These are representative sample records from Science.gov related to your search topic.
For comprehensive and current results, perform a real-time search at Science.gov.
1

Small rocket exhaust plume data  

NASA Technical Reports Server (NTRS)

During recent cryodeposit tests with an 0.18-N thruster, the mass flux in the plume back field was measured for the first time for nitrogen, carbon dioxide, and a mixture of nitrogen, hydrogen, and ammonia at various inlet pressures. This mixture simulated gases that would be generated by a hydrazine plenum attitude propulsion system. The measurements furnish a base upon which to build a mathematical model of plume back flow that will be used in predicting the mass distribution in the boundary region of other plumes. The results are analyzed and compared with existing analytical predictions.

Chirivella, J. E.; Moynihan, P. I.; Simon, W.

1972-01-01

2

Implementation of microwave transmissions for rocket exhaust plume diagnostics  

NASA Astrophysics Data System (ADS)

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

Coutu, Nicholas George

3

Environmental Effects of Space Shuttle Solid Rocket Motor Exhaust Plumes  

NASA Technical Reports Server (NTRS)

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

Hwang, B.; Pergament, H. S.

1976-01-01

4

Temperature Measurement of Solid Rocket Motor Exhaust Plume by Absorption-Emission SPECTROSCOPY1  

Microsoft Academic Search

A method, which measures the temperature of solid rocket motor exhaust plume, was developed by employing an improved sodium line reversal process. The formula for calculating the temperature was improved and simplified. The temporal temperature-time distributions of the exhaust plume of double base propellant rocket motors were given by the established method. The maximum time resolution and accuracy for the

Dong Yang; Houqian Xu; Junde Wang; Baochang Zhao

2001-01-01

5

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

NASA Technical Reports Server (NTRS)

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.

Chirivella, J. E.

1975-01-01

6

Retro rocket plume actuated heat shield exhaust ports  

Microsoft Academic Search

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 port

Colleen Marrese-Reading; J. St. Vaughn; R. Frisbee; P. Zell; K. Hamm; J. Corliss; S. Gayle; R. Pain; D. Rooney; A. Ramos; D. Lewis; J. Shepherd; K. Inaba

2009-01-01

7

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

NASA Technical Reports Server (NTRS)

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

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

1975-01-01

8

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

NASA Technical Reports Server (NTRS)

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

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

2008-01-01

9

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

NASA Technical Reports Server (NTRS)

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).

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

1984-01-01

10

Analysis of mechanisms and the nature of radiation from aluminum oxide in different phase states in solid rocket exhaust plumes  

Microsoft Academic Search

We discuss issues of mechanisms and the nature of Al2O3 radiation in the plume of solid-propellant rocket motors. It was disclosed that the Al2O3 optical constants responsible for liquid and solid particle radiation in the plume can be specified from a single point of view in terms of liquid semiconductor properties with the zone structure having a forbidden bandwidth around

N. A. Anfimov; G. F. Karabadyak; B. A. Khmelinin; Y. A. Plastinin; A. V. Rodionov

1993-01-01

11

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

NASA Technical Reports Server (NTRS)

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.

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

2008-01-01

12

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)

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.

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

2000-01-01

13

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)

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.

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

2000-01-01

14

Hybrid plume plasma rocket  

NASA Technical Reports Server (NTRS)

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 wall 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.

Chang, Franklin R. (inventor)

1989-01-01

15

Particle behavior in solid propellant rocket motors and plumes  

Microsoft Academic Search

The particle size distribution inside the combustion chamber and the changes that occurred across the exhaust nozzle were measured in a subscale solid propellant rocket motor with a 2 percent aluminized end-burning propellant grain and a highly underexpanded nozzle. A combination of diagnostic techniques were used. Size distributions in the exhaust plume were determined by a Single Particle Counter, a

John D. McCrorie II

1992-01-01

16

Empirical Scaling Laws of Rocket Exhaust Cratering  

NASA Technical Reports Server (NTRS)

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.

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

2005-01-01

17

Frequency-Dependent FDTD Simulation of the Interaction of Microwaves With Rocket-Plume  

Microsoft Academic Search

The ionized exhaust plumes of solid rocket motors may interfere with RF transmission under certain flight conditions. To understand the important physical processes involved, we measured microwave attenuation and phase delay due to the exhaust plume during sea-level static firing tests for a full-scale solid propellant rocket motor. The measured data were compared with the results of a detailed simulation

Kiyoshi Kinefuchi; Ikkoh Funaki; Takashi Abe

2010-01-01

18

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

E-print Network

than the surrounding atmosphere. When the air flow around the rocket separates from the missile, the pressure distribution along the rocket changes, and thrust efficiency and fin effectiveness are reduced. Numerical Simulations of Plume Interactions... difference in pressure causes large exhaust plumes and strong interactions between the external flow and plume. The interactions occurring between the missile's expanding exhaust plume and the air passing the side of the missile can cause changes...

Lawton, Stephen Hayes

1989-01-01

19

Gasdynamic propagation of rocket exhaust products in the upper atmosphere  

NASA Astrophysics Data System (ADS)

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.

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

2011-12-01

20

Exhaust Nozzle Plume and Shock Wave Interaction  

NASA Technical Reports Server (NTRS)

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

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

2013-01-01

21

Radiation from advanced solid rocket motor plumes  

NASA Technical Reports Server (NTRS)

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.

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

1994-01-01

22

DESIGN CRITERIA FOR ROCKET EXHAUST SCRUBBERS  

EPA Science Inventory

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

23

An expert system for spectroscopic analysis of rocket engine plumes  

NASA Technical Reports Server (NTRS)

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

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

1991-01-01

24

Evaluation of computation codes for rocket plume's infrared signature by using measurements on a small scale aluminized composite propellant motor  

Microsoft Academic Search

A new experiment has been conducted with a composite propellant rocket motor in order to get two kinds of information: first one is concerning the physical and optical properties of aluminates particles that are emitted in the plume exhaust; second one is concerning the spectral and spatial repartition of radiance in the plume infrared images. The size distribution and the

A. Boischot; A. Roblin; L. Hespel; I. Dubois; P. Prevot; T. Smithson

2006-01-01

25

Rocket exhaust ground cloud/atmospheric interactions  

NASA Technical Reports Server (NTRS)

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

Hwang, B.; Gould, R. K.

1978-01-01

26

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

NASA Technical Reports Server (NTRS)

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.

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

1976-01-01

27

Exhaust gas composition measurement. [liquid monopropellant rocket engine performance tests  

NASA Technical Reports Server (NTRS)

The design, installation, checkout, and operation of an exhaust gas composition measurement system for collecting and analyzing the exhaust gas from a liquid monopropellant rocket engine are described. Design guidelines are given for the critical components of each portion of the system to provide an exhaust gas composition measurement which meets the performance criteria specified.

1979-01-01

28

Combustion Model of Supersonic Rocket Exhausts in an Entrained Flow Enclosure  

NASA Technical Reports Server (NTRS)

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.

Vu, Bruce; Oliveira, Justin

2011-01-01

29

Flow and Combustion in the Base-Wall Region of a Rocket Exhaust Plums  

Microsoft Academic Search

A model for predicting physical characteristics of combustion in the recirculating, chemically-reacting flow in the base region of a rocket exhaust plume is described. The nozzle jet and the free stream are considered as supersonic or subsonic, mixing turbulently to form an axisymmetric and compressible free-boundary layer. The turbulence is accounted for by a two-equation model, which solves the transport

N. C. MARKATOS; D. B. SPALDING; D. G. TATCHELL; A. C. H. MACE

1982-01-01

30

Monopropellant thruster exhaust plume contamination measurements  

NASA Technical Reports Server (NTRS)

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

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

1977-01-01

31

Rocket-Plume Spectroscopy Simulation for Hydrocarbon-Fueled Rocket Engines  

NASA Technical Reports Server (NTRS)

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 as an impurity in the propellants and/or these can form in the boundary layer as a result of interaction of the hot plume with the atmosphere during the ground testing of engines. Ten additional electronic band systems of these five molecules have been included into the code. A comprehensive literature search was conducted to obtain the most accurate values for the molecular and the spectral parameters, including Franck-Cordon factors and electronic transition moments for all ten band systems. For each elemental transition in the RPSSC, six spectral parameters - Doppler broadened line width at half-height, pressure-broadened line width at half-height, electronic multiplicity of the upper state, electronic term energy of the upper state, Einstein transition probability coefficient, and the atomic line center - are required. Input files have been created for ten elements of Ni, Fe, Cr, Co, Cu, Ca, Mn, Al, Ag, and Pd, which retain only relatively moderate to strong transitions in 300 to 430 nm spectral range for each element. The number of transitions in the input files is 68 for Ni; 148 for Fe; 6 for Cr; 87 for Co; 1 for Ca; 3 for Mn; 2 each for Cu, Al, and Ag; and 11 for Pd.

Tejwani, Gopal D.

2010-01-01

32

Rocket plume flowfield characterization using laser Rayleigh scattering  

NASA Technical Reports Server (NTRS)

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.

Zupanc, Frank J.; Weiss, Jonathan M.

1992-01-01

33

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

NASA Astrophysics Data System (ADS)

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

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

2013-09-01

34

Instrumentation for In-Flight SSME Rocket Engine Plume Spectroscopy  

NASA Technical Reports Server (NTRS)

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.

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

1994-01-01

35

Instrumentation for in-flight SSME rocket engine plume spectroscopy  

NASA Astrophysics Data System (ADS)

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.

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

1994-06-01

36

Submillimeter-wave properties of thermospheric rocket plumes  

NASA Technical Reports Server (NTRS)

The problem of detecting rocket plumes at thermospheric altitudes with satellite-borne submillimeter-wave radiometers is examined theoretically. To estimate the sizes of plume signatures contrasted against a 250-K earth background or in self-emission against the cold sky, a computer program has been developed to predict plume brightness temperatures and optical depths of rotational lines of plume molecular constituents (e.g., H2O) as a function of distance from the nozzle. The methods employed in the computations are described in general terms, and examples are presented to indicate that detectable H2O signatures extending to several thousand nozzle diameters should exist at plume altitudes above 250 km.

Litvak, M. M.; Weiss, J. A.; Dionne, G. F.

1980-01-01

37

Effect of water to ablative performance under solid rocket exhaust environment  

NASA Astrophysics Data System (ADS)

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

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

1993-06-01

38

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

NASA Astrophysics Data System (ADS)

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

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

2008-12-01

39

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

NASA Technical Reports Server (NTRS)

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

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

2004-01-01

40

Rocket Exhaust Cratering: A Significant Challenge for Space Exploration  

NASA Technical Reports Server (NTRS)

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

Metzger, Philip

2005-01-01

41

Atmospheric scavenging of hydrochloric acid. [from rocket exhaust  

NASA Technical Reports Server (NTRS)

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

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

1975-01-01

42

Improvement of Rocket Engine Plume Analysis Techniques  

NASA Technical Reports Server (NTRS)

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.

Smith, S. D.

1982-01-01

43

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

NASA Technical Reports Server (NTRS)

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

Srivastava, Ashok N.; Buntine, Wray

1995-01-01

44

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

NASA Astrophysics Data System (ADS)

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

Srivastava, Ashok N.; Buntine, Wray

45

Numerical Simulation of Rocket Exhaust Interaction with Lunar Soil  

NASA Technical Reports Server (NTRS)

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

Liever, Peter; Tosh, Abhijit; Curtis, Jennifer

2012-01-01

46

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

NASA Astrophysics Data System (ADS)

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

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

2012-10-01

47

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

SciTech Connect

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

Kinefuchi, K. [Department of Aeronautics and Astronautics, University of Tokyo, 7-3-1, Hongo, Bunkyo-ku, Tokyo 113-8656 (Japan); Funaki, I.; Shimada, T.; Abe, T. [Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, 3-1-1, Yoshinodai, Chuo-ku, Sagamihara, Kanagawa 252-5210 (Japan)

2012-10-15

48

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

NASA Technical Reports Server (NTRS)

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.

Wang, K. C.

1993-01-01

49

ASRM subscale plume deflector testing. [advanced solid rocket motor  

NASA Technical Reports Server (NTRS)

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

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

1992-01-01

50

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

NASA Technical Reports Server (NTRS)

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.

Penny, M. M.

1975-01-01

51

Detection of Metallic Compounds in Rocket Plumes Using Ion Probes  

NASA Technical Reports Server (NTRS)

This grant experimentally verified that ion probes can consistently detect metallic compounds in a hybrid rocket plume. Two electrostatic detection methods were tested. The first method used an unbiased ion probe. It responded to collisions or near collisions with charged particulates. The amplitude of the response to metallic ions always exceeded that of the combustion products. The second device was a cylindrical Gaussian surface that surrounded, but did not touch, the plume. A charge imbalance in the plume induced a current in cylinder that was detected by a sensitive amplifier. The probe was more sensitive to metallic compounds than the cylinder. However, the Gaussian cylinder demonstrated sufficient sensitivity to warrant serious future consideration. Since the cylinder is nonintrusive, it is particularly attractive. Apparently, ions formed during combustion transfer to the metallic impurities. The formation of these metallic ions slows the ion recombination rate and helps preserve charges in the plume. The electrostatic detectors, in turn, respond to the charges carried by the metallic impurities.

Dunn, Robert W.

1998-01-01

52

Studies of the exhaust products from solid propellant rocket motors  

NASA Technical Reports Server (NTRS)

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.

Dawbarn, R.; Kinslow, M.

1976-01-01

53

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

NASA Technical Reports Server (NTRS)

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.

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

1993-01-01

54

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

NASA Technical Reports Server (NTRS)

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.

Baran, W. J.

1974-01-01

55

Vortex wake and exhaust plume interaction, including ground effect  

NASA Astrophysics Data System (ADS)

Computational modeling and studies of the near-field wake-vortex turbulent flows, far-field turbulent wake- vortex/exhaust-plume interaction for subsonic and High Speed Civil Transport (HSCT) airplane, and wake- vortex/exhaust-plume interaction with the ground are carried out. The three-dimensional, compressible Reynolds-Averaged Navier-Stokes (RANS) equations are solved using the implicit, upwind, Roe-flux-differencing, finite-volume scheme. The turbulence models of Baldwin and Lomax, one-equation model of Spalart and Allmaras and two-equation shear stress transport model of Menter are implemented with the RANS solver for turbulent-flow modeling. For the near-field study, computations are carried out on a fine grid for a rectangular wing with a NACA-0012 airfoil section and a rounded tip. The focus of study is the tip-vortex development, the near-wake-vortex roll-up, and validation of the results with the available experimental data. For the far-field study, the computations of wake-vortex interaction with the exhaust-plume of a single engine of a medium-size subsonic aircraft in a holding condition and two engines of a HSCT in a cruise condition are carried out using an overlapping zonal method for several miles downstream. The overlapping zonal method has been carefully developed and investigated for accurate and efficient calculations of the far-field wake-vortex flow. The results of the subsonic flow are compared with those of a Parabolized Navier-Stokes (PNS) solver known as the UNIWAKE code. Next, the problem of wake-vortex/ground interaction is investigated. For the simulation of this problem, typical velocity profiles of a tip vortex with and without the exhaust-plume temperature profiles are used for inflow boundary conditions and the computations are carried out using the overlapping zonal method for long distances downstream. The effects of the exhaust-plume temperature on the vortex descent, ground boundary-layer separation, vortex rebound and vortex decay are studied and validated with the available experimental data. A parametric study, which covers the effects of atmospheric conditions such as axial wind, crosswind, wind shear, turbulence and, Reynolds number on vortex motion and dynamics near the ground, is also carried out.

Adam, Ihab Gaber

56

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

NASA Technical Reports Server (NTRS)

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

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

1990-01-01

57

Hyper-spectral imaging of aircraft exhaust plumes  

NASA Astrophysics Data System (ADS)

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.

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

2008-10-01

58

Nuclear thermal rocket plume interactions with spacecraft. Final report  

SciTech Connect

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

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

1997-05-01

59

Polar Mesospheric Clouds and Rocket Exhaust in the Arctic Middle Atmosphere: Lidar Observations and Analysis  

NASA Astrophysics Data System (ADS)

We report observations of polar mesospheric clouds (PMC) and rocket exhaust by ground-based lidar at Poker Flat Research Range (PFRR), Chatanika, Alaska (65°N, 147°W). The PMC observations have been made in late summer over several years in years when space shuttle launches both did and did not occur. The rocket exhaust observations have been made in late winter and spring on three nights when Black Brandt XII (two) and X (one) rockets were launched at PFRR. The PMCs are observed at altitudes between 80 and 86 km during visual displays. The rocket exhaust is observed at altitudes between 66 and 82 km, with the strongest echoes from the cloud at the higher altitudes. The aerosol backscatter ratios of the clouds and exhaust have magnitudes from 0.1 to 100. We consider the areal extent, seasonal evolution, and environmental conditions of the PMCs as observed by satellites (i.e., EOS-Aura/OMI, NOAA/SBUV, and EOS-Aura/MLS). We analyze the structure of the PMCs in different years in terms of current microphysical models and analyze the characteristics of the clouds in terms of the influence of space shuttle exhaust. We consider the formation of the rocket exhaust in terms of the combustion products of the rocket fuel and the environmental conditions measured by satellites (i.e., UARS/MLS). We compare and contrast the structure of the PMCs and rocket exhaust and discuss them as indicators of atmospheric conditions.

Collins, R. L.; Deland, M. T.; Lieberman, R. S.; Walker, G. W.

