocket effluents. Final report

    SciTech Connect

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

    1980-06-01

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

  9. 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.

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

    NASA Technical Reports Server (NTRS)

    Meissner, Dana L.

    1993-01-01

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

  11. Solid rocket booster thermal radiation model, volume 1

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  12. 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.

  13. Ultraviolet Emission from Rocket Motor Plumes

    DTIC Science & Technology

    1994-06-01

    who..t’ "-. bsorption band is centred near 255 nm Ozone concentrations vary with location and show regular htne-,oral variations over diurnal and...34Flame Spectroscopy " (John Wiley and Sons, New York, 1965) 25. Yates, G. J., Wilke, M.. King. N and Lumpkn. A. "Ultraviolet imaging of hydrogen

  14. 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.

  15. Splash Plumes

    NASA Astrophysics Data System (ADS)

    Davies, J. H.

    2006-12-01

    I have discovered a new class of thermal upwellings in mantle convection simulations which are not rooted in a thermal boundary layer (ref 1). Since they look a bit like water droplet splashes, I have abbreviated these `plumes not rooted in thermal boundary layers' as `splash plumes'. These mantle convection simulations are high resolution ( ~ 22km spacing) 3D spherical simulations at Earth-like vigour. They have a chondritic rate of internal heating and bottom heating that straddles expected Earth values. There is a realistic depth variation in viscosity, with a lithosphere and lower mantle more viscous than upper mantle. The mantle is compressible with the coefficient of thermal expansion decreasing with depth. Some models have phase transitions. The surface of the models is driven by 119Myr of recent plate motion history. At the end of most simulations (present day) we discover many examples of hot mid-mantle thermal anomalies in the shape of bowls which have hot cylindrical plumes rising from the rim. They originate at a range of depths and are not rooted in thermal boundary layers. These splash plumes are formed from hot mantle collecting beneath the surface, and then a cold instability from the surface descending onto the sheet of hot underlying material pushing it down into the mantle and forming a bowl. The plumes are formed by instabilities coming from the bowl rim edge. In fact the downwellings can push the sheets all the way to the core mantle boundary in certain cases where it is then difficult to tell splash plumes apart from `traditional plumes'. Splash plumes might provide explanations for weak, short-lived plumes that do not seem to have deep roots (e.g. Eifel). If the surface boundary condition is made free-slip (ref 2), rather than be driven by recent plate motion history, we do not discover splash plumes but rather large steady strong thermal boundary layer plumes. Therefore while the discovery of splash plumes is interesting, potentially a more

  16. Two-dimensional motions of rockets

    NASA Astrophysics Data System (ADS)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

  17. 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.

  18. Geomagnetic effects caused by rocket exhaust jets

    NASA Astrophysics Data System (ADS)

    Lipko, Yuriy; Pashinin, Aleksandr; Khakhinov, Vitaliy; Rahmatulin, Ravil

    2016-09-01

    In the space experiment Radar-Progress, we have made 33 series of measurements of geomagnetic variations during ignitions of engines of Progress cargo spacecraft in low Earth orbit. We used magneto-measuring complexes, installed at observatories of the Institute of Solar-Terrestrial Physics of Siberian Branch of the Russian Academy of Sciences, and magnetotelluric equipment of a mobile complex. We assumed that engine running can cause geomagnetic disturbances in flux tubes crossed by the spacecraft. When analyzing experimental data, we took into account space weather factors: solar wind parameters, total daily mid-latitude geomagnetic activity index Kp, geomagnetic auroral electrojet index AE, global geomagnetic activity. The empirical data we obtained indicate that 18 of the 33 series showed geomagnetic variations in various time ranges.

  19. CFD analysis of a rocket exhaust diffuser

    NASA Astrophysics Data System (ADS)

    Bose, Tarit K.; Thanawala, R. H.; Annamalai, K.

    1992-11-01

    The nature of the complex shock structure responsible for the pressure recovery phenomenon in supersonic diffusers is investigated by means of a theoretical CFD analysis using a newly developed computer program for Navier-Stokes solution of an ejector system, and the Prandtl mixing length to model the turbulent boundary layer. The pressure recovery characteristics of an ejector diffuser system was studied for various geometric and flow conditions. A comparison of the results with those of pressure measurements along the diffuser length in an experimental facility showed discrepancies, which are attributed to the boundary conditions imposed.

  20. 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.

  1. NO sub X Deposited in the Stratosphere by the Space Shuttle Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Pergament, H. S.; Thorpe, R. D.; Hwang, B.

    1975-01-01

    The possible effects of the interaction of the plumes from the two solid rocket motors (SRM) from the space shuttles and mixing of the rocket exhaust products and ambient air in the base recirculation region on the total nitrous oxide deposition rate in the stratosphere were investigated. It was shown that these phenomena will not influence the total NOx deposition rate. It was also shown that uncertainties in the particle size of Al2O3, size distributions and particle/gas drag and heat transfer coefficients will not have a significant effect on the predicted NOx deposition rate. The final results show that the total mass flow of NOx leaving the plume at 30 km altitude is 4000 g./sec with a possible error factor of 3. For a vehicle velocity of 1140 meter/sec this yields an NOx deposition rate of about 3.5 g./meter. The corresponding HCl deposition rate at this altitude is about a factor of 500 greater than this value.

  2. 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)

  3. 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

  4. 12. Building 202 exhaust scrubber facility, looking southwest from catwalk ...

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

    12. Building 202 exhaust scrubber facility, looking southwest from catwalk located northeast of Building 202 pump house. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  5. 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

  6. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  8. Tvashtar's Plume

    NASA Technical Reports Server (NTRS)

    2007-01-01

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

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

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

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

  9. 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.

  10. 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)

  11. 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.

  12. 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.

  13. 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.

  14. Vertical Landing Aerodynamics of Reusable Rocket Vehicle

    NASA Astrophysics Data System (ADS)

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

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

  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. 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...

  17. 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...

  18. 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...

  19. 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...

  20. 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...

  1. Isotopic mapping of groundwater perchlorate plumes.

    PubMed

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

    2012-01-01

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

  2. Ice nuclei measurements from solid rocket motor effluents

    NASA Technical Reports Server (NTRS)

    Hindman, E. E., II

    1980-01-01

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

  3. 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.