2010-12-01

60

Active chlorine and nitric oxide formation from chemical rocket plume afterburning  

NASA Technical Reports Server (NTRS)

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

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

1994-01-01

61

PREDICTING ENGINE PARAMETERS USING THE OPTICAL SPECTRUM OF THE SPACE SHUTTLE MAIN ENGINE EXHAUST PLUME  

Microsoft Academic Search

The Optical Plume Anomaly Detection (OPAD) sys- tem is under development to predict engine anoma- lies and engine parameters of the Space Shuttle's Main Engine (SSME). The anomaly detection is based on abnormal metal concentrations in the op- tical spectrum of the rocket plume. Such abnor- malities could be indicative of engine corrosion or other malfunctions. Here, we focus on

Ashok N. Srivastava; Wray Buntine

62

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

Microsoft Academic Search

The Optical Plume Anomaly Detection (OPAD) system is under development to predict engine anomalies and engine parameters of the Space Shuttle's Main Engine (SSME). The anomaly detection is based on abnormal metal concentrations in the optical spectrum of the rocket plume. Such abnormalities could be indicative of engine corrosion or other malfunctions. Here, we focus on the second task of

Ashok N. Srivastava; Wray Buntine

1995-01-01

63

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

Microsoft Academic Search

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

P. R. Ryason

1975-01-01

64

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)

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.

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

65

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

NASA Technical Reports Server (NTRS)

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

Everson, John; Nelson, H. F.

1993-01-01

66

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

NASA Technical Reports Server (NTRS)

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.

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

67

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

NASA Astrophysics Data System (ADS)

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.

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

1991-12-01

68

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

NASA Technical Reports Server (NTRS)

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

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

1992-01-01

69

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

NASA Technical Reports Server (NTRS)

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

Susko, M.

1979-01-01

70

Analyzing rocket plume spectral data with neural networks  

NASA Astrophysics Data System (ADS)

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

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

71

Analyzing rocket plume spectral data with neural networks  

SciTech Connect

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

Whitaker, K.W.; Krishnakumar, K.S.; Benzing, D.A.

1995-09-01

72

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

NASA Technical Reports Server (NTRS)

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

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

1982-01-01

73

The effect of searchlight emission on radiation from solid rocket plumes  

Microsoft Academic Search

Measurements of exhaust plume radiation from aluminized solid-propellant motors indicate that both the intensity and spatial distribution of this radiation are confounded by searchlight emission. Failure to account for this can lead to significant errors in the predicted in-flight radiative heat transfer to spacecraft. Several techniques for the identification and quantification of searchlight radiation were attempted. Comparison of the exit

R. A. Reed; K. S. Beale; D. F. Frazine; D. W. Neese; F. G. Sherrell; D. W. Roberds; S. M. Oliver

1992-01-01

74

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

NASA Technical Reports Server (NTRS)

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.

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

1993-01-01

75

The AlCl absorption feature in solid rocket plume radiation  

Microsoft Academic Search

An absorption feature in the mid-UV spectral region of the radiation from aluminized solid-propellent rocket plumes has been identified as the X 1 Sigma - A 1 Pi transition of aluminum chloride. A proposed model of the radiative transfer has the continuum radiation from the aluminum oxide particulate core passing through the outer gas boundary. A line-by-line calculation is then

W. K. McGregor; J. A. Drakes; K. S. Beale; F. G. Sherrell

1992-01-01

76

The Role of HOx in Super-and Subsonic Aircraft Exhaust Plumes  

NASA Technical Reports Server (NTRS)

The generation of sulfuric acid aerosols in aircraft exhaust has emerged as a critical issue in determining the impact of supersonic aircraft on stratospheric ozone. It has long been held that the first step in the mechanism of aerosol formation is the oxidation of SO2 emitted from the engine by OH in the exhaust plume. We report in situ measurements of OH and HO2 in the exhaust plumes of a supersonic (Air France Concorde) and a subsonic (NASA ER-2) aircraft in the lower stratosphere. These measurements imply that reactions with OH are responsible for oxidizing only a small fraction of SO2 (2%), and thus cannot explain the large number of particles observed in the exhaust wake of the Concorde.

Hanisco, T. F.; Wennberg, P. O.; Cohen, R. C.; Anderson, J. G.; Fahey, D. W.; Keim, E. R.; Gao, R. S.; Wamsley, R. C.; Donnelly, S. G.; DelNegro, L. A.

1997-01-01

77

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

NASA Technical Reports Server (NTRS)

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.

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

78

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

NASA Astrophysics Data System (ADS)

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

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

2003-09-01

79

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

NASA Technical Reports Server (NTRS)

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.

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

1980-01-01

80

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

NASA Astrophysics Data System (ADS)

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

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

2014-01-01

81

Rocket exhaust effluent modeling for tropospheric air quality and environmental assessments  

NASA Technical Reports Server (NTRS)

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.

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

1977-01-01

82

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

NASA Technical Reports Server (NTRS)

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

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

1977-01-01

83

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

NASA Technical Reports Server (NTRS)

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

Anderson, Bruce

2004-01-01

84

A nonintrusive method for the measurement of infrared characteristics from engine exhaust plume  

NASA Astrophysics Data System (ADS)

Nonintrusive measurements of infrared characteristics from engine exhaust plume are required for emission control or target tracking, due to the advantage of online measurement without affecting the exhaust plume. Conventional nonintrusive measurement techniques, e.g. the passive Fourier-transform infrared (FTIR) absorption spectrometry, lack prior knowledge of backgrounds and consume time to measure the complete infrared characteristics. Hence, an improved but simple nonintrusive method is proposed. Accordingly, a prototype system with a Mid-wave infrared imager has been developed and tested for the measurement of vehicle engine exhaust plume. Subsequently, the time-variant effective transmittance and emissivity is determined. Compared to the passive FTIR absorption spectrometry, this method incorporates a known background into the measurement and is more adequate for recording the rapidly changing exhaust plume radiation. Therefore, the accurate value of the transmittance and emissivity can be obtained. Further analysis reveals that the imager could be replaced with a dispersive spectrometer, which makes it feasible to acquire the absolute transmittance and emissivity with respect to wavelength. Thus, the concentration of specific toxic gases could be calculated following the radiance inversion technique.

Xiao, Xizhong; Wang, Yueming; Miao, Bin; Lang, Junwei; Wang, Shengwei; Zhuang, Xiaoqiong; Zhou, Feng; Wang, Jianyu

2013-12-01

85

Stratospheric aircraft exhaust plume and wake chemistry studies  

NASA Technical Reports Server (NTRS)

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

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

1992-01-01

86

Nozzle installation effects on the noise from supersonic exhaust plumes  

NASA Astrophysics Data System (ADS)

The sensitivity of screech coupling in supersonic jets to nozzle installation geometry is explored as a function of nozzle shape, spacing, and orientation. The coupling phenomenon is shown to be a function of geometry for a variety of twin axisymmetric and rectangular nozzle configurations as well as for a single jet in proximity to a solid surface. Rapid plume merging or close proximity to a wall are shown to minimize the noise increment due to coupling. Twin impinging supersonic plumes experience more complex aeroacoustic interactions. The acoustic near field is dominated by screech and impingement tones, but the fuselage undersurface dynamic loads are primarily due to impingement of the unsteady upwash fountain flow on the fuselage undersurface.

Wlezien, R. W.

1992-04-01

87

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

NASA Technical Reports Server (NTRS)

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.

Pearce, B. E.

1983-01-01

88

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

NASA Technical Reports Server (NTRS)

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

Roberts, F. E., III

1991-01-01

89

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

NASA Technical Reports Server (NTRS)

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

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

1995-01-01

90

An overview of in-flight plume diagnostics for rocket engines  

NASA Technical Reports Server (NTRS)

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

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

1992-01-01

91

Crew Launch Vehicle Mobile Launcher Solid Rocket Motor Plume Induced Environment  

NASA Technical Reports Server (NTRS)

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.

Vu, Bruce T.; Sulyma, Peter

2008-01-01

92

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

NASA Astrophysics Data System (ADS)

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 2007 formed bright PMCs between 75 and 85°N a day after launch. These are the highest latitude Arctic PMCs formed by shuttle exhaust reported to date.

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

2014-02-01

93

Numerical investigation of plane plume exhausting from wedge-like micronozzle  

NASA Astrophysics Data System (ADS)

Numerical simulations of near field of a plume exhausting from a plane wedge-like micronozzle into vacuum are performed using two different kinetic approaches: one of the model kinetic equations (ellipsoid statistical model) and Direct Simulation Monte Carlo method. Adequacy and accuracy of the model kinetic equation as applied to such strongly non-equilibrium flows are studied by comparing the results with the data of DSMC simulations.

Shershnev, A. A.; Kudryavtsev, A. N.

2014-12-01

94

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

SciTech Connect

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

Jenkins, R.A.

1990-06-07

95

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

NASA Technical Reports Server (NTRS)

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.

Castner, Raymond; Lake, Troy

2012-01-01

96

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

NASA Technical Reports Server (NTRS)

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.

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

1991-01-01

97

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

NASA Technical Reports Server (NTRS)

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

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

1982-01-01

98

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

NASA Technical Reports Server (NTRS)

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 (that employed parabolic or temporal approximations) and the measured data. Finally to address the major uncertainty in the near-field plume modeling related to the plume processing of sulfur compounds and advanced model was developed to evaluate its impact on the chemical processes in the near wake. A comprehensive aerosol model is developed and it is coupled with chemical kinetics and the axisymmetric turbulent jet flow models. The integrated model is used to simulate microphysical processes in the near-field jet plume, including sulfuric acid and water binary homogeneous nucleation, coagulation, non-equilibrium heteromolecular condensation, and sulfur-induced soot activation. The formation and evolution of aerosols are computed and analyzed. The computed results show that a large number of ultra-fine (0.3--0.6 nm in radius) volatile HSO4 - HO embryos are generated in the near-field plume. These embryos further grow in size by self coagulation and condensation. Soot particles can be activated by both heterogeneous nucleation and scavenging of H2SO4-H2O aerosols. These activated soot particles can serve as water condensation nuclei for contrail formation. Conditions under which ice contrails can form behind aircrafts are studied. The sensitivities of the threshold temperature for contrail formation with respect to aircraft propulsion efficiency, relative humidity, and ambient pressure are evaluated. The computed aerosol properties for different extent of fuel sulfur conversion to S(VI) (SO3 and H2SO4) in engine are examined and the results are found to be sensitive to this conversion fraction.

Menon, Suresh

1998-01-01

99

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

NASA Technical Reports Server (NTRS)

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

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

1971-01-01

100

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

NASA Technical Reports Server (NTRS)

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

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

1972-01-01

101

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

NASA Technical Reports Server (NTRS)

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.

Pellett, G. L.

1981-01-01

102

Dynamic Analysis of a Building Under Rocket Engine Plume Acoustic Load  

NASA Technical Reports Server (NTRS)

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

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

2001-01-01

103

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

NASA Technical Reports Server (NTRS)

The base pressure environment was investigated for the first and second stage mated vehicle in a supersonic flow field from Mach 1.55 through 2.20 with simulated rocket engine exhaust plumes. The pressure environment was investigated for the orbiter at various vent port locations at these same freestream conditions. The Mach number environment around the base of the model with rocket plumes simulated was examined. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of the tested configurations are shown.

Hawthorne, P. J.

1976-01-01

104

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

NASA Technical Reports Server (NTRS)

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.

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

1973-01-01

105

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

NASA Technical Reports Server (NTRS)

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

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

2003-01-01

106

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

NASA Technical Reports Server (NTRS)

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.

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

1990-01-01

107

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

SciTech Connect

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.

Chacko, C.C.; Mendillo, M.

1980-09-01

108

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)

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.

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

2014-01-01

109

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)

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.

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

1970-01-01

110

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

NASA Astrophysics Data System (ADS)

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

Brandenburg, John; Behringer, Robert; Clarke, Abraham

2009-11-01

111

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

NASA Technical Reports Server (NTRS)

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

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

1981-01-01

112

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

NASA Astrophysics Data System (ADS)

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

Wang, Yan Jason; Zhang, K. Max

2012-11-01

113

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)

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.

Susko, M.

1976-01-01

114

Study of high altitude plume impingement  

NASA Technical Reports Server (NTRS)

The radiation intensities are determined in the base region of the space shuttle due to solid particle radiation emanating from the solid rocket motors of the shuttle. Results of an analysis of the Titan 3 and simulated solid rocket motor radiation intensities are presented. The gas particle flow fields of the Titan 3 nozzle and plume and a space shuttle solid rocket motor nozzle and plume are described. The gaseous Titan 3 flow fields are discussed utilizing the results of flow fields generated by a gaseous and two phase method-of-characteristics computer programs. A two phase computer flow field analysis program was developed. An outflow correction theory is developed which will be used to modify existing convection heat transfer methods for better heat transfer predictions on bodies immersed in rocket exhaust plumes.

Mcanally, J. V.; Smith, S. D.

1973-01-01

115

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

NASA Technical Reports Server (NTRS)

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.

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

2011-01-01

116

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

NASA Technical Reports Server (NTRS)

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

Bruckner, A. P.

1972-01-01

117

OPAD data analysis. [Optical Plumes Anomaly Detection  

NASA Technical Reports Server (NTRS)

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

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

1993-01-01

118

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

NASA Technical Reports Server (NTRS)

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

Powers, W. T.

1989-01-01

119

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

NASA Astrophysics Data System (ADS)

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 ?? = 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.

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

120

Prediction of the Size of Aluminum-Oxide Particles in Exhaust Plumes of Solid Rocket Motors  

Microsoft Academic Search

The processes of coagulation and aerodynamic fragmentation of liquid particles of aluminum oxide in an accelerating gas flow in the Laval nozzle are analyzed. A formula obtained by an approximate analytical solution of equations of a two-phase flow is proposed to calculate the characteristic particle diameter at the nozzle exit. The limiting particle diameter in the nozzle throat calculated theoretically

O. B. Kovalev

2002-01-01

121

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

NASA Technical Reports Server (NTRS)

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

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

1980-01-01

122

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)

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.

Smith, S. D.

1984-01-01

123

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

NASA Technical Reports Server (NTRS)

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

Hawthorne, P. J.

1976-01-01

124

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

NASA Technical Reports Server (NTRS)

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.

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

1973-01-01

125

On the fast zonal transport of the STS-121 space shuttle exhaust plume in the lower thermosphere  

NASA Astrophysics Data System (ADS)

Meier et al. (2011) reported rapid eastward transport of the STS-121 space shuttle (launch: July 4, 2006) main engine plume in the lower thermosphere, observed in hydrogen Lyman ? images by the GUVI instrument onboard the TIMED satellite. In order to study the mechanism of the rapid zonal transport, diagnostic tracer calculations are performed using winds from the Thermosphere Ionosphere Mesosphere Electrodynamics General Circulation Model (TIME-GCM) simulation of July, 2006. It is found that the strong eastward jet at heights of 100-110 km, where the exhaust plume was deposited, results in a persistent eastward tracer motion with an average velocity of 45 m/s. This is generally consistent with, though faster than, the prevailing eastward shuttle plume movement with daily mean velocity of 30 m/s deduced from the STS-121 GUVI observation. The quasi-two-day wave (QTDW) was not included in the numerical simulation because it was found not to be large. Its absence, however, might be partially responsible for insufficient meridional transport to move the tracers away from the fast jet in the simulation. The current study and our model results from Yue and Liu (2010) explain two very different shuttle plume transport scenarios (STS-121 and STS-107 (launch: January 16, 2003), respectively): we conclude that lower thermospheric dynamics is sufficient to account for both very fast zonal motion (zonal jet in the case of STS-121) and very fast meridional motion to polar regions (large QTDW in the case of STS-107).

Yue, Jia; Liu, Han-Li; Meier, R. R.; Chang, Loren; Gu, Sheng-Yang; Russell, James, III

2013-03-01

126

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

NASA Technical Reports Server (NTRS)

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

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

1976-01-01

127

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

E-print Network

takeoff, cruise, and landing. After 10 h, the plume treatment decreased grid-scale ozone production by 33 the impact of modeling photochemistry from aircraft emissions in an expanding plume versus at the grid scale in an atmospheric model. Differences in model treatments for a single flight occurred at all altitudes during

Jacobson, Mark

128

Dilution and aerosol dynamics within a diesel car exhaust plume—CFD simulations of on-road measurement conditions  

NASA Astrophysics Data System (ADS)

Vehicle particle emissions are studied extensively because of their health effects, contribution to ambient PM levels and possible impact on climate. The aim of this work was to obtain a better understanding of secondary particle formation and growth in a diluting vehicle exhaust plume using 3-d information of simulations together with measurements. Detailed coupled computational fluid dynamics (CFD) and aerosol dynamics simulations have been conducted for H 2SO 4-H 2O and soot particles based on measurements within a vehicle exhaust plume under real conditions on public roads. Turbulent diffusion of soot and nucleation particles is responsible for the measured decrease of number concentrations within the diesel car exhaust plume and decreases coagulation rates. Particle size distribution measurements at 0.45 and 0.9 m distance to the tailpipe indicate a consistent soot mode (particle diameter Dp˜50 nm) at variable operating conditions. Soot mode number concentrations reached up to 10 13 m -3 depending on operating conditions and mixing. For nucleation particles the simulations showed a strong sensitivity to the spatial dilution pattern, related cooling and exhaust H 2SO 4(g). The highest simulated nucleation rates were about 0.05-0.1 m from the axis of the plume. The simulated particle number concentration pattern is in approximate accordance with measured concentrations, along the jet centreline and 0.45 and 0.9 m from the tailpipe. Although the test car was run with ultralow sulphur fuel, high nucleation particle ( Dp?15 nm) concentrations (>10 13 m -3) were measured under driving conditions of strong acceleration or the combination of high vehicle speed (>140 km h -1) and high engine rotational speed (>3800 revolutions per minute (rpm)). Strong mixing and cooling caused rapid nucleation immediately behind the tailpipe, so that the highest particle number concentrations were recorded at a distance, x=0.45 m behind the tailpipe. The simulated growth of H 2SO 4-H 2O nucleation particles was unrealistically low compared with measurements. The possible role of low and semi-volatile organic components on the growth processes is discussed. Simulations for simplified H 2SO 4-H 2O-octane-gasoil aerosol resulted in sufficient growth of nucleation particles.

Uhrner, U.; von Löwis, S.; Vehkamäki, H.; Wehner, B.; Bräsel, S.; Hermann, M.; Stratmann, F.; Kulmala, M.; Wiedensohler, A.

129

Digital filtering of plume emission spectra  

NASA Technical Reports Server (NTRS)

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

Madzsar, George C.

1990-01-01

130

Digital filtering of plume emission spectra  

SciTech Connect

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

Madzsar, G.C.

1990-07-01

131

Ionospheric effects of rocket exhaust products - Skylab and HEAO-C  

NASA Astrophysics Data System (ADS)

A description is presented of a new computer model of the mesosphere, thermosphere, and ionosphere, taking into account a comparison of computed results with experimental data. The model was developed in the course of a study of ionospheric depletion effects of large rocket launches through the F layer. The computer code uses a two-dimensional array of Eulerian mesh cells in Cartesian coordinates x horizontal (in the geomagnetic meridian plane), and z vertical. The code integrates the chemical/photochemical kinetic equations for 30 individual chemical species in each of 315 cells. Solar radiation, scatter UV, cosmic rays, and precipitating electrons are considered. Diffusion rates are computed for each neutral species under the influence of gravity. The midlatitude ionosphere in connection with the launch of Skylab I on May 14, 1973 is considered, and attention is given to rocket launches and two-dimensional model comparisons.

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

132

Ionospheric shock waves triggered by rockets  

NASA Astrophysics Data System (ADS)

This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC) derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK) Taepodong-2 and 2013 South Korea (SK) Korea Space Launch Vehicle-II (KSLV-II) rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100-600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800-1200 m s-1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800-1200 m s-1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.

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

2014-09-01

133

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)

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.

Hilst, G. R.

1974-01-01

134

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

NASA Technical Reports Server (NTRS)

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.

Harris, Michael F.; Vu, Bruce T.

2012-01-01

135

The dominant effect of alumina on nearfield plume radiation  

Microsoft Academic Search

Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Combustion of aluminum produces condensed alumina particles. Besides causing performance losses in the nozzle, the condensed Al2O3 particles are the major source of primary smoke in the exhaust plume. The particulate matter can also have major effects upon the plume i.r. signature. High number

David Laredo; David W. Netzer

1993-01-01

136

An application of model testing for the study of rocket exhaust cloud properties  

NASA Technical Reports Server (NTRS)

The application of the 6.4-percent Shuttle Model Test Facility to the study of the Shuttle exhaust cloud properties is discussed with emphasis on the properties related to the production of the deposition (submillimeter drops composed of an acidic liquid and alumina solids) which occurs with each launch. An analysis of test data suggests that the major fraction of the liquid in the deposition is produced directly through the interaction between the exhaust and the deluge water spray and then modified by rapid scavenging of wet acidic aluminum oxide particles. Based on this conclusion, the possibility arises that the acid in the deposition can be neutralized by addition of a base to the deluge water. Current work with the 6.4-percent model is directed toward verification of this hypothesis.

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

1983-01-01

137

Influence of ejector technique on infrared radiation of the exhaust plume outside rectangular nozzle  

NASA Astrophysics Data System (ADS)

For accurate knowledge of the impact of ejector technique on infrared radiation of the plume, the physical model of the rectangular nozzle is established. The 3-D flow field outside the rectangular nozzle is simulated by numerical method with software Fluent6.3 pre and post the application of ejector technique, then the data of the flow fields, such as temperature, pressure and density and so on, are obtained, and according to the characteristic of the rectangular nozzle plume the computational domain of infrared radiation was established. This paper uses Lorentz linear statistical narrow-band model to calculate the mean absorption coefficient of the plume in the narrow band. Then it uses Finite Volume Method(FVM) to solve the radiation transmission equations in gas medium, and it obtains the total intensity distribution in 3~5?m of the plume radiation pre and post the application of ejector technique. The results shows that the infrared radiant of the rectangular nozzle decreases significantly by 80% after the application of ejector technique.

Feng, Yunsong; Lu, Yuan; Qiao, Ya

2013-09-01

138

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

Microsoft Academic Search

A space shuttle launch deposits 700 t of water in the atmosphere. Some of this water is released into the upper mesosphere and lower thermosphere where it may be directly detected by a limb sounding satellite instrument. We report measurements of water vapour plumes from shuttle launches made by the Microwave Limb Sounder (MLS) on the Aura satellite. Approximately 50%

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

2010-01-01

139

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

Microsoft Academic Search

A space shuttle launch deposits 700 tonnes of water in the atmosphere. Some of this water is released into the upper mesosphere and lower thermosphere where it may be directly detected by a limb sounding satellite instrument. We report measurements of water vapour plumes from shuttle launches made by the Microwave Limb Sounder (MLS) on the Aura satellite. Approximately 50%-65%

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

2011-01-01

140

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

NASA Technical Reports Server (NTRS)

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.

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

1973-01-01

141

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

NASA Technical Reports Server (NTRS)

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

Goswami, Kisholoy

2011-01-01

142

Analysis of reacting flowfields in low-thrust rocket engines and plumes  

NASA Astrophysics Data System (ADS)

The mixing and combustion processes in small gaseous hydrogen-oxygen thrusters and plumes are studied by means of a computational model developed as a general purpose analytic procedure for solving low speed, reacting, internal flowfields. The model includes the full Navier-Stokes equations coupled with species diffusion equations for a hydrogen-oxygen reaction kinetics system as well as the option to use either the k-Epsilon or q-Omega low Reynolds number, two-equation turbulence models. Solution of the governing equations is accomplished by a finite-volume formulation with central-difference spatial discretizations and an explicit, four-stage, Runge Kutta time-integration procedure. The Runge-Kutta scheme appears to provide efficient convergence when applied to the calculation of turbulent, reacting flowfields in these small thrusters. Appropriate boundary conditions are developed to properly model propellant mass flowrates and regenerative wall cooling. The computational method is validated against measured engine performance parameters on a global level, as well as experimentally obtained exit plane and plume flowfield properties on a local level. The model does an excellent job of predicting the measured performance trends of an auxiliary thruster as a function of O/F ratio, although the performance levels are consistently underpredicted by approximately 4 percent. These differences arise because the extent to which the wall coolant layer and combustion gases mix and react is underpredicted. Predictions of velocity components, temperature and species number densities in the near-field plume regions of several low-thrust engines show reasonable agreement with experimental data obtained by two separate laser diagnostic techniques. Discrepancies between the predictions and measurements are primarily due to three-dimensional mixing processes which are not accounted for in the analysis. Both comparisons with experiment and the evident reason for errors in absolute levels of predicted quantities suggest the method should prove valuable for predicting parametric trends for design studies. In addition, issues such as numerical stability, robustness and computational efficiency are addressed. These include the evaluation of a numerically compatible two-equation turbulence model and the implementation of a time-derivative preconditioning method for convergence enhancement of low Mach number, chemically reacting flows.

Weiss, Jonathan Mitchell

143

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

SciTech Connect

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.

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

1995-11-01

144

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)

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.

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

1980-01-01

145

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

SciTech Connect

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

Larson, R.S.

1986-06-01

146

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)

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.

Rhein, R. A.

1973-01-01

147

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

NASA Technical Reports Server (NTRS)

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

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

1983-01-01

148

Low altitude plume impingement handbook  

NASA Technical Reports Server (NTRS)

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

Smith, Sheldon D.

1991-01-01

149

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

NASA Technical Reports Server (NTRS)

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

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

1993-01-01

150

Space Shuttle Plume and Plume Impingement Study  

NASA Technical Reports Server (NTRS)

The extent of the influence of the propulsion system exhaust plumes on the vehicle performance and control characteristics is a complex function of vehicle geometry, propulsion system geometry, engine operating conditions and vehicle flight trajectory were investigated. Analytical support of the plume technology test program was directed at the two latter problem areas: (1) definition of the full-scale exhaust plume characteristics, (2) application of appropriate similarity parameters; and (3) analysis of wind tunnel test data. Verification of the two-phase plume and plume impingement models was directed toward the definition of the full-scale exhaust plume characteristics and the separation motor impingement problem.

Tevepaugh, J. A.; Penny, M. M.

1977-01-01

151

Characteristics of aerosol particles and trace gases in ship exhaust plumes  

NASA Astrophysics Data System (ADS)

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 ?m. 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-sulfur fuel, the chemical submicron aerosol fraction is mainly composed of hydrocarbon-like organic aerosol (HOA) species. These include PAHs that are adsorbed onto the high number of ultrafine particles. Nevertheless, the chemical composition, typical particle sizes as well as emitted gaseous components vary substantially dependent on the engine or ship type, engine operation condition and fuel mixture. This results in cargo vessels compared to tankers, passenger ships and river boats being the largest polluters influencing the Elbe shipping lane areas by high amounts of NOx, SO2, CO2, PAH, BC and ultrafine particulate matter. The tropospheric ozone chemistry in this area is also substantially affected particularly due to the increasing number of Elbe-passing ships. As onshore regions can be influenced by aged shipping plumes, trajectory pathways and transportation times were examined. As a consequence of the plumes' aging, variations of the organic fraction of the mass spectral fingerprints were found. Eyring, V. et al. (2010), Atmospheric Environment, 44, 4735-4771.

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

2011-12-01

152

Plume base flow simulation technology  

NASA Technical Reports Server (NTRS)

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.

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

1983-01-01

153

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

NASA Technical Reports Server (NTRS)

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

Smith, S. D.

1984-01-01

154

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

SciTech Connect

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

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

1991-12-09

155

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

NASA Technical Reports Server (NTRS)

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

Dougherty, N. Sam; Liu, Baw-Lin

1991-01-01

156

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

NASA Technical Reports Server (NTRS)

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

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

1975-01-01

157

One-Dimensional Rocket Launch  

NSDL National Science Digital Library

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

Christian, Wolfgang; Belloni, Mario

2006-01-12

158

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

NASA Technical Reports Server (NTRS)

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

Bowyer, J. M.

1977-01-01

159

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

NASA Technical Reports Server (NTRS)

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

Capo, M. A.; Mickle, R.

1975-01-01

160

A tandem mirror plasma source for hybrid plume plasma studies  

NASA Technical Reports Server (NTRS)

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

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

1985-01-01

161

Ablative Rocket Deflector Testing and Computational Modeling  

NASA Technical Reports Server (NTRS)

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

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

2010-01-01

162

Plume interference with space shuttle range safety signals  

NASA Technical Reports Server (NTRS)

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

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

1979-01-01

163

Instrumentation of UALR labscale hybrid rocket motor  

NASA Astrophysics Data System (ADS)

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

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

2006-05-01

164

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

NASA Technical Reports Server (NTRS)

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.

Hendershot, K. C.

1977-01-01

165

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

NASA Technical Reports Server (NTRS)

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.

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

166

Rocket Engine Clustering and Vehicle Integration as Influenced by Base Thermal Environments  

NASA Technical Reports Server (NTRS)

Clustered rocket engines create severe thermal environments in the base of rocket vehicle stages. Boosters burning hydrocarbon fuels experience severe radiant heating early in flight; as the plumes interact at higher altitudes, convective heating becomes significant. For hydrogen-fueled upper stages radiation is not important, but convective heating is severe during the entire stage operation. Predicted and measured heating rates are discussed. The base region thermal environments of stages with clustered engines present a variety of engine/vehicle interaction problems. Components and structures in the base region, including the rocket engines, cannot survive radiant and convective heating from engine exhausts without such remedies as protective insulation, shielding, air-scooping, and proper disposal of the fuel-rich turbine exhaust gases. Different thermal protection concepts evolve for booster and upper stages due to the differences in ground test and flight environments. Solutions to the engine/vehicle interaction and design integration problems are described.

Hopson, George D.; McAnelly, William B.

1966-01-01

167

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

NASA Astrophysics Data System (ADS)

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.

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

2002-12-01

168

High energy laser transmissions through missile exhaust plumes: the implications for high angle-of-attack (AOA) scenarios  

Microsoft Academic Search

Summary form only given. Performance of new generation joint airborne and seaborne (ABL\\/SeBL) high energy laser (HEL) systems may prove very reliable for the early, mid and late boost-phase missile targeting, especially within an air and sea battle space environment. This paper examines late boost-phase target exhaust reverse flows (which tend to engulf the ascending missiles); angle-of-attack (AOA) coupled with

C. A. Paiva; H. S. Slusher

2002-01-01

169

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

NASA Technical Reports Server (NTRS)

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

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

2005-01-01

170

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)

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.

O'Brien, Brian

2009-05-01

171

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)

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.

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

1977-01-01

172

Space Shuttle Plume Simulation Effect on Aerodynamics  

NASA Technical Reports Server (NTRS)

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.

Hair, L. M.

1978-01-01

173

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)

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.

Falanga, Ralph A.; Janos, Joseph J.

1961-01-01

174

An analytical and experimental investigation of resistojet plumes  

NASA Technical Reports Server (NTRS)

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

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

1987-01-01

175

An analytical and experimental investigation of resistojet plumes  

NASA Technical Reports Server (NTRS)

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

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

1987-01-01

176

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

NASA Astrophysics Data System (ADS)

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.

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

2002-11-01

177

Delta 2 Explosion Plume Analysis Report  

NASA Technical Reports Server (NTRS)

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

Evans, Randolph J.

2000-01-01

178

The Ultraviolet Plume Instrument (UVPI)  

Microsoft Academic Search

The Ultraviolet Plume Instrument (UVPI) was launched aboard the Low-power Atmospheric Compensation Experiment (LACE) satellite on 14 Feb. 1990. Both the spacecraft and the UVPI were sponsored by the Directed Energy Office of the Strategic Defense Initiative Organization. The mission of the UVPI was to obtain radiometrically calibrated images of rocket plumes at high altitude and background image data of

D. M. Horan

1993-01-01

179

Linear Spectral Analysis of Plume Emissions Using an Optical Matrix Processor  

NASA Technical Reports Server (NTRS)

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

Gary, C. K.

1992-01-01

180

Atmospheric scavenging exhaust  

NASA Technical Reports Server (NTRS)

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.

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

1977-01-01

181

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

NASA Technical Reports Server (NTRS)

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

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

2014-01-01

182

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds  

NASA Technical Reports Server (NTRS)

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Huebner, Lawrence D.; Witte, David W.; Andrews, Earl H., Jr.

2004-01-01

183

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

NASA Astrophysics Data System (ADS)

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

Nakashima, Yuki; Heki, Kosuke

2014-07-01

184

Liquid Booster Module (LBM) plume flowfield model  

NASA Technical Reports Server (NTRS)

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.

Smith, S. D.

1981-01-01

185

Measurement and infrared image prediction of a heated exhaust flow  

Microsoft Academic Search

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

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

1994-01-01

186

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

NASA Technical Reports Server (NTRS)

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.

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

2013-01-01

187

SAFE Testing Nuclear Rockets Economically  

NASA Astrophysics Data System (ADS)

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 Rover/NERVA programs in the 1960s, the rocket exhaust can no longer be vented to the open atmosphere. As a consequence, previous studies have examined the feasibility of building a large-scale version of the Nuclear Furnace Scrubber that was demonstrated in 1971. We have investigated an alternative that would deposit the rocket exhaust along with any entrained fission products directly into the ground. The Subsurface Active Filtering of Exhaust, or SAFE, concept would allow variable sized engines to be tested for long times at a modest expense. A system overview, results of preliminary calculations, and cost estimates of proof of concept demonstrations are presented. The results indicate that a nuclear rocket could be tested at the Nevada Test Site for under $20 M.

Howe, Steven D.; Travis, Bryan; Zerkle, David K.

2003-01-01

188

Safe testing nuclear rockets economically  

SciTech Connect

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

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

2002-01-01

189

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

NASA Astrophysics Data System (ADS)

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

Furuya, T.; Heki, K.

2008-03-01

190

Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode  

NASA Technical Reports Server (NTRS)

The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations are being conducted at Marshall Space Flight Center to benchmark the FDNS code for RBCC engine operations for such configurations. The primary fluid physics of interests are the mixing and interaction of the rocket plume and secondary flow, subsequent combustion of the fuel rich rocket exhaust with the secondary flow and combustion of the injected afterburner flow. The CFD results are compared to static pressure along the RBCC duct walls, Raman Spectroscopy specie distribution data at several axial locations, net engine thrust and entrained air for the SLS cases. The CFD results compare reasonably well with the experimental results.

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

2000-01-01

191

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

NASA Technical Reports Server (NTRS)

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

Anderson, William E.

2005-01-01

192

Solid rocket booster thermal radiation model, volume 1  

NASA Technical Reports Server (NTRS)

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

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

1976-01-01

193

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

NASA Technical Reports Server (NTRS)

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

Boyle, W.; Conine, B.

1978-01-01

194

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

NASA Technical Reports Server (NTRS)

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.

Meissner, Dana L.

1993-01-01

195

Rocket Engines  

NSDL National Science Digital Library

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

2011-07-27

196

Rocket Launchers  

NSDL National Science Digital Library

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

Museum, Chicago C.

2010-01-01

197

Methylhydrazinium nitrate. [rocket plume deposit chemistry  

NASA Technical Reports Server (NTRS)

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

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

1983-01-01

198

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines  

NASA Technical Reports Server (NTRS)

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

Morris, Christopher I.

2005-01-01

199

Rain scavenging of solid rocket exhaust clouds  

NASA Technical Reports Server (NTRS)

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

Dingle, A. N.