  4. Atomic hydrogen rocket engine

    NASA Technical Reports Server (NTRS)

    Etters, R. D.; Flurchick, K.

    1981-01-01

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

  5. 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.

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

    NASA Technical Reports Server (NTRS)

    Plencner, R. M.

    1977-01-01

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

  8. 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.

  9. Magnetic Detachment and Plume Control in Escaping Magnetized Plasma

    SciTech Connect

    P. F. Schmit and N. J. Fisch

    2008-11-05

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

  10. 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...

  11. 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).

  12. 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.

  13. 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.

  14. Modelling of thruster plume induced erosion

    NASA Astrophysics Data System (ADS)

    Alred, John; Boeder, Paul; Mikatarian, Ron; Pankop, Courtney; Schmidl, William

    2003-09-01

    One source of external induced contamination on the International Space Station (ISS) is thruster plume exhausts. The contamination from these plumes onto ISS sensitive surfaces is due to liquid drops of unreacted or partially reacted propellants. However, the drag acceleration of these particles (drops) from the exhaust gases produces high velocity (~km/s) drops that will mechanically damage surfaces in the exhaust. Previous space flight experiments on the Space Shuttle Orbiter which studied thruster plume induced contamination also demonstrated the pitting nature of these particles. The External Contamination/Plasma Team of the Boeing ISS Program Office in Houston has developed an approach to modeling the mechanical erosion on surfaces due to the impact of particles in thruster plumes. This approach melds damage simulation data from a smooth particle hydrodynamics (SPH) code from Los Alamos National Laboratory (LANL) into Boeing's own contamination computer tool (NASAN-II). The Boeing team has conducted several analyses simulating bipropellant thruster droplets impacting ISS sensitive surfaces. Computational results of various thrusters firing onto the ISS, at different build-stages, were completed and show a concern for particular solar array orientations during attitude control firings. Mitigation techniques for minimizing the erosion effects have also been determined and are presented.

  15. 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.

  16. ASSESSMENT OF PLUME DIVING

    EPA Science Inventory

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

  17. 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.

  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. 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.

  20. Problems of the mathematical description of rocket engines as plants

    NASA Astrophysics Data System (ADS)

    Kiforenko, B. N.

    2012-09-01

    Mathematical models of liquid-propellant, nuclear, and electric rocket engines are presented that more fully describe thrust generation than the classical models do. The optimal control of engine thrust is analyzed within the framework of Mayer's general variational problem. It is shown that the control of a rocket engine satisfying the necessary optimality conditions belongs to the boundary arc of the feasible control set between the point of maximum thrust and the point of maximum exhaust velocity

  1. 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.

  2. Diagnostics of an argon arcjet plume with a diode laser.

    PubMed

    Zhang, F Y; Komurasaki, K; Iida, T; Fujiwara, T

    1999-03-20

    The diode-laser absorption technique was applied for simultaneous velocity and temperature measurements of an argon plume exhausted by an arcjet. The Ar I absorption line at 811.531 nm was taken as the center absorption line. The velocity and the temperature were derived from the Doppler shift in the absorption profiles and the full width at half-maximum of the plume absorption profile, respectively. From the measured plume velocity and temperature, the total enthalpy of the exhausted plume, the thrust efficiency, and the thermal efficiency of the arcjet were derived, and the performance of the arcjet was examined. The results are demonstrated to agree with results derived by other methods, and the technique can be applied to the measurement of other arcjet systems without much modification.

  3. Solid Rocket Motor Backflow Analysis For CONTOUR Mishap Investigation

    DTIC Science & Technology

    2005-07-13

    to thank the members of the CONTOUR MIB, especially Mr. Craig Tooley , NASA/GSFC, for their support. He also appreciates the efforts of Messrs. Lou... Tooley , CONTOUR Mishap Investigation Board, SAI-12-627/MSW-2, 20 May 2003, pp. 1-13. 6. M. Woronowicz, “Development of a Novel Free Molecule Rocket Plume

  4. 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.

  5. Research on infrared radiation signatures of high-altitude plume based on DSMC method

    NASA Astrophysics Data System (ADS)

    Bao, Xingdong; Mao, Hongxia; Wu, Jie; Dong, Yanbing

    2016-10-01

    Exhaust plume flow field have the characteristics of high temperature, high speed and multi-species flow. Exhaust plume infrared signal are important basis of diagnosing, detecting and identifying plume spectrum. This paper focuses on the infrared radiation characteristics of high-altitude plume. The plume flows exhausted from a micro-nozzle of a low-thrust engine at high-altitude have been simulated numerically through using a DSMC method. Both the properties of plume flow at high altitude and the non-equilibrium effect related to rarefied gases are analyzed. Results are given numerically in good agreement with high-altitude plume observations. With the fields of pressure, temperature and main components of the exhaust plume as input data, the line-by line method was used to calculate the 2 5μm infrared spectral radiation properties of the plume. Different flight conditions are considered to analyze the influence on the infrared radiation characteristics. Some interesting conclusion are finally achieved.

  6. 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.

  7. 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.

  8. Measurement and infrared image prediction of a heated exhaust flow

    NASA Astrophysics Data System (ADS)

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

    1994-06-01

    The focus of the current 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 CFD codes through infrared imaging. The technique of reducing the 3D field-variable domain to a 2D infrared image invokes the use of an inverse Monte-Carlo ray trace algorithm and an infrared band model for exhaust gases. This paper describes an experiment in which the above- mentioned field variables were carefully measured. Data from this experiment in the form of velocity plots are shown. The inverse Monte-Carlo ray trace technique is described. Finally, an experimentally obtained infrared image is directly compared to an infrared image predicted from the measured field variables.

  9. 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

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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).

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. A comparison of two methods of measuring particle size of Al2O3 produced by a small rocket motor

    NASA Technical Reports Server (NTRS)

    Dobbins, R. A.; Strand, L. D.

    1969-01-01

    The size of aluminum oxide particles produced by small rocket motors is determined by tank collection and spectrophotometry. The size of the particulate determines loss in thrust due to particle lag, particulate radiant heat transfer, acoustic attenuation and impingement and rocket plume structure and properties.

  3. 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.

  4. 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

  5. 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

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

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

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

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

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

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

  8. 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. 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).

  10. 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.

  11. 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.

  12. 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

  13. 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

  14. Liquid Rocket Engine Testing

    DTIC Science & Technology

    2016-10-21

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

  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. Rocket astronomy - an overview

    NASA Astrophysics Data System (ADS)

    Friedman, H.

    The history of rocket astronomy is retold, with emphasis on solar investigations. The use of captured V-2 rockets after World War II was followed by the Aerobee, which exceeded the V-2's altitude and was more reliable. The V-2 has made first-generation investigations in X-ray and UV photometry, which was followed by studies of the solar cycle X-ray variablity, the X-ray corona, and solar flares. Nike rockets played an important role in these investigations. The role of rockets in galactic X-ray astronomy is briefly described.

  17. 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.

  18. Four dimensional reconstruction and analysis of plume images

    NASA Astrophysics Data System (ADS)

    Dhawan, Atam P.; Disimile, Peter J.; Peck, Charles, III

    Results of a time-history based three-dimensional reconstruction of cross-sectional images corresponding to a specific planar location of the jet structure are reported. The experimental set-up is described, and three-dimensional displays of time-history based reconstruction of the jet structure are presented. Future developments in image analysis, quantification and interpretation, and flow visualization of rocket engine plume images are expected to provide a tool for correlating engine diagnostic features with visible flow structures.

  19. 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.

  20. Mechanisms of Exhaust Pollutants and Plume Formation in Continuous Combustion.

    DTIC Science & Technology

    1984-06-01

    IN CONTINUOUS COMBUSTION E AFOSR 78-3586 G.S. Samuelsen UCI -ARTR-84-7 ci~LAi June 1984 Ř Mechanical Engineering Univgrsity of California Irvine...ORGANIZATION ’b- OFFICE SYMBOL 74. NAME OF MONITORING ORGANIZATION University of California I IfiapphuabielArFreOfi fSintfcRsac UCI Combustion ...SUPPLEMENTARY NOTATION 17 COSATI CODES iI19. SUBJaECT TERI,.S (Continue on pWar~tlee~r aid ida ify by blocir nurnbar, ’efLO TGROUP I sue. R ’ Combustion and

  1. Nuclear thermal rockets using indigenous Martian propellants

    SciTech Connect

    Zubrin, R.M.

    1989-01-01

    This paper considers a novel concept for a Martian descent and ascent vehicle, called NIMF (for nuclear rocket using indigenous Martian fuel), the propulsion for which will be provided by a nuclear thermal reactor which will heat an indigenous Martian propellant gas to form a high-thrust rocket exhaust. The performance of each of the candidate Martian propellants, which include CO2, H2O, CH4, N2, CO, and Ar, is assessed, and the methods of propellant acquisition are examined. Attention is also given to the issues of chemical compatibility between candidate propellants and reactor fuel and cladding materials, and the potential of winged Mars supersonic aircraft driven by this type of engine. It is shown that, by utilizing the nuclear landing craft in combination with a hydrogen-fueled nuclear thermal interplanetary vehicle and a heavy lift booster, it is possible to achieve a manned Mars mission in one launch. 6 refs.

  2. Automated and Manual Rocket Crater Measurement Software

    NASA Technical Reports Server (NTRS)

    Metzger, Philip; Immer, Christopher

    2012-01-01

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

  3. Interpretation of Core Length in Shear Coaxial Rocket Injectors from X-ray Radiography Measurements (Briefing Charts)

    DTIC Science & Technology

    2014-06-01

    Shear coaxial jets can be found in a number of combustion devices – Turbofan engine exhaust, air blast furnaces, and liquid rocket engines ...case where both fluids are supercritical – Common in modern boost-class liquid rocket engines • Current focus is on two-phase coaxial jets for... rocket engine applications – Common in upper-stage engines and during throttled conditions, startup and shutdown transients of boost-class engines

  4. 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

  5. Pollutant monitoring of aircraft exhaust with multispectral imaging

    NASA Astrophysics Data System (ADS)

    Berkson, Emily E.; Messinger, David W.

    2016-10-01

    Communities surrounding local airports are becoming increasingly concerned about the aircraft pollutants emitted during the landing-takeoff (LTO) cycle, and their potential for negative health effects. Chicago, Los Angeles, Boston and London have all recently been featured in the news regarding concerns over the amount of airport pollution being emitted on a daily basis, and several studies have been published on the increased risks of cancer for those living near airports. There are currently no inexpensive, portable, and unobtrusive sensors that can monitor the spatial and temporal nature of jet engine exhaust plumes. In this work we seek to design a multispectral imaging system that is capable of tracking exhaust plumes during the engine idle phase, with a specific focus on unburned hydrocarbon (UHC) emissions. UHCs are especially potent to local air quality, and their strong absorption features allow them to act as a spatial and temporal plume tracer. Using a Gaussian plume to radiometrically model jet engine exhaust, we have begun designing an inexpensive, portable, and unobtrusive imaging system to monitor the relative amount of pollutants emitted by aircraft in the idle phase. The LWIR system will use two broadband filters to detect emitted UHCs. This paper presents the spatial and temporal radiometric models of the exhaust plume from a typical jet engine used on 737s. We also select filters for plume tracking, and propose an imaging system layout for optimal detectibility. In terms of feasibility, a multispectral imaging system will be two orders of magnitude cheaper than current unobtrusive methods (PTR-MS) used to monitor jet engine emissions. Large-scale impacts of this work will include increased capabilities to monitor local airport pollution, and the potential for better-informed decision-making regarding future developments to airports.

  6. Numerical Investigation of Twin-Nozzle Rocket Plume Phenomenology

    DTIC Science & Technology

    2007-11-02

    of three-dimensionality, turbulent viscosity, afterburning, near-field shock structure, finite-rate kinetic chemistry , internozzle geometric spacing...stress model, kinetic chemistry especially at lower altitudes, and nozzle exit profile are important parameters that should be considered in the

  7. 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)

  8. 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.

  9. 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)

  10. Measuring Model Rocket Acceleration.

    ERIC Educational Resources Information Center

    Jenkins, Randy A.