1978-01-01

200

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

E-print Network

sections of aluminum-oxide rocket exhaust particles are sensitive to slight deviations from their nominal (aluminum-oxide) particles. The mass-mean radii of Al2O3 particles in upper-stage II and III booster rocketJOURNAL OF SPACECRAFT AND ROCKETS Vol. 42, No. 4, July­August 2005 Lidar Backscatter Properties

201

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

NASA Technical Reports Server (NTRS)

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

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

2001-01-01

202

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

NASA Technical Reports Server (NTRS)

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.

Davoudzadeh, Farhad; Liu, Nan-Suey

2003-01-01

203

Balloon Rocket  

NSDL National Science Digital Library

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

Minnesota, Science M.

1995-01-01

204

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

205

Plasma Diagnostics Development for Advanced Rocket Engines  

NASA Astrophysics Data System (ADS)

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

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

2000-10-01

206

Space shuttle exhaust cloud properties  

NASA Technical Reports Server (NTRS)

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

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

1983-01-01

207

Rocket Pinwheel  

NSDL National Science Digital Library

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

2012-06-26

208

Devising rocket power for smaller engines  

SciTech Connect

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).

Burruss, R. [GEOMET Technologies Inc., Germantown, MD (United States)

1996-04-01

209

Automatic Control of Rockets  

Microsoft Academic Search

\\u000a To model a rocket’s attitude plant with gimbaled nozzle and reaction jet actuators. To design and analyze a rocket’s roll,\\u000a pitch, and yaw control systems using single-variable and multi-variable methods.

Ashish Tewari

210

Program Computes Sound Pressures at Rocket Launches  

NASA Technical Reports Server (NTRS)

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.

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

2005-01-01

211

Nuclear fuels in space rockets  

Microsoft Academic Search

There are three different types of rocket engines; solid propelled rockets, liquid propelled rockets and nuclear rockets. Nuclear rockets work by routing an appropriate gas through a nuclear rector. The reactor is at high temperature. Gas expands as it leaves the nozzle, producing a high amount of thrust. Nuclear rockets don't need an oxidizer and they require much less fuel

Ahmet Yayli; A. A. Aksit

2003-01-01

212

Action-Reaction! Rocket  

NSDL National Science Digital Library

Students construct a rocket from a balloon propelled along a guide string. They use this model to learn about Newton's three laws of motion, examining the effect of different forces on the motion of the rocket.

Integrated Teaching And Learning Program

213

Film Canister Rocket  

NSDL National Science Digital Library

In this activity, learners construct and launch rockets using simple materials and their understanding of chemical reactions. Learners can experiment by modifying their rocket designs (shapes) or "fuel packets" (baking soda).

Boston, Wgbh

2002-01-01

214

Antarctic mesospheric clouds formed from space shuttle exhaust Michael H. Stevens  

E-print Network

., 2003]. [3] Recently it was shown that space shuttle main engine exhaust injected near 110 km altitudeAntarctic mesospheric clouds formed from space shuttle exhaust Michael H. Stevens E.O. Hulburt transport of a space shuttle exhaust plume into the southern hemisphere two days after a January, 2003

Chu, Xinzhao

215

Pop Rocket Variables  

NSDL National Science Digital Library

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

216

Pulse Detonation Rocket MHD Power Experiment  

NASA Technical Reports Server (NTRS)

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

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

2002-01-01

217

STS-98 Emits Plume of Smoke  

NASA Technical Reports Server (NTRS)

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

2001-01-01

218

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

NASA Technical Reports Server (NTRS)

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.

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

1992-01-01

219

Plume Busters  

NSDL National Science Digital Library

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

Macfarlane, P.; Bohling, Geoffrey

220

Environmentally compatible solid rocket propellants  

NASA Technical Reports Server (NTRS)

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).

Jacox, James L.; Bradford, Daniel J.

1995-01-01

221

Rockets for spin recovery  

NASA Technical Reports Server (NTRS)

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

Whipple, R. D.

1980-01-01

222

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...

223

Mississippi Plumes  

NSDL National Science Digital Library

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

Center, Nasa G.; Day, Earth S.

224

Analysis on Impulse Characteristics of PDRE with Exhaust Measurements  

NASA Astrophysics Data System (ADS)

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.

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

2014-06-01

225

Rocket Wind Tunnel  

NSDL National Science Digital Library

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

Nasa

2012-05-15

226

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

227

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

228

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

229

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

230

Meteorological assessment of SRM exhaust products' environmental impact. Final report  

Microsoft Academic Search

The environmental impact of solid rocket motor (SRM) exhaust products discharged into the free air stream upon the launching of space vehicles that depend upon SRM boosters to obtain large thrust was assessed. The emission of AlâOâ to the troposphere from the SRMs in each Shuttle launch is considered. The AlâOâ appears as particles suitable for heterogeneous nucleation of hydrochloric

Dingle

1982-01-01

231

Automated and Manual Rocket Crater Measurement Software  

NASA Technical Reports Server (NTRS)

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.

Metzger, Philip; Immer, Christopher

2012-01-01

232

Antarctic mesospheric clouds formed from space shuttle exhaust  

NASA Astrophysics Data System (ADS)

New satellite observations reveal lower thermospheric transport of a space shuttle exhaust plume into the southern hemisphere two days after a January, 2003 launch. A day later, ground-based lidar observations in Antarctica identify iron ablated from the shuttle's main engines. Additional satellite observations of polar mesospheric clouds (PMCs) show a burst that constitutes 10-20% of the PMC mass between 65-79°S during the 2002-2003 season, comparable to previous results for an Arctic shuttle plume. This shows that shuttle exhaust can be an important global source of both PMC formation and variability.

Stevens, Michael H.; Meier, R. R.; Chu, Xinzhao; DeLand, Matthew T.; Plane, John M. C.

2005-07-01

233

Sounding rocket lessons learned  

NASA Technical Reports Server (NTRS)

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

Wessling, Francis C.; Maybee, George W.

1991-01-01

234

Pop! Rocket Launcher  

NSDL National Science Digital Library

In this activity, learners construct a simple air pressure launcher for paper rockets. Learners stomp or jump on an empty 2-liter bottle and force the air inside through connected plastic pipes to propel a paper rocket. The launching activity should be done in an open space like a gymnasium or cafeteria or can be conducted outside on a calm day.

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

2013-01-30

235

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

NASA Technical Reports Server (NTRS)

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

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

1977-01-01

236

Combustor and nozzle effects on particulate behavior in solid rocket motors  

NASA Astrophysics Data System (ADS)

An investigation was conducted using a subscale solid rocket motor to measure the effect of nozzle residence time on the behavior of Al2O3 particles to assess the applicability of subscale motor data to full-scale motors and to measure the effects of nozzle entrance particle size distribution on the slag accumulated with submerged nozzles. Although particles as large as 140 micrometers were present at the nozzle entrance, most of the particulate mass was contained in much smaller particles. This observation is in good agreement with the small mass that accumulated above the submerged nozzle. It was found that both particle breakup and collision coalescence occurred across the exhaust nozzle, with a significant increase in the mass fraction of small (less than 2 micrometers) particles. Increasing the nozzle residence time enhanced particle breakup but did not affect the maximum plume particle size. Thus, full-scale motors are expected to have a higher percentage of mass in particles less than 2 micrometers than subscale motors but with similar diameters of the largest particles.

Yakin, Bulent

1993-12-01

237

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

NASA Technical Reports Server (NTRS)

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

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

1986-01-01

238

Scientific debate: Mantle plumes  

NSDL National Science Digital Library

After a preliminary discussion of hotspots (emphasizing the generic term melting anomalies), the mantle plume hypothesis, and alternative hypotheses, students are assigned roles for a debate on the mantle plume controversy. Students conduct an in-class debate, presenting arguments from opposite sides of the plume debate. After the debate students write a reflection paper on their perspective on the debate.

Jordan, Brennan

239

Exhaust gas recirculation control  

Microsoft Academic Search

A transducer regulates an operating pressure which positions an exhaust gas recirculation control valve pintle to provide exhaust gas recirculation at rates which maintain the pressure in the recirculation passage upstream of the valve pintle equal to a reference pressure; exhaust gas recirculation thus varies with engine exhaust backpressure and accordingly is substantially proportional to induction air flow. The transducer

Haka

1979-01-01

240

Characterization of rocket propellant combustion products  

SciTech Connect

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

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

1991-12-09

241

Monitoring Engine Vibrations And Spectrum Of Exhaust  

NASA Technical Reports Server (NTRS)

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

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

1991-01-01

242

Andoya Rocket Range  

NSDL National Science Digital Library

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

243

HYDROGEN-OXYGEN ROCKETS  

NSDL National Science Digital Library

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

Reierson, David

244

GPS Sounding Rocket Developments  

NASA Technical Reports Server (NTRS)

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.

Bull, Barton

1999-01-01

245

Rocket engine numerical simulation  

NASA Technical Reports Server (NTRS)

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.

Davidian, Ken

1993-01-01

246

Rocket engine numerical simulator  

NASA Technical Reports Server (NTRS)

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.

Davidian, Ken

1993-01-01

247

NASA: Rocket Activities  

NSDL National Science Digital Library

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

248

Russian Rocket Engine Test  

NASA Technical Reports Server (NTRS)

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

1998-01-01

249

Rocket Motor Microphone Investigation  

NASA Technical Reports Server (NTRS)

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.

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

2010-01-01

250

Rocket Launch Probability  

NSDL National Science Digital Library

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

Exner, Nicholas

2009-01-13

251

GPS Sounding Rocket Developments  

NASA Technical Reports Server (NTRS)

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.

Bull, Barton

1999-01-01

252

Rocket Science 101  

NSDL National Science Digital Library

On occasion, when one is asked to describe a common activity or simple concept, the other party may exclaim, âÂÂWell, itâÂÂs not exactly rocket science.â Well, this website is exactly that: rocket science. To be more precise, NASA has created this elegant and visually stimulating demonstration website that allows guests the opportunity to learn how two different types of rockets (the Delta II and the Atlas V) are constructed. First-time visitors will most likely want to take advantage of the short tutorial that explains the basic part of the launch vehicle, how it is constructed, and how all of these parts effectively help launch a NASA spacecraft. After looking over this section, visitors will want to get started on constructing their own rocket. They will have the opportunity to learn about different parts of each device, and then select each item for the rocket. At the conclusion, visitors will get to see a demonstration of how each rocket works during flight.

253

Rocket Engine Oscillation Diagnostics  

NASA Technical Reports Server (NTRS)

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.

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

2002-01-01

254

Transonic, Axisymmetric Flow Over Nozzle Afterbodies With Supersonic Jet Exhausts  

NASA Technical Reports Server (NTRS)

Predictions require less computation than Navier-Stokes solutions. RAXJET computer program predicts transonic, axisymmetric flow over nozzle afterbodies with supersonic jet exhausts and includes effects of boundarylayer displacement, separation, jet entrainment, and inviscid jet plume blockage. RAXJET written in FORTRAN IV.

Wilmoth, R. G.

1983-01-01

255

Controls on plume heat flux and plume excess temperature  

Microsoft Academic Search

Plume heat flux and plume excess temperature in the upper mantle inferred from surface observations may pose important constraints on the heat flux from the core and mantle internal heating rate. This study examined the relationship between plume heat flux Q p , core-mantle boundary (CMB) heat flux Q cmb and plume excess temperature DeltaT plume in thermal convection using

Wei Leng; Shijie Zhong

2008-01-01

256

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

NASA Technical Reports Server (NTRS)

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.

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

1991-01-01

257

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

NASA Technical Reports Server (NTRS)

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.

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

2013-01-01

258

Heat Exhaustion, First Aid  

MedlinePLUS

newsletter | contact Share | Heat Exhaustion, First Aid A A A Heat exhaustion signs and symptoms can include heavy perspiration; nausea; lightheadedness; severe thirst; dilated pupils; and red or pale, ...

259

Volcanic Plume Measurements with UAV (Invited)  

NASA Astrophysics Data System (ADS)

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

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

2013-12-01

260

Exhaust gas recirculation control  

SciTech Connect

In an internal combustion engine, recirculation of exhaust gases is controlled to maintain the control pressure in a zone of the recirculation passage proportional to a reference pressure and thus to provide exhaust gas recirculation as a proportion of induction air flow. A duty cycle modulated valve controls an exhaust backpressure port and an atmospheric pressure port to create the reference pressure, whereby the proportion of exhaust gases recirculated is established by the duty cycle and is independent of the induction air flow.

Stoltman, D.D.

1983-08-23

261

Two stage turbine for rockets  

NASA Technical Reports Server (NTRS)

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.

Veres, Joseph P.

1993-01-01

262

Two stage turbine for rockets  

NASA Astrophysics Data System (ADS)

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.

Veres, Joseph P.

1993-11-01

263

THE ENVIRONMENT CREATED BY AN OPEN-AIR SOLID ROCKET PROPELLANT FIRE  

Microsoft Academic Search

A 91 kg (200 lbm) block of aluminized solid rocket propellant was burned in open air to simulate an accidental propellant fire. A suite of remote optical instruments measured the temperature and radiative properties of the plume. Solid molybdenum calorimeters provided data for heat flux estimates. Various refractory oxide and metallic witness samples placed in the fire provided temperature benchmarks and insight

L. W. HUNTER; Y. CHANG; H. N. OGUZ; J. T. WILKERSON; A. M. LENNON; R. P. CAIN; B. G. CARKHUFF; M. E. THOMAS; S. C. WALTS; C. A. MITCHELL; D. W. BLODGETT; D. H. TERRY

2007-01-01

264

Laser Rayleigh and Raman Diagnostics for Small Hydrogen/oxygen Rockets  

NASA Technical Reports Server (NTRS)

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

Degroot, Wilhelmus A.; Zupanc, Frank J.

1993-01-01

265

Exhaust gas recirculation control  

Microsoft Academic Search

In an internal combustion engine, recirculation of exhaust gases is controlled to maintain the control pressure in a zone of the recirculation passage proportional to a reference pressure and thus to provide exhaust gas recirculation as a proportion of induction air flow. A duty cycle modulated valve controls an exhaust backpressure port and an atmospheric pressure port to create the

Stoltman

1983-01-01

266

Exhaust gas recirculation control  

Microsoft Academic Search

A transducer creates an operating pressure that positions a control valve to provide exhaust gas recirculation at rates which maintain the control pressure in the recirculation passage between the valve and an orifice equal to a reference pressure; exhaust gas recirculation thus varies with engine exhaust backpressure and accordingly is substantially proportional to induction air flow. The transducer also varies

Vogelsberg

1979-01-01

267

The Optimal Bottle Rocket Lauch  

NSDL National Science Digital Library

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

Menzies, Margaret

268

Rockets in World War I  

NASA Technical Reports Server (NTRS)

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.

2004-01-01

269

Dynamics of Thermochemical Plumes  

NASA Astrophysics Data System (ADS)

We investigate the dynamics of thermo-chemical plumes to enlighten the fundamental differences with purely thermal plumes. The key features of our 3D numerical model include: (1) a compressible mantle with an endothermic phase transition at 670km depth, (2) a mantle 'wind' induced by the imposed surface plate motion, (3) twenty million active tracers simulate denser material initially in the lowermost mantle, (4) plumes form naturally i.e., without imposing any temperature perturbation. First, we investigate the widely accepted head-tail structure of plumes. Our results show that thermo-chemical plumes reaching the surface may or may not have a head since, in some cases, only a narrow 'tail' of hot material is able to ascend in the upper mantle. Therefore, we suggest that the existence of a large igneous province at the onset of hotspot volcanism is not a valid prerequisite for a deep plume origin. Second, we investigate the entrainment of deep heterogeneities. Our results show the generation of narrow, long lasting, distinct filaments in the plume's tail. Therefore, the plume conduit is laterally heterogeneous, rather than concentrically zoned. Third, we calculate the shear wave velocity anomalies in the lower mantle, using the temperature field and the distribution of chemical heterogeneities provided by the convection model. The great variety of plume's shapes and sizes differs strikingly from the expected 'mushroom' shape of purely thermal plumes, bearing important implications for the interpretation of seismologically detected plumes. Finally, our model predictions will be compared with a variety of observations in the Central Pacific.

Farnetani, C. G.; Samuel, H.

2004-12-01

270

Mechanical analysis on rocket propellants  

Microsoft Academic Search

The mechanical properties of solid rocket propellants are very important for good functioning of rocket motors. During use\\u000a and storage the mechanical properties of rocket propellants are changing, due to chemical and mechanical influences such as\\u000a thermal reactions, oxidation reactions or vibrations. These influences can result in malfunctioning, leading to an unwanted\\u000a explosion of the rocket motor. Most of modern

G. Herder; F. P. Weterings; W. P. C. de Klerk

2003-01-01

271

Rocket center Peenemuende - Personal memories  

NASA Technical Reports Server (NTRS)

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.

Dannenberg, Konrad; Stuhlinger, Ernst

1993-01-01

272

Overview of rocket engine control  

NASA Technical Reports Server (NTRS)

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

Lorenzo, Carl F.; Musgrave, Jeffrey L.

1991-01-01

273

Density-current plumes.  

PubMed

Diurnal solar heating produces an unstable warm zone just off the bottom of the inshore regions of a salt lake. The warm water rises in plumes in which brine shrimp become entrapped through apparently negative photokinetic behavior. The plumes of concentrated shrimp resemble those composed of insects in air. PMID:17775163

Mason, D T

1966-04-15

274

Density-Current Plumes  

Microsoft Academic Search

Diurnal solar heating produces an unstable warm zone just off the bottom of the inshore regions of a salt lake. The warm water rises in plumes in which brine shrimp become entrapped through apparently negative photokinetic behavior. The plumes of concentrated shrimp resemble those composed of insects in air.