    1993-01-01

    Presents an experiment that measures the acceleration and velocity of a model rocket. Lift-off information is transmitted to a computer that creates a graph of the velocity. Discusses the analysis of the computer-generated data and differences between calculated and experimental velocity and acceleration of several rocket types. (MDH)

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  12. Entrainment by Lazy Plumes

    NASA Astrophysics Data System (ADS)

    Kaye, Nigel; Hunt, Gary

    2004-11-01

    We consider plumes with source conditions that have a net momentum flux deficit compared to a pure plume - so called lazy plumes. We examine both the case of constant buoyancy flux and buoyancy flux linearly increasing with height. By re-casting the plume conservation equations (Morton, Taylor & Turner 1956) for a constant entrainment coefficient ((α)) in terms of the plume radius (β) and the dimensionless parameter (Γ=frac5Q^2 B4α M^5/2) we show that the far-field flow in a plume with a linear internal buoyancy flux gain is a constant velocity lazy plume with reduced entrainment and radial growth rate. For highly lazy source conditions we derive first-order approximate solutions which indicate a region of zero entrainment near the source. These phenomena have previously been observed, however, it has often been assumed that reduced entrainment implies a reduced (α). We demonstrate that a constant (α) formulation is able to capture the behaviour of these reduced entrainment flows. Morton, B. R., Taylor, G. I. & Turner, J. S. (1956), Turbulent gravitational convection from maintained and instantaneous sources.', Proc. Roy. Soc. 234, 1-23.

  13. 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.

  14. RADAR ATTENUATION BY SOLID-PROPELLANT-ROCKET EXHAUST.

    DTIC Science & Technology

    source of attenuation with both plastisol -nitrocellulose and inert-binder composite propellants. Aluminum is a second major cause of radar attenuation...concentrations and collision frequencies in the product gases; for plastisol -nitrocellulose propellants attenuation at one frequency can be inferred from measured

  15. 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.

  16. 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.

  17. 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.

  18. 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).

  19. 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...

  20. Microphysical properties of the Shuttle exhaust cloud

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    A data base describing the properties of the exhaust cloud produced by the launch of the STS has been developed based on data from a series of ground and aircraft based measurements made during the launches of STS 2, 3, and 4. Aircraft observations were performed during the STS-3 launch with a NOAA WP-3D Orion hurricane research aircraft which contained instrumentation for cloud condensation nucleus and ice nucleus counting, Aitken particle counting, and pH determination. Ground observations were conducted at 50 different sites, as well as in the direct exhaust from the solid rocket booster flame trench at all three launches. The data is analyzed in order to determine any possible adverse impacts of the exhaust products on human health and/or the environment. Analyses of the exhaust cloud measurements indicate that in the case of the ground cloud where plenty of large water drops are present and considerable scavenging and fallout of aerosol takes place, possible adverse impacts of the remaining aerosols (CCN and IN) on natural precipitation processes which may occur in the launch area hours after the launch are remote. However, it is determined that under certain atmospheric conditions there could be short term adverse effects on visibility.

  1. 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.

  2. 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

  3. Plume Measurement System (PLUMES) Calibration Experiment

    DTIC Science & Technology

    1994-08-01

    Atle Lohrmann SonTek, Inc. 7940 Silverton Avenue, No. 105 San Diego, California 92126 and Craig Huhta JIMAR University of Hawaii, Honolulu, Hawaii 96822...Measurement System (PLUMES) Calibration Experiment by Age Lohrmann SonTek, Inc. 7940 Silverton Avenue, No. 105 San Diego, CA 92126 Craig Huhta JIMAR...PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) &. PERFORMING ORGANIZATION SonTek, Inc., 7940 Silverton Avenue, No. 105, San Diego, CA 92126 REPORT NUMBER

  4. 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.

  5. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

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

  6. 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.

  7. 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.

  8. The History of Rockets.

    ERIC Educational Resources Information Center

    Newby, J. C.

    1988-01-01

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

  9. 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.

  10. Rocket engine numerical simulator

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

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

  11. 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.

  12. 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)

  13. 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...

  14. 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.

  15. 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.

  16. 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.

  17. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

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

  18. 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.

  19. 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.

  20. 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

  1. 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.

  2. Four dimensional reconstruction and analysis of plume images

    NASA Astrophysics Data System (ADS)

    Dhawan, Atam P.; Peck, Charles, III; Disimile, Peter

    1991-05-01

    A number of methods have been investigated and are under current investigation for monitoring the health of the Space Shuttle Main Engine (SSME). Plume emission analysis has recently emerged as a potential technique for correlating the emission characteristics with the health of an engine. In order to correlate the visual and spectral signatures of the plume emission with the characteristic health monitoring features of the engine, the plume emission data must be acquired, stored, and analyzed in a manner similar to flame emission spectroscopy. The characteristic visual and spectral signatures of the elements vaporized in exhaust plume along with the features related to their temperature, pressure, and velocity can be analyzed once the images of plume emission are effectively acquired, digitized, and stored on a computer. Since the emission image varies with respect to time at a specified planar location, four dimensional visual and spectral analysis need to be performed on the plume emission data. In order to achieve this objective, feasibility research was conducted to digitize, store, analyze, and visualize the images of a subsonic jet in a cross flow. The jet structure was made visible using a direct injection flow visualization technique. The results of time-history based three dimensional reconstruction of the cross sectional images corresponding to a specific planar location of the jet structure are presented. The experimental set-up to acquire such data is described and three dimensional displays of time-history based reconstructions of the jet structure are discussed.

  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. Augmentation of Rocket Propulsion: Physical Limits

    NASA Technical Reports Server (NTRS)

    Taylor, Charles R.

    1996-01-01

    Rocket propulsion is not ideal when the propellant is not ejected at a unique velocity in an inertial frame. An ideal velocity distribution requires that the exhaust velocity vary linearly with the velocity of the vehicle in an inertial frame. It also requires that the velocity distribution variance as a thermodynamic quantity be minimized. A rocket vehicle with an inert propellant is not optimal, because it does not take advantage of the propellant mass for energy storage. Nor is it logical to provide another energy storage device in order to realize variable exhaust velocity, because it would have to be partly unfilled at the beginning of the mission. Performance is enhanced by pushing on the surrounding because it increases the reaction mass and decreases the reaction jet velocity. This decreases the fraction of the energy taken away by the propellant and increases the share taken by the payload. For an optimal model with the propellant used as fuel, the augmentation realized by pushing on air is greatest for vehicles with a low initial/final mass ratio. For a typical vehicle in the Earth's atmosphere, the augmentation is seen mainly at altitudes below about 80 km. When drag is taken into account, there is a well-defined optimum size for the air intake. Pushing on air has the potential to increase the performance of rockets which pass through the atmosphere. This is apart from benefits derived from "air breathing", or using the oxygen in the atmosphere to reduce the mass of an on-board oxidizer. Because of the potential of these measures, it is vital to model these effects more carefully and explore technology that may realize their advantages.

  5. 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.

  6. Laser Rayleigh and Raman diagnostics for small hydrogen/oxygen rockets

    NASA Astrophysics Data System (ADS)

    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.

  7. Laser Rayleigh and Raman diagnostics for small hydrogen/oxygen rockets

    NASA Astrophysics Data System (ADS)

    de Groot, Wim A.; Zupanc, Frank J.

    1993-05-01

    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 a 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.

  8. Nuclear Rocket Technology Conference

    NASA Technical Reports Server (NTRS)

    1966-01-01

    The Lewis Research Center has a strong interest in nuclear rocket propulsion and provides active support of the graphite reactor program in such nonnuclear areas as cryogenics, two-phase flow, propellant heating, fluid systems, heat transfer, nozzle cooling, nozzle design, pumps, turbines, and startup and control problems. A parallel effort has also been expended to evaluate the engineering feasibility of a nuclear rocket reactor using tungsten-matrix fuel elements and water as the moderator. Both of these efforts have resulted in significant contributions to nuclear rocket technology. Many successful static firings of nuclear rockets have been made with graphite-core reactors. Sufficient information has also been accumulated to permit a reasonable Judgment as to the feasibility of the tungsten water-moderated reactor concept. We therefore consider that this technoIogy conference on the nuclear rocket work that has been sponsored by the Lewis Research Center is timely. The conference has been prepared by NASA personnel, but the information presented includes substantial contributions from both NASA and AEC contractors. The conference excludes from consideration the many possible mission requirements for nuclear rockets. Also excluded is the direct comparison of nuclear rocket types with each other or with other modes of propulsion. The graphite reactor support work presented on the first day of the conference was partly inspired through a close cooperative effort between the Cleveland extension of the Space Nuclear Propulsion Office (headed by Robert W. Schroeder) and the Lewis Research Center. Much of this effort was supervised by Mr. John C. Sanders, chairman for the first day of the conference, and by Mr. Hugh M. Henneberry. The tungsten water-moderated reactor concept was initiated at Lewis by Mr. Frank E. Rom and his coworkers. The supervision of the recent engineering studies has been shared by Mr. Samuel J. Kaufman, chairman for the second day of the

  9. Apollo experience report: Thermal protection from engine-plume environments

    NASA Technical Reports Server (NTRS)

    Taylor, J. T.

    1972-01-01

    Portions of the combined Apollo spacecraft (the command and service module and the lunar module) are subjected to the impingement of hot exhaust gases from the various propulsion systems of the modules. The configurations of the vehicles and the sources of impinging engine plumes are described. A typical Apollo mission is outlined. Protection and design-verification methods are discussed. Finally, recommendations are made for future spacecraft programs.

  10. Where Plumes Live

    NASA Astrophysics Data System (ADS)

    King, S. D.

    2004-12-01

    From the perspective of fluid dynamics, `Plumes or not?' might be the wrong question. Let me begin by defining a few terms. Plume with a `P' is the well-known thermal structure with thin (order 100 km) tail and large, bulbous head that originates at the core-mantle boundary. The thin tail/large, bulbous-head morphology has been generated in a number of laboratory and numerical experiments. It can be seen, for example, on the cover of the famous fluid dynamics text by Batchelor. There is a clearly-defined range of parameters for which this structure is the preferred solution for instabilities arising from a bottom boundary layer in a convecting fluid. For example, a strong temperature-dependent rheology is needed. By contrast, plume with a `p' is any cylindrical or quasi-cylindrical instability originating from a thermal (or thermo-chemical) boundary layer. In fluid dynamics plume is sometimes used interchangeable with jet. Unless there is a very small temperature drop across the core-mantle boundary or a rather remarkable balance between temperature and composition at the base of the mantle, there are almost certainly plumes. (Note the little p.) Are these plumes the thermal structures with thin (order 100 km) tails and large bulbous heads or could they be broad, hot regions such as the degree 2 pattern seen in global seismic tomography images of the lower mantle, or the disconnected droplets seen in chaotic convection? To study this question, I will present a sequence of numerical `experiments' that illustrate the morphology of instabilities from a basal thermal boundary layer, i.e., plumes. Some of the aspects I will present include: spherical geometry, temperature-and pressure-dependence of rheology, internal heating, pressure-dependent coefficient of thermal expansion, variable coefficient of thermal diffusivity, phase transformations, and compositional layering at the base of the mantle. The goal is to map out the parameters and conditions where Plumes live

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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...

  16. 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.

  17. EUV analysis of polar plumes

    NASA Technical Reports Server (NTRS)

    Ahmad, I. A.; Withbroe, G. L.

    1977-01-01

    Three polar plumes were studied using Skylab Mg X and O VI data. The plumes lie within the boundaries of a polar coronal hole. We find that the mean temperature of the plumes is about 1.1 million K and that they have a small vertical temperature gradient. Densities are determined and found consistent with white light analyses. The variation of density with height in the plumes is compared with that expected for hydrostatic equilibrium. As is the case for other coronal features, polar plumes will be a source of solar wind if the magnetic field lines are open. On the basis of the derived plume model and estimates of the numbers of plumes in polar coronal holes, it appears that polar plumes contain about 15% of the mass in a typical polar hole and occupy about 10% of the volume.

  18. 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)

  19. 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

  20. Density and optical properties of SPARCS plumes

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  1. 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.