David T. Mason

1966-01-01

275

COOLING TOWER PLUME MODEL  

EPA Science Inventory

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

276

The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics  

NASA Technical Reports Server (NTRS)

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The 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.

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

2014-01-01

277

Solid Rocket Booster Recovery  

NASA Technical Reports Server (NTRS)

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.

1982-01-01

278

Solid Rocket Booster Recovery  

NASA Technical Reports Server (NTRS)

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.

1982-01-01

279

Solid Rocket Booster Separation  

NASA Technical Reports Server (NTRS)

This Quick Time movie shows the Space Shuttle Solid Rocket Booster (SRB) separation from the external tank (ET). After separation, the boosters fall to the ocean from which they are retrieved 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. That is equivalent to 44 million horsepower, or the combined power of 400,000 subcompact cars.

1998-01-01

280

Rocket Activity "Hero Engine"  

E-print Network

was a spinning copper sphere that was propelled by a thrust produced by a jet of steam. The engine was an early34 Rocket Activity Pop Can "Hero Engine" Objective To investigate Newton's third law of motion engines out of soft drink cans and investigate ways to increase the action-reaction thrust produced

Provancher, William

281

This Is Rocket Science!  

NASA Astrophysics Data System (ADS)

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

Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

2013-09-01

282

Water Rocket Launch  

NSDL National Science Digital Library

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

Ieee

2014-06-18

283

Dr. Goddard Transports Rocket  

NASA Technical Reports Server (NTRS)

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

1974-01-01

284

Thiokol Solid Rocket Motors  

NASA Technical Reports Server (NTRS)

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.

Graves, S. R.

2000-01-01

285

Liquid rocket valve assemblies  

NASA Technical Reports Server (NTRS)

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.

1973-01-01

286

Liquid Rocket Engine Testing  

NASA Technical Reports Server (NTRS)

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

Rahman, Shamim

2005-01-01

287

Russian Rocket Engine Test  

NASA Technical Reports Server (NTRS)

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

1998-01-01

288

Solar rocket absorber  

SciTech Connect

A solar rocket absorber of rhenium tubes is used to provide heated liquid hydrogen to a thruster. The rhenium tubes are wrapped in a closed shape having an opening for receiving solar radiation for heating the liquid propellant. The vessel of rhenium tubes is held by a carbon shell which is further encased in a reradiation shield to prevent heat loss.

Robinson, P. I.

1985-07-16

289

Tsiolkovsky Rocket Designs  

NASA Technical Reports Server (NTRS)

By the end of the 19th Century, a Russian theorist, Konstantian Tsiolkovsky, was examining the fundamental scientific theories behind rocketry. He made some pioneering studies in liquid chemical rocket concepts and recommended liquid oxygen and liquid hydrogen as the optimum propellants. In the 1920's, Tsiolkovsky analyzed and mathematically formulated the technique for staged vehicles to reach escape velocities from Earth.

2004-01-01

290

Controls on plume heat flux and plume excess temperature  

Microsoft Academic Search

Plume heat flux and plume excess temperature in the upper mantle inferred from surface observations may pose important constraints on the heat flux from the core and mantle internal heating rate. This study examined the relationship between plume heat flux Qp, core-mantle boundary (CMB) heat flux Qcmb and plume excess temperature ?Tplume in thermal convection using both numerical modeling and

Wei Leng; Shijie Zhong

2008-01-01

291

Rapid Mars transits with exhaust-modulated plasma propulsion  

NASA Technical Reports Server (NTRS)

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

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

1995-01-01

292

Rockets using Liquid Oxygen  

NASA Technical Reports Server (NTRS)

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

Busemann, Adolf

1947-01-01

293

Seismological images of plumes  

NASA Astrophysics Data System (ADS)

The existence, physical character and morphology of mantle plumes have been widely debated for over forty years. Geodynamical and seismological investigations, the two main approaches for inferring deep Earth structure, lead to very different interpretations of Earth dynamics. The existence of at least one thermal boundary layer within the mantle (i.e., core-mantle boundary) and estimates of convective vigor from surface heat flow observations still provides a theoretical argument for the well-known “head” and “tail” plume. However associated thin cylindrical structures reaching from the Earth’s surface to the core are not clearly distinguished in blurry tomographic images. An inability to image fine-scale plume structures in the lower mantle is due to the inherently poor (compared to geodynamical models) and highly heterogeneous spatial resolution of tomographic models, and perhaps also to erroneous characterization of Earth-physics in geodynamical models. We generate synthetic tomographic images of mantle plumes using a three stage technique. First we develop 2D axisymmetric spherical models of plausible plume structures extending the models of [Lin and van Keken, 2006] for example by including the effects of compressibility and depth dependent viscosity. We maximize computational efficiency by employing a series of locally refined meshes tailored to the position of the plume “head” and “tail”. Initial investigations show that the plume velocity is sensitive to the thickness of the initial thermal boundary layer and the harmonic initial condition. The plume morphology is additionally altered when compressibility and depth-dependent viscosity is implemented. Secondly, we map thermal structure from the geodynamic models to seismic velocities using constraints from mineral physics data [Xu et al., 2008], and finally we apply the tomographic filter of the S20RTS global shear wave tomographic model [Ritsema et al., 2007] to the images of seismic wave speed to determine the tomographic expression of our plumes.

Smith, H. E.; Styles, E. E.; van Keken, P. E.; Goes, S. D.; Ritsema, J. E.

2009-12-01

294

Rocket center Peenemünde — Personal memories  

NASA Astrophysics Data System (ADS)

Von Braun built his first rockets as a young teenager. At 14, he started making plans for rockets for human travel to the Moon and Mars. The German Army began a rocket program in 1929. Two years later, Colonel (later General) Becker contacted von Braun who experimented with rockets in Berlin, gave him a contract in 1932, and, jointly with the Air Force, in 1936 built the rocket center Peenemünde where von Braun and his team developed the A-4 (V-2) rocket under Army auspices, while the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes. Albert Speer, impressed by the work of the rocketeers, allowed a modest growth of the Peenemünde project; this brought Dannenberg to the von Braun team in 1940. Hitler did not believe in rockets; he ignored the A-4 project until 1942 when he began to support it, expecting that it could turn the fortunes of war for him. He drastically increased the Peenemünde work force and allowed the transfer of soldiers from the front to Peenemünde; that was when Stuhlinger, in 1943, came to Peenemünde as a Pfc.-Ph.D. Later that year, Himmler wrenched the authority over A-4 production out of the Army's hands, put it under his command, and forced production of the immature rocket at Mittelwerk, and its military deployment against targets in France, Belgium, and England. Throughout the development of the A-4 rocket, von Braun was the undisputed leader of the project. Although still immature by the end of the war, the A-4 had proceeded to a status which made it the first successful long-range precision rocket, the prototype for a large number of military rockets built by numerous nations after the war, and for space rockets that launched satellites and traveled to the Moon and the planets.

Dannenberg, Konrad; Stuhlinger, Ernst

295

Exhaust gas recirculation control  

SciTech Connect

A regulating unit senses the pressures in two zones of a recirculation passage to create a control pressure, and a transducer regulates an operating pressure which positions a control valve to provide exhaust gas recirculation at rates which establish the pressures in the zones necessary to maintain the control pressure equal to a reference pressure. Exhaust gas recirculation thus varies with engine exhaust backpressure and accordingly is a proportion of induction air flow with the proportion being ruled by the regulating unit.

Stoltman, D.D.

1980-04-08

296

Validation of scramjet exhaust simulation technique at Mach 6  

NASA Technical Reports Server (NTRS)

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

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

1979-01-01

297

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)....

298

Sulfur plumes off Namibia  

NASA Technical Reports Server (NTRS)

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

2002-01-01

299

Method of hybrid plume plasma propulsion  

NASA Technical Reports Server (NTRS)

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.

Chang, Franklin R. (Inventor)

1990-01-01

300

ISRO's solid rocket motors  

NASA Astrophysics Data System (ADS)

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 developed. The first and second stages of 1 and 0.8 m dia respectively used low carbon steel casing and PBAN propellant. The first stage used segmented construction with a total propellant weight of 8600 kg. The second stage employed about 3 tonnes of the same propellant. The third and fourth stages were of GFRP construction and employed respectively 1100 and 275 kg of CTPB type propellants. Nozzle expansion ratios upto 30 were employed and delivered vacuum lsp of 2766 Ns/kg realized. The fourth stage motor was subsequently used as the apogee motor for orbit injection of India's first geosynchronous satellite—APPLE. All these motors have been flight proven a number of times. Further design improvements have been incorporated and these motors continue to be in use. Starting in 1984 design for a large booster was undertaken. This booster employs a nominal propellant weight of 125 tonne in a 2.8 m dia casing. The motor is expected to be qualified for flight test in 1989. Side by side a high performance motor housing nearly 7 tonnes of propellant in composite casing of 2 m dia and having flex nozzle control system is also under development for upper stage application. Details of the development of the motors, their leading specifications and performance are described.

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

1989-08-01

301

Laser rocket system analysis  

NASA Technical Reports Server (NTRS)

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

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

1979-01-01

302

Hybrid rocket performance  

NASA Technical Reports Server (NTRS)

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

Frederick, Robert A., Jr.

1992-01-01

303

Small rocket tornado probe  

SciTech Connect

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.

Colgate, S.A.

1982-01-01

304

EPDM rocket motor insulation  

NASA Technical Reports Server (NTRS)

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.

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

2003-01-01

305

EPDM rocket motor insulation  

NASA Technical Reports Server (NTRS)

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.

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

2004-01-01

306

EPDM rocket motor insulation  

NASA Technical Reports Server (NTRS)

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.

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

2008-01-01

307

Duplex tab exhaust nozzle  

NASA Technical Reports Server (NTRS)

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

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

2012-01-01

308

Exhaust backpressure tester  

Microsoft Academic Search

This patent describes a method for measuring exhaust backpressure in an internal combustion engine. It comprises: providing a pressure indicating device of the type having an elongate probe which communicates fluid pressure to an interior portion of the device; locating a wall of a manifold, pipe, muffler, catalytic converter or which is in fluid communication with an exhaust port of

1989-01-01

309

Exhaust gas recirculation control  

Microsoft Academic Search

A regulating unit senses the pressures in two zones of a recirculation passage to create a control pressure, and a transducer regulates an operating pressure which positions a control valve to provide exhaust gas recirculation at rates which establish the pressures in the zones necessary to maintain the control pressure equal to a reference pressure. Exhaust gas recirculation thus varies

Stoltman

1980-01-01

310

Gasdynamic approach to small plumes computation  

NASA Astrophysics Data System (ADS)

The semi-inverse marching characteristics scheme SIMA was extended to treat rotational flows; it is applied to computation of free plumes, starting out from non-uniform nozzle exit flow that reflects substantial viscous effects. For lack of measurements of exit flow in small nozzles, the exit plane flow is approximated by introducing a Power Law Interpolation (PLI) between the exit plane center and lip values. Exit plane flow variables thus approximated, are Mach number, pressure, flow angle and stagnation temperature. This choice is guided by gasdynamic considerations of exhaust flow from small nozzles into vacuum. The PLI is adjusted so as to obtain a match between computations and measurements at intermediate range from the nozzle. Computed plumes were found to be in good agreement with five different sets of small plume experiments. Comparative computations were performed using two alternate methods: the Boynton-Simons point-source approximation, and SIMA computation that started out from a uniform exit flow. It is demonstrated that for small nozzles having an exit flow dominated by viscous effects, the combined SIMA/PLI computational method is reasonably accurate and is dearly superior to either of the two alternate methods.

Genkin, L.; Baer, M.; Falcovitz, J.

1993-07-01

311

Liquid Rocket Engine Testing Overview  

NASA Technical Reports Server (NTRS)

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

Rahman, Shamim

2005-01-01

312

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

NASA Astrophysics Data System (ADS)

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.

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

1994-02-01

313

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

NASA Technical Reports Server (NTRS)

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.

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

1994-01-01

314

13Space Shuttle Atlantis (STS-135) -Plume speed This sequence of images  

E-print Network

it in the opposite direction to its exhaust. The plume of gas is ejected at high speed from the Shuttle main engines13Space Shuttle Atlantis (STS-135) - Plume speed This sequence of images shows the historic launch of the Space Shuttle Atlantis (STS-135) on July 8, 2011 at 11:29 a.m. EDT, from launch pad 39A at the NASA Cape

315

Optimum rocket propulsion for energy-limited transfer  

NASA Technical Reports Server (NTRS)

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 Saturn.

Zuppero, Anthony; Landis, Geoffrey A.

1991-01-01

316

Magellan Aerodynamic Characteristics During the Termination Experiment Including Thruster Plume-Free Stream Interaction  

NASA Technical Reports Server (NTRS)

Results are presented on the aerodynamic characteristics of the Magellan spacecraft during the October 1994 Termination Experiment, including the effects of the thruster engine exhaust plumes upon the molecular free stream around the spacecraft and upon the aerodynamics coefficients. As Magellan passed through the Venusian atmosphere, the solar arrays were turned in opposite directions relative to the free stream creating a torque on the spacecraft. The spacecraft control system was programmed to counter the effects of this torque with attitude control engines to maintain an inertially fixed attitude. The orientation and reaction engine telemetry returned from Magellan are used to create a model of the aerodynamic torques. Geometric models of the Magellan spacecraft are analyzed with the aid of both free molecular and Direct Simulation Monte Carlo codes. The simulated aerodynamic torques determined are compared to the measured torques. The Direct Simulation Monte Carlo method is also used to model the attitude engine exhaust plumes, the free stream disturbance caused by these plumes, and the resulting torques acting on the spacecraft compared to no-exhaust plume cases. The effect of the exhaust plumes was found to be sufficiently large that thrust reversal is possible.

Cestero, Francisco J.; Tolson, Robert H.

1998-01-01

317

Contamination control and plume assessment of low-energy thrusters  

NASA Technical Reports Server (NTRS)

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

Scialdone, John J.

1993-01-01

318

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

NASA Astrophysics Data System (ADS)

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

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

2014-11-01

319

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

NASA Technical Reports Server (NTRS)

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

Ondler, R. M.

1989-01-01

320

Space shuttle exhausted aluminum oxide: A measured particle size distribution  

Microsoft Academic Search

Aluminum oxide (A2O3) particles were collected from the space shuttle exhaust plume immediately following the launch of STS-34 on October 18, 1989. A2O3 samples were obtained at 2.4, 3.0, 3.2, and 7.4 km in altitude. The samples were analyzed using scanning electron microscopy to develop particle size distributions. There were no indications that the particle size distribution changed as a

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

1991-01-01

321

ActivitiesActivities RocketsRockets TelescopesTelescopes  

E-print Network

to midnight · Keynote talk by Dr. Carolyn Porco · Astro Q & A · Evening Star-Talks · Observe Saturn · Sky-Talks · Observe Saturn · Sky observing · Concessions available Dr. Carolyn PorcoDr. Carolyn Porco Saturday 10 amActivitiesActivities RocketsRockets TelescopesTelescopes Observe Saturn Observe Saturn Exhibits

New Hampshire, University of

322

Exhaust gas recirculation control  

SciTech Connect

A switching member simultaneously establishes a reference pressure and selects the pressure in one of two zones of a recirculation passage to create a control pressure, and a transducer regulates an operating pressure which positions a control valve to provide exhaust gas recirculation at rates which establish the pressures in the zones necessary to maintain the control pressure equal to the reference pressure. Exhaust gas recirculation thus varies with engine exhaust backpressure and accordingly is a proportion of induction air flow with the proportion being ruled by the switching member.

Haka, R. J.; Stoltman, D. D.

1980-02-05

323

Catalytic Microtube Rocket Igniter  

NASA Technical Reports Server (NTRS)

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

Schneider, Steven J.; Deans, Matthew C.

2011-01-01

324

Collapse in Thermal Plumes  

NASA Astrophysics Data System (ADS)

Collapsing thermal plumes have been investigated through experimental and numerical simulations. Collapsing plumes are an uncommon fluid dynamical phenomenon, usually seen when the buoyancy source is turned off. A series of fluid dynamical experiments were conducted on thermal plumes at a variety of temperature and viscosity contrasts, in a 26.5 cm^3 cubic tank heated by a constant temperature heater 2 cm in diameter and no-slip bottom and top surfaces. Working fluids included Lyle's Golden Syrup and ADM's Liquidose 436 syrup, which have strongly-temperature dependent viscosity and high Pr number (10^3-10^7 at experimental conditions). Visualisation included white light shadowgraphs and PIV of the central plane. Temperature contrasts ranged from 3-60°C, and two differing forms of collapse were identified. At very low temperature differences 'no rise' collapse was discovered, where the plumes stagnate in the lower third of the tank before collapsing. At temperature differences between 10-23°C normal evolution occurred until 'lens shape' collapse developed between midway and two-thirds of the distance from the base. The lens shape originated in the top of the conduit and was present throughout collapse. At temperatures above ?T=23°C the plumes follow the expected growth and shape and flatten out at the top of the tank. Thermal collapse remains difficult to explain given experimental conditions (continuous heating). Instead it is possible that small density differences arising from crystallization at ambient temperatures changes plume buoyancy-inducing collapse. We show results on the evolution of the refractive index of the syrup through time to ascertain this possibility. Preliminary numerical results using Fluidity will be presented to explore a greater parameter range of viscosity contrasts and tank aspect ratios.