  2. Enceladus' Water Vapour Plumes

    NASA Technical Reports Server (NTRS)

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

    2006-01-01

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

  3. Double Diffusive Plumes

    NASA Astrophysics Data System (ADS)

    Sutherland, Bruce; Lee, Brace

    2008-11-01

    Sour gas flares attempt to dispose of deadly H2S gas through combustion. What does not burn rises as a buoyant plume. But the gas is heavier than air at room temperature, so as the rising gas cools eventually it becomes negatively buoyant and descends back to the ground. Ultimately, our intent is to predict the concentrations of the gas at ground level in realistic atmospheric conditions. As a first step towards this goal we have performed laboratory experiments examining the structure of a steady state plume of hot and salty water that rises buoyantly near the source and descends as a fountain after it has cooled sufficiently. We call this a double-diffusive plume because its evolution is dictated by the different (turbulent) diffusivities of heat and salt. A temperature and conductivity probe measures both the salinity and temperature along the centreline of the plume. The supposed axisymmetric structure of the salinity concentration as it changes with height is determined by light-attenuation methods. To help interpret the results, a theory has been successfully adapted from the work of Bloomfield and Kerr (2000), who developed coupled equations describing the structure of fountains. Introducing a new empirical parameter for the relative rates of turbulent heat and salt diffusion, the predictions are found to agree favourably with experimental results.

  4. 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…

  5. PLUME and research sotware

    NASA Astrophysics Data System (ADS)

    Baudin, Veronique; Gomez-Diaz, Teresa

    2013-04-01

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

  6. Buoyant plume calculations

    SciTech Connect

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

    1985-01-01

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

  7. MK 66 Rocket Signature Reduction

    DTIC Science & Technology

    1982-04-01

    Indian Head, Maryland. ’The objec- tive of the study was to reduce the visible signature of the rocket motor. The rocket motor used for demonstration tests...15 6. Actual Emmiissions . . . . . . ........... . 16 7. Human Eye Adjusted Emmissions ..................... .. 16 8. Cross...altered. Additives are commonly used in gun propellants for elimination of muzzle flash. Their use in tactical rockets has been very limited, and

  8. 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…

  9. Hermes A-1 Test Rockets

    NASA Technical Reports Server (NTRS)

    1950-01-01

    The first Hermes A-1 test rocket was fired at White Sand Proving Ground (WSPG). Hermes was a modified V-2 German rocket, utilizing the German aerodynamic configuration; however, internally it was a completely new design. Although it did not result in an operational vehicle, the information that was gathered in the process contributed directly to the development of the Redstone rocket.

  10. Overview of rocket engine control

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Musgrave, Jeffrey L.

    1991-01-01

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

  11. 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.

  12. Solid Rocket Booster Recovery

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The towing ship, Liberty, towed a recovered solid rocket booster (SRB) for the STS-5 mission to Port Canaveral, Florida. The recovered SRB would be inspected and refurbished for reuse. The Shuttle's SRB's and solid rocket motors (SRM's) are the largest ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds. The requirement for reusability dictated durable materials and construction to preclude corrosion of the hardware exposed to the harsh seawater environment. The SRB contains a complete recovery subsystem that includes parachutes, beacons, lights, and tow fixture.

  13. Solid Rocket Booster Recovery

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The towing ship, Liberty, towed a recovered solid rocket booster (SRB) for the STS-3 mission to Port Canaveral, Florida. The recovered SRB would be inspected and refurbished for reuse. The Shuttle's SRB's and solid rocket motors (SRM's) are the largest ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds. The requirement for reusability dictated durable materials and construction to preclude corrosion of the hardware exposed to the harsh seawater environment. The SRB contains a complete recovery subsystem that includes parachutes, beacons, lights, and tow fixture.

  14. Jupiter Rocket Engine

    NASA Technical Reports Server (NTRS)

    2004-01-01

    Engine for the Jupiter rocket. The Jupiter vehicle was a direct derivative of the Redstone. The Army Ballistic Missile Agency (ABMA) at Redstone Arsenal, Alabama, continued Jupiter development into a successful intermediate ballistic missile, even though the Department of Defense directed its operational development to the Air Force. ABMA maintained a role in Jupiter RD, including high-altitude launches that added to ABMA's understanding of rocket vehicle operations in the near-Earth space environment. It was knowledge that paid handsome dividends later.

  15. 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.

  16. 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.

  17. Dynamics of thermochemical plumes: 2. Complexity of plume structures and its implications for mapping mantle plumes

    NASA Astrophysics Data System (ADS)

    Lin, Shu-Chuan; van Keken, Peter E.

    2006-03-01

    The mantle plume hypothesis provides explanations for several major observations of surface volcanism. The dynamics of plumes with purely thermal origin has been well established, but our understanding of the role of compositional variations in the Earth's mantle on plume formation is still incomplete. In this study we explore the structures of plumes originating from a thermochemical boundary layer at the base of the mantle in an attempt to complement fluid dynamical studies of purely thermal plumes. Our numerical experiments reveal diverse characteristics of thermochemical plumes that frequently deviate from the classic features of plumes. In addition, owing to the interplay between the thermal and compositional buoyancy forces, the morphology, temperature, and flow fields in both the plume head and plume conduit are strongly time-dependent. The entrainment of the dense layer and secondary instabilities developed in the boundary layer contribute to lateral heterogeneities and enhance stirring processes in the plume head. Our models show that substantial topography of the compositional layer can develop simultaneously with the plumes. In addition, plumes may be present in the lower mantle for more than 70 million years. These features may contribute to the large low seismic velocity provinces beneath the south central Pacific, the southern Atlantic Ocean, and Africa. Our model results support the idea that the dynamics of mantle plumes is much more complicated than conventional thinking based on studies of purely thermal plumes. The widely used criteria for mapping mantle plumes, such as a vertically continuous low seismic velocity signature and strong surface topography swell, may not be universally applicable. We propose that the intrinsic density contrast of the distinct composition may reduce the associated topography of some large igneous provinces such as Ontong Java.

  18. 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…

  19. 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.

  20. 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

  1. Rocket Auroral Correlator Experiment

    NASA Technical Reports Server (NTRS)

    LaBelle, James

    2003-01-01

    Dartmouth College provided a multi-channel high- and low- frequency wave receivers, including active sensors on deployable booms, to the Rocket Auroral Correlator Experiment launched from Poker Flat, Alaska, in January 2002. College also performed preliminary analysis of the data. Details are outlined in chronological order.

  2. 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.

  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. 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.

  5. 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.

  6. 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…

  7. 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.

  8. 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)

  9. Sampling by mantle plumes : the legacy of the plume source

    NASA Astrophysics Data System (ADS)

    Brandeis, G.; Touitou, F.; Davaille, A.