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

2013-12-01

325

Rocket + Science = Dialogue  

NASA Technical Reports Server (NTRS)

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

Morris,Bruce; Sullivan, Greg; Burkey, Martin

2010-01-01

326

SlideRocket  

NSDL National Science Digital Library

As this is a New Year, there will be a need for new presentations. SlideRocket makes pesky presentation troubles go away, as you can access PowerPoint presentations from any locations, collaborate with colleagues around the world, and also integrate dynamic data, charts, and graphs quite seamlessly. Some of the more advanced features are only available via the pay versions of the product, but the free version is easy and engaging. This version is compatible with all operating systems, including Linux.

327

Rocket Noise Prediction Program  

NASA Technical Reports Server (NTRS)

A comprehensive, automated, and user-friendly software program was developed to predict the noise and ignition over-pressure environment generated during the launch of a rocket. The software allows for interactive modification of various parameters affecting the generated noise environment. Predictions can be made for different launch scenarios and a variety of vehicle and launch mount configurations. Moreover, predictions can be made for both near-field and far-field locations on the ground and any position on the vehicle. Multiple engine and fuel combinations can be addressed, and duct geometry can be incorporated efficiently. Applications in structural design are addressed.

Margasahayam, Ravi; Caimi, Raoul

1999-01-01

328

Viscoelastic rocket grain fracture analysis  

Microsoft Academic Search

A viscoelastic fracture analysis has been developed for rocket grain fracture predictions. The fracture analysis uses a stress intensity factor technique to predict crack velocity histories under thermal and pressurization loading conditions. The theory is compared with two-dimensional pressurized tests of two typical rocket motor geometries using the viscoelastic material, Solithane 113.

E. C. Francis; C. H. Carlton; R. E. Thompson

1974-01-01

329

Coal-Fired Rocket Engine  

NASA Technical Reports Server (NTRS)

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

Anderson, Floyd A.

1987-01-01

330

What fuel for a rocket?  

E-print Network

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

E. N. Miranda

2012-08-13

331

Air-Breathing Rocket Engines  

NASA Technical Reports Server (NTRS)

This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

1998-01-01

332

Beyond the thermal plume paradigm  

NASA Astrophysics Data System (ADS)

Geodynamic models of thermo-chemical plumes rising in a mantle wind suggest that we should abandon some paradigms based on the dynamics of purely thermal axisymmetric plumes. The head-tail structure is possible but not unique and the lack of a plume head does not preclude a deep origin. Our results suggest that the surface expression of some thermo-chemical plumes may be a headless, age-progressive volcanic chain. Plume tails are laterally heterogeneous, rather than concentrically zoned, because deep heterogeneities are sheared into distinct and long-lasting filaments that will be successively sampled by different volcanoes, as the oceanic plate moves over the plume tail. Finally, calculated S-wave velocity anomalies are consistent with recent plume tomographic images, showing that compositional heterogeneities in the lowermost mantle favour the coexistence of a great variety of plume shapes and sizes.

Farnetani, C. G.; Samuel, H.

2005-04-01

333

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....

334

Determination of alloy content from plume spectral measurements  

NASA Technical Reports Server (NTRS)

The mathematical derivation for a method to determine the identities and amounts of alloys present in a flame where numerous alloys may be present is described. This method is applicable if the total number of elemental species from all alloys that may be in the flame is greater than or equal to the total number of alloys. Arranging the atomic spectral line emission equations for the elemental species as a series of simultaneous equations enables solution for identity and amount of the alloy present in the flame. This technique is intended for identification and quantification of alloy content in the plume of a rocket engine. Spectroscopic measurements reveal the atomic species entrained in the plume. Identification of eroding alloys may lead to the identification of the eroding component.

Madzsar, George C.

1991-01-01

335

Rhenium Rocket Manufacturing Technology  

NASA Technical Reports Server (NTRS)

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

1997-01-01

336

Entake or exhaust valve actuator  

Microsoft Academic Search

Intake or exhaust valve actuator assembly is described for an internal combustion engine for hydraulically opening and closing an intake or exhaust valve for admitting intake gases from an intake conduit into a combustion chamber or permitting exhaust gases to escape from the combustion chamber into an exhaust conduit, the engine including a piston which oscillates in the combustion chamber,

Smietana

1993-01-01

337

Enceladus' Water Vapour Plumes  

NASA Technical Reports Server (NTRS)

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

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

2006-01-01

338

Double Diffusive Plumes  

Microsoft Academic Search

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

Bruce Sutherland; Brace Lee

2008-01-01

339

Buoyant plume calculations  

SciTech Connect

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

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

1985-01-01

340

PLUME and research sotware  

NASA Astrophysics Data System (ADS)

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.

Baudin, Veronique; Gomez-Diaz, Teresa

2013-04-01

341

COLD WEATHER PLUME STUDY  

EPA Science Inventory

While many studies of power plant plume transport and transformation have been performed during the summer, few studies of these processes during the winter have been carried out. Accordingly, the U.S. Environmental Protection Agency and the Electric Power Research Institute join...

342

Calibration and Application of Plume Radiation Code: PARRAD  

NASA Technical Reports Server (NTRS)

The next generation of Reusable Launch Vehicle (RLV) design requires careful considerations towards Thermal Protection System (TPS) not only in the high-heating regions, also in the base region. The use of hydrocarbon based liquid propulsion engine and the construction, launch and operational cost limitations places a significant burden on the base region TPS design. The radiation from the rocket-plume is expected to dominate base heating during ascent. The objective of this paper is to evaluate the accuracy of a plume radiation modeling with the PARRAD code. The original PARRAD code has been modified to work with single and multiple /overset grids and plume-field solution generated by computational codes, such as GASP. These modifications will be detailed in the proposed paper. To establish accuracy, simulations of the Space Shuttle Main Engine nozzle were performed and the radiative heating predicted by PARRAD are compared with ground based measurements. These results and pertinent discussion of issues related to accuracy will be detailed. In addition, predictions for the X-34 RLV, yet to be performed, will be presented and the significance (or lack) of base heating due to plume radiation will be discussed.

Venkatapathy, Ethiraj; Babikian, Dikran; Davies, Carol B.; Palmer, Grant; Cavolowsky, John A. (Technical Monitor)

1995-01-01

343

Impact of aircraft plume dynamics on airport local air quality  

NASA Astrophysics Data System (ADS)

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

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

2013-08-01

344

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

NASA Technical Reports Server (NTRS)

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.

Lane, John E.; Metzger, Philip T.

2014-01-01

345

Dr. Robert H. Goddard and His Rocket  

NASA Technical Reports Server (NTRS)

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

2004-01-01

346

Rocketing into Adaptive Inquiry  

NSDL National Science Digital Library

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

Beverly A. Joyce

2002-01-01

347

Sampling by mantle plumes : the legacy of the plume source  

NASA Astrophysics Data System (ADS)

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.

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

2013-12-01

348

Exhaust backpressure tester  

SciTech Connect

This patent describes a method for measuring exhaust backpressure in an internal combustion engine. It comprises: providing a pressure indicating device of the type having an elongate probe which communicates fluid pressure to an interior portion of the device; locating a wall of a manifold, pipe, muffler, catalytic converter or which is in fluid communication with an exhaust port of the internal combustion engine; creating a bore through the wall of a size sufficient to just receive the probe therethrough; inserting the probe in the bore in unsealed and unthreaded relation therewith; reading the backpressure indicated by the device; withdrawing the probe from the bore; and inserting a plug into the bore. The plug having a diameter sufficient to frictionally engage the radially inner surface of the bore thereby plugging the bore against exhaust leakage.

Freeman, F.F.

1989-12-12

349

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

NASA Technical Reports Server (NTRS)

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

1972-01-01

350

Hyperventilation and exhaustion syndrome  

PubMed Central

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 (R2 = 0.752) by scores in KES and HADS. The brief Grounding training contributed to a near significant reduction in hyperventilation (F = 2.521, p < 0.124) and to significant reductions in exhaustion scores and scores of depression and anxiety. The conclusion is that hyperventilation is common in exhaustion syndrome patients and that it can be reduced by systematic physical therapy such as Grounding. PMID:24134551

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

2014-01-01

351

Hyperventilation and exhaustion syndrome.  

PubMed

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

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

2014-12-01

352

Exhaust gas recirculation system  

SciTech Connect

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

Rachedi, S.H.

1983-08-30

353

Brassicaceae (Mustard family) Yellow rocket  

E-print Network

Brassicaceae (Mustard family) Yellow rocket Barbarea vulgaris R. Br. Life cycle Erect winter annual to identifying Christmas tree weeds. #12;Brassicaceae (Mustard family) Stems Erect, hairless and up to 3 feet

354

Scanning thermal plumes  

Microsoft Academic Search

Over a three-year period 800 thermal line scans of power plant plumes were made by an airborne scanner, with ground truth measured concurrently at the plants. Computations using centered finite differences in the thermal scanning imagery show a lower bound in the horizontal temperature gradient in excess of 1.6 C\\/m. Gradients persist to 3 m below the surface. Vector plots

F. L. Scarpace; R. P. Madding; T. Green

1975-01-01

355

Simulation of a Rocket Base Combined Cycle Exchange Inlet at Subsonic Conditions  

NASA Astrophysics Data System (ADS)

Rocket Based Combined Cycle (RBCC) engines combine the high specific impulses of air breathing engines and the large operation envelop of rockets. Such engines incorporate 4 modes of operation with the first three modes relying on the performance of a mixing duct. The performance improves with a longer mixing duct but the problem with a long mixing duct is that it increases the overall engine weight. Thus, there have been studies done by other research groups to decrease this mixing duct length. Research has been ongoing at Carleton University to design a RBCC engine concept that can potentially reduce the mixing duct length by improving mixing. This is done by using a design that expands the rocket exhaust from a singular throat through multiple clovers to a semi-annular profile. The current study focuses on the subsonic free stream flight conditions in order to analyze the rocket air interaction by using this profile. From simulations performed in ANSYS CFX 12.1, it is clear that any abrupt changes to the geometry should be avoided when designing the rocket flow path. Then in the exchange inlet / mixing duct simulations, by varying the mixing duct outlet pressure, it is found that mixing improves since the mass flow rate of air and Mach number decreases. Moreover, a comparison is done with a more conventional design that places a single rocket along the centerline (SRC). It is found that the current design outperforms the SRC configuration in terms of mixing for up to 4 mixing duct diameters downstream.

Yuen, Tommy Shi Chun

356

Solid rocket motor internal insulation  

NASA Technical Reports Server (NTRS)

Internal insulation in a solid rocket motor is defined as a layer of heat barrier material placed between the internal surface of the case propellant. The primary purpose is to prevent the case from reaching temperatures that endanger its structural integrity. Secondary functions of the insulation are listed and guidelines for avoiding critical problems in the development of internal insulation for rocket motors are presented.

Twichell, S. E. (editor); Keller, R. B., Jr.

1976-01-01

357

Exhaust gas recirculation system  

Microsoft Academic Search

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

Rachedi

1983-01-01

358

Exhaust gas recirculation control  

Microsoft Academic Search

A switching member simultaneously establishes a reference pressure and selects the pressure in one of two zones of a recirculation passage to create a control pressure, and a transducer regulates an operating pressure which positions a control valve to provide exhaust gas recirculation at rates which establish the pressures in the zones necessary to maintain the control pressure equal to

R. J. Haka; D. D. Stoltman

1980-01-01

359

Effect of gaseous and solid simulated jet plumes on an 040A space shuttle launch configuration at m=1.6 to 2.2  

NASA Technical Reports Server (NTRS)

The effect of plume-induced flow separation and aspiration effects due to operation of both orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the Space Shuttle Vehicle is determined. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold-gas jet supplied by an auxiliary 200-atm air supply system and solid-body plume simulators. The aerodynamic effects produced by these two simulation procedures are compared. The parameters most significantly affected by the jet plumes are pitching moment, elevon control effectiveness, axial force, and orbiter wing loads. The solid rocket motor (SRM) plumes have the largest effect on the aerodynamic characteristics. The effect of the orbiter plumes in combination with the SRM plumes is also significant. Variations in the nozzle design parameters and configuration changes can reduce the jet plume-induced aerodynamic effects.

Dods, J. B., Jr.; Brownson, J. J.; Kassner, D. L.; Blackwell, K. L.; Decker, J. P.; Roberts, B. B.

1974-01-01

360

Flow fields of low pressure vent exhausts  

NASA Technical Reports Server (NTRS)

The flow field produced by low pressure gas vents are described based on experimental data obtained from tests in a large vacuum chamber. The gas density, pressure, and flux at any location in the flow field are calculated based on the vent plume description and the knowledge of the flow rate and velocity of the venting gas. The same parameters and the column densities along a specified line of sight traversing the plume are also obtained and shown by a computer generated graphical representation. The fields obtained with a radically scanning Pitot probe within the exhausting gas are described by a power of the cosine function, the mass rate, and the distance from the exit port. The field measurements were made for gas at pressures ranging from 2 to 50 torr venting from pipe fittings with diameters to 3/16 to 1-1/2 inches I.D. (4.76 to 38.1 mm). The N2 mass flow rates ranged from 2E-4 to 3.7E-1 g/s.

Scialdone, John J.

1990-01-01

361

Flow fields of low pressure vent exhausts  

NASA Technical Reports Server (NTRS)

The flow field produced by low pressure gas vents are described based on experimental data obtained from tests in a large vacuum chamber. The gas density, pressure, and flux at any location in the flow field are calculated based on the vent plume description and the knowledge of the flow rate and velocity of the venting gas. The same parameters and the column densities along a specified line of sight traversing the plume are also obtained and shown by a computer-generated graphical representation. The fields obtained with a radially scanning Pitot probe within the exhausting gas are described by a power of the cosine function, the mass rate and the distance from the exit port. The field measurements were made for gas at pressures ranging from 2 to 50 torr venting from pipe fittings with diameters of 3/16 inch to 1-1/2 inches I.D. (4.76 mm to 38.1 mm). The N(2) mass flow rates ranged from 2E-4 to 3.7E-1 g/s.

Scialdone, John J.

1989-01-01

362

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

NASA Astrophysics Data System (ADS)

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

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

2010-12-01

363

Formation, Dynamics, and Impact of Plasmaspheric Plumes  

Microsoft Academic Search

Workshop on Plasmaspheric Drainage Plumes, Taos, New Mexico, 9-13 October 2006 Plasmaspheric plumes result from erosion of the plasmasphere. The Institute of Geophysics and Planetary Physics (IGPP) Workshop on Plasmaspheric Drainage Plumes was convened in Taos, N. M., on 9-13 October 2006 to examine outstanding questions about the formation and dynamics of plumes, and the impact of plumes on the

Jerry Goldstein; Joseph Borovsky; John Foster; Donald Carpenter

2007-01-01

364

Computational fluid dynamics analysis of Space Shuttle main engine multiple plume flows at high-altitude flight conditions  

Microsoft Academic Search

Computational fluid dynamics (CFD) analysis is providing verification of Space Shuttle flight performance details and is being applied to Space Shuttle Main Engine Multiple plume interaction flow field definition. Advancements in real-gas CFD methodology that are described have allowed definition of exhaust plume flow details at Mach 3.5 and 107,000 ft. The specific objective includes the estimate of flow properties

N. S. Dougherty; J. B. Holt; B. L. Liu; S. L. Johnson

1992-01-01

365

Rocket Science 101 Interactive Educational Program  

NASA Technical Reports Server (NTRS)

To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

2007-01-01

366

Upwelling relaxation and estuarine plumes  

NASA Astrophysics Data System (ADS)

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

Rao, Shivanesh; Pringle, James; Austin, Jay

2011-09-01

367

EUVS Sounding Rocket Payload  

NASA Technical Reports Server (NTRS)

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

Stern, Alan S.

1996-01-01

368

Deuterium microbomb rocket propulsion  

E-print Network

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

Friedwardt Winterberg

2008-12-02

369

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

NASA Astrophysics Data System (ADS)

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

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

370

Resistive MHD Simulations of Laminar Round Jets with Application to Magnetic Nozzle Flows  

E-print Network

in the exhaust plume of the magnetoplasma rocket known as VASIMRR. This rocket has great potential for reducing the travel time for deep space exploration missions. However, it is very difficult to investigate detachment in ground-based experiments because...

Araya, Daniel

2012-02-14

371

Multistage exhaust gas recirculation system  

Microsoft Academic Search

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

D. C. Ahrns; S. H. Rachedi

1983-01-01

372

Midwave infrared imaging Fourier transform spectrometry of combustion plumes  

NASA Astrophysics Data System (ADS)

A midwave infrared (MWIR) imaging Fourier transform spectrometer (IFTS) was used to successfully capture and analyze hyperspectral imagery of combustion plumes. Jet engine exhaust data from a small turbojet engine burning diesel fuel at a low rate of 300 cm3/min was collected at 1 cm -1 resolution from a side-plume vantage point on a 200x64 pixel window at a range of 11.2 meters. Spectral features of H2O, CO, and CO2 were present, and showed spatial variability within the plume structure. An array of thermocouple probes was positioned within the plume to aid in temperature analysis. A single-temperature plume model was implemented to obtain spatially-varying temperatures and plume concentrations. Model-fitted temperatures of 811 +/- 1.5 K and 543 +/- 1.6 K were obtained from plume regions in close proximity to thermocouple probes measuring temperatures of 719 K and 522 K, respectively. Industrial smokestack plume data from a coal-burning stack collected at 0.25 cm-1 resolution at a range of 600 meters featured strong emission from NO, CO, CO2, SO 2, and HCl in the spectral region 1800-3000 cm-1. A simplified radiative transfer model was employed to derive temperature and concentrations for clustered regions of the 128x64 pixel scene, with corresponding statistical error bounds. The hottest region (closest to stack centerline) was 401 +/- 0.36 K, compared to an in-stack measurement of 406 K, and model-derived concentration values of NO, CO2, and SO2 were 140 +/- 1 ppmV, 110,400 +/- 950 ppmV, and 382 +/- 4 ppmV compared to in-stack measurements of 120 ppmV (NOx), 94,000 ppmV, and 382 ppmV, respectively. In-stack measurements of CO and HCl were not provided by the stack operator, but model-derived values of 19 +/- 0.2 ppmV and 111 +/- 1 ppmV are reported near stack centerline. A deployment to Dugway Proving Grounds, UT to collect hyperspectral imagery of chemical and biological threat agent simulants resulted in weak spectral signatures from several species. Plume detection of methyl salicilate was achieved from both a stack release and explosive detonation, although spectral identification was not accomplished due to weak signal strength.