    2013-12-01

    Plumes in the Earth's mantle are considered to be at the origin of intraplate volcanism (or hotspots). They continue to fascinate the scientific community by the heterogeneity of the material they sample on the surface of our planet. To characterize what part of the mantle is sampled by plumes, we have developed a laboratory model for laminar thermal plumes at high Prandtl number, in a fluid whose viscosity depends strongly on the temperature. This study describes the precise phenomenology of the plume and proposes scaling laws for the speed and temperature of the conduit of the plume. We show a strong dependence of these features of the plume with the Rayleigh number and viscosity ratio. Our visualization technique allows for the simultaneous non-intrusive measurements of the temperature, deformation and velocity fields. By calculating numerically the advection of passive markers through the experimental velocity field, we found that (1) the hot center of the plume conduit only consists of fluid which has passed through the thermal boundary layer ("TBL") at the bottom of the tank from which the plume was issued. Moreover, as material is stretched by velocity gradients, it is also in the thermal boundary layer that most of the material stretching occurs (2). The fluid is then transported in the conduit without lateral mixing, and further stretched vertically by the lateral velocity gradients. Since it is only the hot upwelling plume center which melts and therefore is sampled by volcanic activity, (1) implies that the plume geochemical signature is representative of the material located in the deep TBL of the mantle from which the plume is issued. On the other hand, (2) implies that filaments, pancakes, and concentric or bimodal zonation of the plume at the surface all result from different distributions of the heterogeneities in the plume source, filaments being the most generic case. Finally, we apply the scaling laws to the case of Hawaii.

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

    EPA Science Inventory

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

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

    NASA Astrophysics Data System (ADS)

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

    2013-08-01

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

  12. Elimination of rocket engine asymmetric loads during tests at sea level

    NASA Technical Reports Server (NTRS)

    Johnson, J. R.

    1966-01-01

    Secondary injection concept eliminates asymmetric loads and may increase thrust rocket engine loads during sea level tests. The concept uses either a tubular manifold with evenly spaced injection ports or secondary fluid injected at the turbine exhaust inlet to the thrust chamber.

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

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  14. Iridium-Coated Rhenium Radiation-Cooled Rockets

    NASA Technical Reports Server (NTRS)

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

    1997-01-01

    Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion, attitude control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for planetary space- craft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber temperature capability. The material system that is currently used for radiation-cooled rockets, a niobium alloy (C103) with a fused silica coating, has a maximum operating temperature of 1370 C. Temperature limitations of C103 rockets force the use of fuel film cooling, which degrades rocket performance and, in some cases, imposes a plume contamination issue from unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium (Ir) coating has demonstrated operation at high temperatures (2200 C) and for long lifetimes (hours). The added thermal margin afforded by iridium-coated rhenium (Ir/Re) allows reduction or elimination of fuel film cooling. This, in turn, leads to higher performance and cleaner spacecraft environments. There are ongoing government- and industry-sponsored efforts to develop flight Ir/ Re engines, with the primary focus on 440-N, apogee insertion engines. Complementing these Ir/Re engine development efforts is a program to address specific concerns and fundamental characterization of the Ir/Re material system, including (1) development of Ir/Re rocket fabrication methods, (2) establishment of critical Re mechanical properly data, (3) development of reliable joining methods, and (4) characterization of Ir/Re life-limiting mechanisms.

  15. Conditionally analyzed mean velocity and turbulence measurements in a plume-induced boundary layer separated flowfield

    NASA Astrophysics Data System (ADS)

    Palko, Carl Wayne

    1998-11-01

    A supersonic plume-induced boundary layer separated (PIBLS) flowfield occurs when an underexpanded exhaust plume obstructs the flow around a rocket causing an oblique shock wave to form on the afterbody. The shock oscillates randomly in the streamwise direction causing an unsteady boundary layer separation that complicates prediction and measurement of PIBLS flowfields. This study provides the first turbulence measurements in a PIBLS flowfield and, consequently, the first benchmark data for evaluating future computational models for such flows. Conditionally analyzed two-component laser Doppler velocimetry (LDV) measurements were made in a planar, two-dimensional PIBLS flow containing an unsteady oblique shock wave formed by the convergence of two supersonic streams past a thick plate. High-speed wall pressure measurements were used to locate the shock wave and, consequently, allow separation of the effects of shock wave motion from the turbulence fluctuations in the velocity measurements of a shock-separated free shear layer. It was found that isolating the large-scale changes in the shock position from the turbulence reduces the experimental scatter rather than substantially changing the shapes or magnitudes of the turbulent stress profiles. Changes in shock motion direction, however, do significantly alter the turbulent stresses. This is the first evidence of the dependence of turbulence level and structure on shock motion direction and the first direct evidence of the effects of changes in shock wave position on turbulence amplification. The shock-induced separation process was found to dramatically increase the streamwise and transverse Reynolds normal stresses (which both peak near reattachment), the primary shear stress, and the normal stress anisotropy. The shock-separated shear layer consists of only a single layer with a large initial growth rate followed by a much smaller growth rate, instead of the two layers found in rapidly expanded shear layers. The

  16. 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.

  17. Modeling the near acoustic field of a rocket during launch

    NASA Technical Reports Server (NTRS)

    Mauritzen, David W.

    1989-01-01

    The design of launch pad structures is critically dependent upon the stresses imposed by the acoustical pressure field generated by the rocket engines during launch. The purpose of this effort is to better describe the acoustical field in the immediate launch area. Since the problem is not analytically tractable, empirical modeling will be employed so that useful results may be obtained for structural design purposes. The plume of the rocket is considered to be a volumetric acoustic source, and is broken down into incremental contributing volumes. A computer program has been written to sum all the contributions to find the total sound pressure level at an arbitrary point. A constant density source is initially assumed and the acoustic field evaluated for several cases to verify the correct operation of the program.

  18. Hyperventilation and exhaustion syndrome.

    PubMed

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

    2014-12-01

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

  19. 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…

  20. Seismic Imaging of Mantle Plumes

    NASA Astrophysics Data System (ADS)

    Nataf, Henri-Claude

    The mantle plume hypothesis was proposed thirty years ago by Jason Morgan to explain hotspot volcanoes such as Hawaii. A thermal diapir (or plume) rises from the thermal boundary layer at the base of the mantle and produces a chain of volcanoes as a plate moves on top of it. The idea is very attractive, but direct evidence for actual plumes is weak, and many questions remain unanswered. With the great improvement of seismic imagery in the past ten years, new prospects have arisen. Mantle plumes are expected to be rather narrow, and their detection by seismic techniques requires specific developments as well as dedicated field experiments. Regional travel-time tomography has provided good evidence for plumes in the upper mantle beneath a few hotspots (Yellowstone, Massif Central, Iceland). Beneath Hawaii and Iceland, the plume can be detected in the transition zone because it deflects the seismic discontinuities at 410 and 660 km depths. In the lower mantle, plumes are very difficult to detect, so specific methods have been worked out for this purpose. There are hints of a plume beneath the weak Bowie hotspot, as well as intriguing observations for Hawaii. Beneath Iceland, high-resolution tomography has just revealed a wide and meandering plume-like structure extending from the core-mantle boundary up to the surface. Among the many phenomena that seem to take place in the lowermost mantle (or D''), there are also signs there of the presence of plumes. In this article I review the main results obtained so far from these studies and discuss their implications for plume dynamics. Seismic imaging of mantle plumes is still in its infancy but should soon become a turbulent teenager.

  1. Engine for Redstone Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    This photograph is of the engine for the Redstone rocket. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the 'reliable workhorse' for America's early space program. As an example of its versatility, the Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile.

  2. 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.

  3. ISRO's solid rocket motors

    NASA Astrophysics Data System (ADS)

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

    1989-08-01

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

  4. 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.

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

    NASA Technical Reports Server (NTRS)

    Lane, John E.; Metzger, Philip T.

    2014-01-01

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

  6. The Aries sounding rocket

    NASA Astrophysics Data System (ADS)

    Dooling, D.

    1980-02-01

    A family of sounding rockets called Aries, using the motors from obsolete Minuteman ICBMs, is described. Payloads for Aries range from 1,500 to 3,500 lb (with a payload diameter of 44 in.) and include various instruments (magnetospheric tracers, X-ray and extreme ultraviolet astronomy and a large X-ray telescope). Prospects for future launching of a two and even three-stage Aries are discussed.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Mantle plumes and continental tectonics.

    PubMed

    Hill, R I; Campbell, I H; Davies, G F; Griffiths, R W

    1992-04-10

    Mantle plumes and plate tectonics, the result of two distinct modes of convection within the Earth, operate largely independently. Although plumes are secondary in terms of heat transport, they have probably played an important role in continental geology. A new plume starts with a large spherical head that can cause uplift and flood basalt volcanism, and may be responsible for regional-scale metamorphism or crustal melting and varying amounts of crustal extension. Plume heads are followed by narrow tails that give rise to the familiar hot-spot tracks. The cumulative effect of processes associated with tail volcanism may also significantly affect continental crust.

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

    NASA Astrophysics Data System (ADS)

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

    2009-12-01

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

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

    NASA Technical Reports Server (NTRS)

    Stephens, J. B. (Editor)

    1974-01-01

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

  15. Exhaust bypass flow control for exhaust heat recovery

    DOEpatents

    Reynolds, Michael G.

    2015-09-22

    An exhaust system for an engine comprises an exhaust heat recovery apparatus configured to receive exhaust gas from the engine and comprises a first flow passage in fluid communication with the exhaust gas and a second flow passage in fluid communication with the exhaust gas. A heat exchanger/energy recovery unit is disposed in the second flow passage and has a working fluid circulating therethrough for exchange of heat from the exhaust gas to the working fluid. A control valve is disposed downstream of the first and the second flow passages in a low temperature region of the exhaust heat recovery apparatus to direct exhaust gas through the first flow passage or the second flow passage.

  16. Hybrid Exhaust Component

    NASA Technical Reports Server (NTRS)

    Pelletier, Gerard D. (Inventor); Logan, Charles P. (Inventor); McEnerney, Bryan William (Inventor); Haynes, Jeffrey D. (Inventor)

    2015-01-01

    An exhaust includes a wall that has a first composite material having a first coefficient of thermal expansion and a second composite material having a second coefficient of the thermal expansion that is less than the first coefficient of thermal expansion.

  17. Diesel engine exhaust

    Integrated Risk Information System (IRIS)

    Diesel engine exhaust ; CASRN N.A . Human health assessment information on a chemical substance is included in the IRIS database only after a comprehensive review of toxicity data , as outlined in the IRIS assessment development process . Sections I ( Health Hazard Assessments for Noncarcinogenic Ef

  18. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2015-09-01

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

  20. Optimum rocket propulsion for energy-limited transfer

    NASA Technical Reports Server (NTRS)

    Zuppero, Anthony; Landis, Geoffrey A.

    1991-01-01

    In order to effect large-scale return of extraterrestrial resources to Earth orbit, it is desirable to optimize the propulsion system to maximize the mass of payload returned per unit energy expended. This optimization problem is different from the conventional rocket propulsion optimization. A rocket propulsion system consists of an energy source plus reaction mass. In a conventional chemical rocket, the energy source and the reaction mass are the same. For the transportation system required, however, the best system performance is achieved if the reaction mass used is from a locally available source. In general, the energy source and the reaction mass will be separate. One such rocket system is the nuclear thermal rocket, in which the energy source is a reactor and the reaction mass a fluid which is heated by the reactor and exhausted. Another energy-limited rocket system is the hydrogen/oxygen rocket where H2/O2 fuel is produced by electrolysis of water using a solar array or a nuclear reactor. The problem is to choose the optimum specific impulse (or equivalently exhaust velocity) to minimize the amount of energy required to produce a given mission delta-v in the payload. The somewhat surprising result is that the optimum specific impulse is not the maximum possible value, but is proportional to the mission delta-v. In general terms, at the beginning of the mission it is optimum to use a very low specific impulse and expend a lot of reaction mass, since this is the most energy efficient way to transfer momentum. However, as the mission progresses, it becomes important to minimize the amount of reaction mass expelled, since energy is wasted moving the reaction mass. Thus, the optimum specific impulse will increase with the mission delta-v. Optimum I(sub sp) is derived for maximum payload return per energy expended for both the case of fixed and variable I(sub sp) engines. Sample missions analyzed include return of water payloads from the moons of Mars and of