Bradley, Kenneth C.

373

Particle size distribution measurements in a subscale motor for the Ariane 5 solid rocket booster  

Microsoft Academic Search

An experimental determination of the combustion-chamber aluminum oxide particle-size distribution for the Ariane 5 Solid Rocket Booster is carried out. A subscale motor using a helium injection technique for quenching the reaction products is designed, manufactured and tested. A 30 percent helium-mass flow rate injection close to the head-end of the combustion chamber is found to give an exhaust aluminum

J. C. Traineau; P. Kuentzmann; M. Prevost; P. Tarrin; A. Delfour

1992-01-01

374

Nuclear thermal rockets using indigenous extraterrestrial propellants  

NASA Technical Reports Server (NTRS)

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

Zubrin, Robert M.

1990-01-01

375

Apollo Video Photogrammetry Estimation Of Plume Impingement Effects  

NASA Technical Reports Server (NTRS)

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

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

2008-01-01

376

Marshall Team Recreates Goddard Rocket  

NASA Technical Reports Server (NTRS)

In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

2003-01-01

377

Effluent sampling of Titan 3 C vehicle exhaust  

NASA Technical Reports Server (NTRS)

Downwind in situ ground-level measurements of the exhaust from a Titan 3 C launch vehicle were made during a normal launch. The measurement activity was conducted as part of an overall program to obtain field data for comparison with the multilayer dispersion model currently being used to predict the behavior of rocket vehicle exhaust clouds. All measurements were confined to land, ranging from the launch pad to approximately 2 kilometers downwind from the pad. Measurement systems included detectors for hydrogen chloride (HCl), carbon dioxide (CO2), and particulates (Al2O3). Airborne and ground-based optical systems were employed to monitor exhaust cloud rise, growth, and movement. These measurement systems, located along the ground track (45 deg azimuth from the launch pad) of the exhaust cloud, showed no effluents attributable to the launch. Some hydrogen chloride and aluminum oxide were detected in the surface wind direction (15 deg azimuth) from the pad. Comparisons with the model were made in three areas: (1) assumption of cloud geometry at stabilization; (2) prediction of cloud stabilization altitude; and (3) prediction of the path of cloud travel. In addition, the importance of elemental analyses of the particulate samples is illustrated.

Gregory, G. L.; Storey, R. W., Jr.

1975-01-01

378

Atmospheric chemistry in volcanic plumes  

PubMed Central

Recent field observations have shown that the atmospheric plumes of quiescently degassing volcanoes are chemically very active, pointing to the role of chemical cycles involving halogen species and heterogeneous reactions on aerosol particles that have previously been unexplored for this type of volcanic plumes. Key features of these measurements can be reproduced by numerical models such as the one employed in this study. The model shows sustained high levels of reactive bromine in the plume, leading to extensive ozone destruction, that, depending on plume dispersal, can be maintained for several days. The very high concentrations of sulfur dioxide in the volcanic plume reduces the lifetime of the OH radical drastically, so that it is virtually absent in the volcanic plume. This would imply an increased lifetime of methane in volcanic plumes, unless reactive chlorine chemistry in the plume is strong enough to offset the lack of OH chemistry. A further effect of bromine chemistry in addition to ozone destruction shown by the model studies presented here, is the oxidation of mercury. This relates to mercury that has been coemitted with bromine from the volcano but also to background atmospheric mercury. The rapid oxidation of mercury implies a drastically reduced atmospheric lifetime of mercury so that the contribution of volcanic mercury to the atmospheric background might be less than previously thought. However, the implications, especially health and environmental effects due to deposition, might be substantial and warrant further studies, especially field measurements to test this hypothesis. PMID:20368458

von Glasow, Roland

2010-01-01

379

Stationary plasma thruster plume characteristics  

Microsoft Academic Search

Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of

Roger M. Myers; David H. Manzella

1994-01-01

380

Mantle plumes and flood basalts  

Microsoft Academic Search

We discuss the geological, geophysical, and petrological observations that constrain the nature of mantle convection in plumes, and show how theoretical models of mantle plumes have developed over the past three decades. The large volumes of lava emplaced in geologically short periods as flood basalts are generated mainly by decompression melting of abnormally hot mantle brought to the base of

R. S. White; D. P. Mckenzie

1995-01-01

381

Automated Rocket Propulsion Test Management  

NASA Technical Reports Server (NTRS)

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

Walters, Ian; Nelson, Cheryl; Jones, Helene

2007-01-01

382

Evidence of Residual Plasmaspheric Plumes  

NASA Astrophysics Data System (ADS)

Plasmaspheric erosion produces plumes of plasma that extend sunward from the main torus. When geomagnetic activity decreases, a given plume loses its sunward orientation, rotating eastward and wrapping itself around the plasmasphere torus. The residual plume is a major feature of the recovery phase plasmasphere, and is suspected to be an important influence upon the loss rates of energetic particles. In this study, comparison between in situ observations of the Los Alamos National Laboratory (LANL) Magnetospheric Plasma Analyzers (MPA) and output of a plasmapause test particle (PTP) simulation for the moderately disturbed interval 18--20 January 2000 reveals evidence of plasmaspheric plumes that wrapped completely around the main torus and lasted for at least 40 hours and possibly as long as 60 hours. The presence of long-lived multiple wrapped residual plumes suggests that the global plasmaspheric density distribution preserves some memory of prior epochs of erosion and recovery.

Goldstein, J.; Thomsen, M.

2007-12-01

383

Anchoring Atmospheric Density Models Using Observed Shuttle Plume Emissions  

NASA Astrophysics Data System (ADS)

Atmospheric number densities at a given low-earth orbit (LEO) altitude can vary by more than an order of magnitude, depending on such parameters as diurnal variations and solar activity. The MSIS atmospheric model, which includes these dependent variables as input, is reported as being accurate to ±15%. Improvement to such models requires accurate direct atmospheric measurement. Here, a means of anchoring atmospheric models is offered through measuring the size and shape of atomic line or molecular band radiance resulting from the atmospheric interaction from rocket engine plumes or gas releases in LEO. Many discrete line or band emissions, ranging from the infrared to the ultraviolet may be suitable. For this purpose we are focusing on NH(A?X), centered at 316 nm. This emission is seen in the plumes of the Shuttle Orbiter PRCS engines, is expected in the plume of any amine fueled engine, and can be observed from remote sensors in space or on the ground. The atmospheric interaction of gas releases or plumes from spacecraft in LEO are understood by comparison of observed radiance with that predicted by Direct Simulation Monte Carlo (DSMC) models. The recent Extended Variable Hard Sphere (EVHS) improvements in treating hyperthermal collisions has produced exceptional agreement between measured and modeled steady-state Space Shuttle OMS and PRCS 190-250 nm Cameron band plume radiance from CO(a?X), which is understood to result from a combination of two- and three-step mechanisms. Radiance from NH(A?X) in far field plumes is understood to result from a simpler single-step process of the reaction of a minor plume species with atomic oxygen, making it more suitable for use in determining atmospheric density. It is recommended that direct retrofire burns of amine fueled engines be imaged in a narrow band from remote sensors to reveal atmospheric number density. In principal the simple measurement of the distance between the engine exit and the peak in the steady-state radiance from LEO spacecraft can indicate atmospheric density to ~1% accuracy. Use of this radiance requires calibration by an accurate independent measurement associated with a well-resolved steady-state image of it.

Dimpfl, W. L.; Bernstien, L. S.

2010-12-01

384

Far ultraviolet astronomy from sounding rockets  

Microsoft Academic Search

The contributions of sounding rocket observations to non-solar far ultraviolet astronomy are reviewed. The instruments and techniques used to investigate the interstellar medium and far-ultraviolet emissions from cool star chromospheres and planetary atmospheres are discussed, and sounding rocket UV observations of Comets Kohoutek and West are described. The use of sounding rockets to provide calibrations of Orbiting Astronomical Observatory measurements

G. R. Carruthers

1976-01-01

385

Premature ignition of a rocket motor.  

SciTech Connect

During preparation for a rocket sled track (RST) event, there was an unexpected ignition of the zuni rocket motor (10/9/08). Three Sandia staff and a contractor were involved in the accident; the contractor was seriously injured and made full recovery. The data recorder battery energized the low energy initiator in the rocket.

Moore, Darlene Ruth

2010-10-01

386

Consort 1 sounding rocket flight  

NASA Technical Reports Server (NTRS)

This paper describes a payload of six experiments developed for a 7-min microgravity flight aboard a sounding rocket Consort 1, in order to investigate the effects of low gravity on certain material processes. The experiments in question were designed to test the effect of microgravity on the demixing of aqueous polymer two-phase systems, the electrodeposition process, the production of elastomer-modified epoxy resins, the foam formation process and the characteristics of foam, the material dispersion, and metal sintering. The apparatuses designed for these experiments are examined, and the rocket-payload integration and operations are discussed.

Wessling, Francis C.; Maybee, George W.

1989-01-01

387

The Advanced Solid Rocket Motor  

NASA Technical Reports Server (NTRS)

The paper describes the Advanced Solid Rocket Motor (ASRM) that is being developed to replace, in 1997, the Redesigned Solid Rocket Motor which currently boosts the Space Shuttle. The ASRM will contain features to improve motor safety (fewer potential leak paths, improved seal materials, stronger case material, and fewer nozzle and case joints), an improved ignition system using through-bulkhead initiators, and highly reproducible manufacturing and inspection techniques with a large number of automated procedures. The ASRM will be able to deliver 12,000 lbs greater payloads to any given orbit of the Shuttle. There are also environmental improvements, realized by waste propellant recovery.

Mitchell, Royce E.

1992-01-01

388

Rocket study of auroral processes  

NASA Technical Reports Server (NTRS)

Abstracts are presented of previously published reports analyzing data from three Echo 3 rocket flights. Particle experiments designed for the Terrier-Malmute flight, the Echo 5 flight, and the Norwegian Corbier Ferdinand 50 flight are described and their flight performance evaluated. Theoretical studies on auroral particle precipitation are reviewed according to observations made in three regions of space: (1) the region accessible to rockets and low altitude satellites (few hundred to a few thousand kilometers); (2) the region extending from 4000 to 8000 km (S3-3 satellite range); and (3) near the equatorial plane (geosynchronous satellite measurements). Questions raised about auroral arc formation are considered.

Arnoldy, R. L.

1981-01-01

389

The multichannel spectral device with transmitting analyzed optical signals by the optical fiber for the liquid propellant rocket engine diagnostics  

NASA Astrophysics Data System (ADS)

Recently spectroscopic methods of diagnostics of the liquid propellant rocket engines for prevention of emergencies become topical. Such diagnostics is based on tracking appearance of spectral lines of engine's constructional materials in spectrum of rocket plume radiation and glow dynamics of these lines. The multichannel optical spectral device considered in this paper makes the contactless spectrum analysis of optical radiations that allows to use this spectral device for diagnostics of the liquid propellant rocket engines. The novelty of this spectral device lies in application of the fiber-optical bundle and N parallel channels of the spectrum analysis [1]. Each channel contains the narrow-band optical filtration which has been set on the certain wave length. The fiber-optical bundle is used for transmitting analyzed optical radiation on a safe distance for the device from the rocket engine. This method of the contactless rocket engine diagnostics allows to except a direct contact of the spectral device with the field of rocket blast radiation and to eliminate negative influence of the engine on the spectral device, for example the acoustic impact.

Vaganov, Mikhail A.; Moskaletz, Oleg D.

2013-09-01

390

Volatile nanoparticle formation and growth within a diluting diesel car exhaust.  

PubMed

A major source of particle number emissions is road traffic. However, scientific knowledge concerning secondary particle formation and growth of ultrafine particles within vehicle exhaust plumes is still very limited. Volatile nanoparticle formation and subsequent growth conditions were analyzed here to gain a better understanding of "real-world" dilution conditions. Coupled computational fluid dynamics and aerosol microphysics models together with measured size distributions within the exhaust plume of a diesel car were used. The impact of soot particles on nucleation, acting as a condensational sink, and the possible role of low-volatile organic components in growth were assessed. A prescribed reduction of soot particle emissions by 2 orders of magnitude (to capture the effect of a diesel particle filter) resulted in concentrations of nucleation-mode particles within the exhaust plume that were approximately 1 order of magnitude larger. Simulations for simplified sulfuric acid-water vapor gas-oil containing nucleation-mode particles show that the largest particle growth is located in a recirculation zone in the wake of the car. Growth of particles within the vehicle exhaust plume up to detectable size depends crucially on the relationship between the mass rate of gaseous precursor emissions and rapid dilution. Chassis dynamometer measurements indicate that emissions of possible hydrocarbon precursors are significantly enhanced under high engine load conditions and high engine speed. On the basis of results obtained for a diesel passenger car, the contributions from light diesel vehicles to the observed abundance of measured nucleation-mode particles near busy roads might be attributable to the impact of two different time scales: (1) a short one within the plume, marked by sufficient precursor emissions and rapid dilution; and (2) a second and comparatively long time scale resulting from the mix of different precursor sources and the impact of atmospheric chemistry. PMID:21516935

Uhrner, Ulrich; Zallinger, Michael; von Löwis, Sibylle; Vehkamäki, Hanna; Wehner, Birgit; Stratmann, Frank; Wiedensohler, Alfred

2011-04-01

391

Solid rocket motor witness test  

NASA Technical Reports Server (NTRS)

The Solid Rocket Motor Witness Test was undertaken to examine the potential for using thermal infrared imagery as a tool for monitoring static tests of solid rocket motors. The project consisted of several parts: data acquisition, data analysis, and interpretation. For data acquisition, thermal infrared data were obtained of the DM-9 test of the Space Shuttle Solid Rocket Motor on December 23, 1987, at Thiokol, Inc. test facility near Brigham City, Utah. The data analysis portion consisted of processing the video tapes of the test to produce values of temperature at representative test points on the rocket motor surface as the motor cooled down following the test. Interpretation included formulation of a numerical model and evaluation of some of the conditions of the motor which could be extracted from the data. These parameters included estimates of the insulation remaining following the tests and the thickness of the charred layer of insulation at the end of the test. Also visible was a temperature signature of the star grain pattern in the forward motor segment.

Welch, Christopher S.

1991-01-01

392

Liquid propellant rocket combustion instability  

NASA Technical Reports Server (NTRS)

The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

Harrje, D. T.

1972-01-01

393

Centrifugal pumps for rocket engines  

NASA Technical Reports Server (NTRS)

The use of centrifugal pumps for rocket engines is described in terms of general requirements of operational and planned systems. Hydrodynamic and mechanical design considerations and techniques and test procedures are summarized. Some of the pump development experiences, in terms of both problems and solutions, are highlighted.

Campbell, W. E.; Farquhar, J.

1974-01-01

394

Empirical Characterization of Plasmaspheric Plumes  

NASA Astrophysics Data System (ADS)

The formation and subsequent development of plasmaspheric plumes in a given convection event is well characterized by a series of distinct phases, each triggered by a change in the strength of global magnetospheric convection. For example, after a prolonged quiet period, an increase in convection strength triggers erosion, and the formation of a sunward-pointing plume that typically spans at least a few hours of dayside magnetic local time (MLT). On the other hand, a convection decrease causes a pre-existing plume to begin rotating eastward, eventually becoming wrapped around the main plasmaspheric torus. Predicted by computational models, plume phases have since proven to be a consistent feature of plasmaspheric dynamics in numerous observations made by the Imager for Magnetopause-to-Aurora Global Exploration (IMAGE) spacecraft during the years 2000-2005. However, currently-existing empirical models for the plasmasphere do not include plumes or plume phases. We present first results of an empirical model of plume density and location using (respectively) measurements by the IMAGE radio plasma imager (RPI) and extreme ultraviolet (EUV) instruments. This new model framework differs differs from current models in two ways. First, it represents the plasmapause as a multi-valued function of L versus MLT. Second, it incorporates the concept of plume phases by parameterizing plasmaspheric density based on superposed epoch analysis. An empirical characterization of plasmaspheric plume density and location is an important step toward better knowledge of the spatial and temporal dependence of critical wave-particle interactions affecting ring current ions and outer radiation belt electrons.

Goldstein, J.; Denton, R. E.; Sandel, B. R.

2008-12-01

395

Comprehensive simultaneous shipboard and airborne characterization of exhaust from a modern container ship at sea.  

PubMed

We report the first joint shipboard and airborne study focused on the chemical composition and water-uptake behavior of particulate ship emissions. The study focuses on emissions from the main propulsion engine of a Post-Panamax class container ship cruising off the central coast of California and burning heavy fuel oil. Shipboard sampling included micro-orifice uniform deposit impactors (MOUDI) with subsequent off-line analysis, whereas airborne measurements involved a number of real-time analyzers to characterize the plume aerosol, aged from a few seconds to over an hour. The mass ratio of particulate organic carbon to sulfate at the base of the ship stack was 0.23 +/- 0.03, and increased to 0.30 +/- 0.01 in the airborne exhaust plume, with the additional organic mass in the airborne plume being concentrated largely in particles below 100 nm in diameter. The organic to sulfate mass ratio in the exhaust aerosol remained constant during the first hour of plume dilution into the marine boundary layer. The mass spectrum of the organic fraction of the exhaust aerosol strongly resembles that of emissions from other diesel sources and appears to be predominantly hydrocarbon-like organic (HOA) material. Background aerosol which, based on air mass back trajectories, probably consisted of aged ship emissions and marine aerosol, contained a lower organic mass fraction than the fresh plume and had a much more oxidized organic component. A volume-weighted mixing rule is able to accurately predict hygroscopic growth factors in the background aerosol but measured and calculated growth factors do not agree for aerosols in the ship exhaust plume. Calculated CCN concentrations, at supersaturations ranging from 0.1 to 0.33%, agree well with measurements in the ship-exhaust plume. Using size-resolved chemical composition instead of bulk submicrometer composition has little effect on the predicted CCN concentrations because the cutoff diameter for CCN activation is larger than the diameter where the mass fraction of organic aerosol begins to increase significantly. The particle number emission factor estimated from this study is 1.3 x 10(16) (kg fuel)(-1), with less than 1/10 of the particles having diameters above 100 nm; 24% of particles (>10 nm in diameter) activate into cloud droplets at 0.3% supersaturation. PMID:19673244

Murphy, Shane M; Agrawal, Harshit; Sorooshian, Armin; Padró, Luz T; Gates, Harmony; Hersey, Scott; Welch, W A; Lung, H; Miller, J W; Cocker, David R; Nenes, Athanasios; Jonsson, Haflidi H; Flagan, Richard C; Seinfeld, John H

2009-07-01

396

COMPARING AND LINKING PLUMES ACROSS MODELING APPROACHES  

EPA Science Inventory

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

397

IMPROVED PREDICTION OF BENDING PLUMES  

EPA Science Inventory

Integral plume models harbor a fundamental, often significant error because the standard implementation of control volumes, or elements, is inconsistent with the overall geometry of the problem. he error, called negative volume anomaly, occurs irregularly, being contingent on the...

398

Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter  

NASA Technical Reports Server (NTRS)

Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

Messing, Wesley E.

1959-01-01

399

Rocket dynamics (2nd revised and enlarged edition)  

NASA Astrophysics Data System (ADS)

The book presents the fundamentals of the dynamics of ballistic and spacecraft-launching rockets. In particular, attention is given to equations of rocket motion, stability and controllability of rockets, stabilization of rocket motion with allowance for liquid propellent motions in the tank, and rocket stabilization with allowance for rocket shell elasticity. The discussion also covers stabilization of the longitudinal oscillations of liquid propellant rockets, stabilization of the rotational motion of rockets and spacecraft along free-flight trajectories, and experimental methods for studying the dynamic characteristics of rockets.

Abgarian, Karlen A.; Kaliazin, Ernst L.; Mishin, Vasilii P.; Rapoport, Il'ia M.

400

Numerical modeling of exhaust smoke dispersion for a generic frigate and comparisons with experiments  

NASA Astrophysics Data System (ADS)

The exhaust smoke dispersion for a generic frigate is investigated numerically through the numerical solution of the governing fluid flow, energy, species and turbulence equations. The main objective of this work is to obtain the effects of the yaw angle, velocity ratio and buoyancy on the dispersion of the exhaust smoke. The numerical method is based on the fully conserved control-volume representation of the fully elliptic Navier-Stokes equations. Turbulence is modeled using a two-equation ( k- ?) model. The flow visualization tests using a 1/100 scale model of the frigate in the wind tunnel were also carried out to determine the exhaust plume path and to validate the computational results. The results show that down wash phenomena occurs for the yaw angles between ? =10° and 20°. The results with different exhaust gas temperatures show that the buoyancy effect increases with the increasing of the exhaust gas temperature. However, its effect on the plume rise is less significant in comparison with its momentum. A good agreement between the predictions and experiment results is obtained.

Ergin, Selma; Dobrucal?, Erinç

2014-06-01

401

Space simulation experiments on reaction control system thruster plumes  

NASA Technical Reports Server (NTRS)

A space simulation procedure was developed for studying rocket plume contamination effects using a 5-pound bipropellant reaction control system thruster. Vacuum chamber pressures of 3 x 10 to the minus 5 torr (70 miles altitude) were achieved with the thruster firing in pulse trains consisting of eight pulses (50 msec on, 100 msec off, and seven minutes between pulse trains). The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. An effort was made to simulate propellant system flow dynamics corresponding to actual spacecraft mission use. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.

Cassidy, J. F.

1972-01-01

402

Space simulation experiments on reaction control system thruster plumes.  

NASA Technical Reports Server (NTRS)

A space simulation procedure was developed for studying rocket plume contamination effects using a 5-lb bipropellant reaction control system thrustor. Vacuum chamber pressures of 0.00003 torr (70 miles altitude) were achieved with the thrustor firing in pulse trains consisting of eight pulses - 50 msec on, 100 msec off, and seven minutes between pulse trains. The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid-helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.

Cassidy, J. F.

1972-01-01

403

Exhaust gas sensors  

SciTech Connect

The automotive industry needed a fast, reliable, under-the-hood method of determining nitrogen oxides in automobile exhaust. Several technologies were pursued concurrently. These sensing technologies were based on light absorption, electrochemical methods, and surface mass loading. The Y-12 plant was selected to study the methods based on light absorption. The first phase was defining the detailed technical objectives of the sensors--this was the role of the automobile companies. The second phase was to develop prototype sensors in the laboratories--the national laboratories. The final phase was testing of the prototype sensors by the automobile industries. This program was canceled a few months into what was to be a three-year effort.

Hiller, J. [Lockheed Martin Energy Systems, Inc., Oak Ridge, TN (United States); Miree, T.J. [Ford Motor Co., Allen Park, MI (United States)

1997-02-09

404

Variable area exhaust nozzle  

NASA Technical Reports Server (NTRS)

An exhaust nozzle for a gas turbine engine comprises a number of arcuate flaps pivotally connected to the trailing edge of a cylindrical casing which houses the engine. Seals disposed within the flaps are spring biased and extensible beyond the side edges of the flaps. The seals of adjacent flaps are maintained in sealing engagement with each other when the flaps are adjusted between positions defining minimum nozzle flow area and the cruise position. Extensible, spring biased seals are also disposed within the flaps adjacent to a supporting pylon to thereby engage the pylon in a sealing arrangement. The flaps are hinged to the casing at the central portion of the flaps' leading edges and are connected to actuators at opposed outer portions of the leading edges to thereby maximize the mechanical advantage in the actuation of the flaps.

Johnston, E. A. (inventor)

1979-01-01

405

Power Exhaust in Fusion Plasmas  

NASA Astrophysics Data System (ADS)

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

Fundamenski, Wojciech

2014-07-01

406

Treatment of power utilities exhaust  

SciTech Connect

Provided is a process for treating nitrogen oxide-containing exhaust produced by a stationary combustion source by the catalytic reduction of nitrogen oxide in the presence of a reductant comprising hydrogen, followed by ammonia selective catalytic reduction to further reduce the nitrogen oxide level in the exhaust.

Koermer, Gerald (Basking Ridge, NJ)

2012-05-15

407

Automotive Fuel and Exhaust Systems.  

ERIC Educational Resources Information Center

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

Irby, James F.; And Others

408

Active Volcanic Plumes on Io  

NASA Technical Reports Server (NTRS)

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

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

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

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

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

1997-01-01

409

Unique nuclear thermal rocket engine  

NASA Astrophysics Data System (ADS)

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

Culver, Donald W.; Rochow Wilcox Space; Nuclear Systems, Richard

1993-01-01

410

Unique nuclear thermal rocket engine  

SciTech Connect

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

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

1993-01-15

411

Relativistic rocket: Dream and reality  

NASA Astrophysics Data System (ADS)

The dream of interstellar flights persists since the first pioneers in astronautics and has never died. Many concepts of thruster capable to propel a rocket to the stars have been proposed and the most suitable among them are thought to be photon propulsion and propulsion by the products of proton-antiproton annihilation in magnetic nozzle. This article addresses both concepts allowing for cross-section of annihilation among other issues in order to show their vulnerability and to indicate the problems. The concept of relativistic matter propulsion is substantiated and discussed. The latter is argued to be the most straightforward way to build-up a relativistic rocket firstly because it is based on the existing technology of ion generators and accelerators and secondly because it can be stepped up in efflux power starting from interplanetary spacecrafts powered by nuclear reactors to interstellar starships powered by annihilation reactors. The problems imposed by thermodynamics and heat disposal are accentuated.

Semyonov, Oleg G.

2014-06-01

412

Thermal stratification potential in rocket engine coolant channels  

NASA Technical Reports Server (NTRS)

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

Kacynski, Kenneth J.

1992-01-01

413

Theory for Plasma Rocket Propulsion  

NASA Astrophysics Data System (ADS)

Electrical propulsion of rockets is developing potentially into the use of 3 different thrusters for future long-distance space missions that primarily involve plasma dynamics. These are the Magnetoplasmadynamic (MPD) Thruster, the Plasma Induction Thruster (PID), and the VASIMIR Thruster. The history of the development of electrical propulsion into these prospects and the current research of particularly the VASIMIR Thruster are reviewed. Theoretical questions that need to be addressed in that development are explored.

Grabbe, Crockett

2009-11-01

414

The SCIFER sounding rocket experiment  

NASA Technical Reports Server (NTRS)

The sounding of the cleft on ion fountain energization region (SCIFER) experiment is described. The purpose of the SCIFER experiment was to study the upper ionosphere and cleft ion fountain by overflying Svalbard (Norway) with sounding rockets. Deep ionospheric density canyons were observed. The SCIFER demonstrated the correlation between accelerated ions, broadband low frequency electric fields, and reduced plasma density at 1400 km altitude in the pre-noon cleft.

Kintner, P. M.; Arnoldy, R.; Pollock, C.; Moore, T.; Holtet, J.; Deehr, C.; Moen, J.

1997-01-01

415

The Structure and Origin of Solar Plumes: Network Plumes  

NASA Astrophysics Data System (ADS)

This study is based upon plumes seen close to the solar limb within coronal holes in the emission from ions formed in the temperature region of 1 MK, in particular, the band of Fe IX 171 Å from EIT on the Solar and Heliospheric Observatory. It is shown, using geometric arguments, that two distinct classes of structure contribute to apparently similar plume observations. Quasi-cylindrical structures are anchored in discrete regions of the solar surface (beam plumes), and faint extended structures require integration along the line of sight (LOS) in order to reproduce the observed brightness. This second category, sometimes called "curtains," are ubiquitous within the polar holes and are usually more abundant than the beam plumes, which depend more on the enhanced magnetic structures detected at their footpoints. It is here proposed that both phenomena are based on plasma structures in which emerging magnetic loops interact with ambient monopolar fields, involving reconnection. The important difference is in terms of physical scale. It is proposed that curtains are composed of a large number of microplumes, distributed along the LOS. The supergranule network provides the required spatial structure. It is shown by modeling that the observations can be reproduced if microplumes are concentrated within some 5 Mm of the cell boundaries. For this reason, we propose to call this second population "network plumes." The processes involved could represent a major contribution to the heating mechanism of the solar corona.

Gabriel, A.; Bely-Dubau, F.; Tison, E.; Wilhelm, K.

2009-07-01

416

Rocket Engine Altitude Simulation Technologies  

NASA Technical Reports Server (NTRS)

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

Woods, Jody L.; Lansaw, John

2010-01-01

417

Validation of Inlet and Exhaust Boundary Conditions for a Cartesian Method  

NASA Technical Reports Server (NTRS)

Inlets and exhaust nozzles are often omitted in aerodynamic simulations of aircraft due to the complexities involved in the modeling of engine details and flow physics. However, the omission is often improper since inlet or plume flows may have a substantial effect on vehicle aerodynamics. A method for modeling the effect of inlets and exhaust plumes using boundary conditions within an inviscid Cartesian flow solver is presented. This approach couples with both CAD systems and legacy geometry to provide an automated tool suitable for parameter studies. The method is validated using two and three-dimensional test problems which are compared with both theoretical and experimental results. The numerical results demonstrate excellent agreement with theory and available data, even for extremely strong jets and very sensitive inlets.

Pandya, Shishir A.; Murman, Scott M.; Aftosmis, Michael J.

2004-01-01

418

A hierachy of dynamic plume models incorporating uncertainty: Volume 2, Stack Exhause Model (SEM): Final report  

Microsoft Academic Search

The Stack Exhaust Model (SEM) is the highest resolution member of a hierarchy of models that predict the mean of time-averaged sampler observations downwind of a fossil-fueled power plant stack, along with an estimate of the variation about this mean. To represent the turbulent atmosphere surrounding the plume compatibly with available meteorological data, a second order closure sub-model is used.

R. I. Sykes; W. S. Lewellen; S. F. Parker; D. S. Henn

1989-01-01

419

Navier-Stokes simulation of plume\\/Vertical Launching System interaction flowfields  

Microsoft Academic Search

The application of Navier-Stokes methodology to the analysis of Vertical Launching System\\/missile exhaust plume interactions is discussed. The complex 3D flowfields related to the Vertical Launching System are computed utilizing the PARCH\\/RNP Navier-Stokes code. PARCH\\/RNP solves the fully-coupled system of fluid, two-equation turbulence (k-epsilon) and chemical species equations via the implicit, approximately factored, Beam-Warming algorithm utilizing a block-tridiagonal inversion procedure.

B. J. York; N. Sinha; S. M. Dash; L. Anderson; L. Gominho

1992-01-01

420

Summary of Experiments Performed to Investigate the Effects of Ion Thruster Plumes on Microwave Propagation  

NASA Technical Reports Server (NTRS)

Electric propulsion systems have now reached a level of maturity where they are being used on operational spacecraft. One concern for the designers however, is the effect of the ion exhaust plumes produced by the systems, on microwave communication with the spacecraft. To better understand these effects, a number of propagation experiments were performed at the NASA Glenn Research Center with an operating ion thruster. This report describes the experiments and presents the results of the data obtained.

Lambert, Kevin M.; Zaman, Afroz J.

1999-01-01

421

Artificial ageing of double base rocket propellant  

Microsoft Academic Search

The ageing of double base rocket propellants (DB rocket propellants), which is a consequence of chemical reactions and physical\\u000a processes that take place over time, has significant effect on their relevant properties (e.g. chemical composition, mechanical\\u000a properties, ballistic properties, etc.). The changes of relevant properties limit the safe and reliable service life of DB\\u000a rocket propellants. This is the reason

S. Mate?i? Mušani?; M. Su?eska

2009-01-01

422

SHARPI/PICTURE Sounding Rocket Telescope  

NASA Technical Reports Server (NTRS)

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

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

2006-01-01

423

Dr. Robert H. Goddard and His Rocket  

NASA Technical Reports Server (NTRS)

Goddard rocket in launching tower at Roswell, New Mexico, March 21, 1940. Fuel was injected by pumps from the fueling platform at left. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets, which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

1940-01-01

424

GASOLINE VEHICLE EXHAUST PARTICLE SAMPLING STUDY  

SciTech Connect

The University of Minnesota collaborated with the Paul Scherrer Institute, the University of Wisconsin (UWI) and Ricardo, Inc to physically and chemically characterize the exhaust plume from recruited gasoline spark ignition (SI) vehicles. The project objectives were: (1) Measure representative particle size distributions from a set of on-road SI vehicles and compare these data to similar data collected on a small subset of light-duty gasoline vehicles tested on a chassis dynamometer with a dilution tunnel using the Unified Drive Cycle, at both room temperature (cold start) and 0 C (cold-cold start). (2) Compare data collected from SI vehicles to similar data collected from Diesel engines during the Coordinating Research Council E-43 project. (3) Characterize on-road aerosol during mixed midweek traffic and Sunday midday periods and determine fleet-specific emission rates. (4) Characterize bulk- and size-segregated chemical composition of the particulate matter (PM) emitted in the exhaust from the gasoline vehicles. Particle number concentrations and size distributions are strongly influenced by dilution and sampling conditions. Laboratory methods were evaluated to dilute SI exhaust in a way that would produce size distributions that were similar to those measured during laboratory experiments. Size fractionated samples were collected for chemical analysis using a nano-microorifice uniform deposit impactor (nano-MOUDI). In addition, bulk samples were collected and analyzed. A mixture of low, mid and high mileage vehicles were recruited for testing during the study. Under steady highway cruise conditions a significant particle signature above background was not measured, but during hard accelerations number size distributions for the test fleet were similar to modern heavy-duty Diesel vehicles. Number emissions were much higher at high speed and during cold-cold starts. Fuel specific number emissions range from 1012 to 3 x 1016 particles/kg fuel. A simple relationship between number and mass emissions was not observed. Data were collected on-road to compare weekday with weekend air quality around the Twin Cities area. This portion of the study resulted in the development of a method to apportion the Diesel and SI contribution to on-road aerosol.

Kittelson, D; Watts, W; Johnson, J; Zarling, D Schauer,J Kasper, K; Baltensperger, U; Burtscher, H

2003-08-24

425

Low thrust chemical rocket technology  

NASA Technical Reports Server (NTRS)

An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher performance propellants were evaluated: Space storable propellants, including liquid oxygen (LOX) as the oxidizer with nitrogen hydrides or hydrocarbon as fuels. Specifically, a LOX/hydrazine engine was designed, fabricated, and shown to have a 95 pct theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1. Further performance improvment can be obtained by the use of LOX/hydrogen propellants, especially for manned spacecraft applications, and specific designs must be developed and advanced through flight qualification.

Schneider, Steven J.

1992-01-01

426

Focused Rocket-Ejector RBCC Experiments  

NASA Technical Reports Server (NTRS)

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

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

427

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2014 CFR

...Section 1507.10 Commercial Practices CONSUMER PRODUCT SAFETY COMMISSION FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall...

2014-01-01

428

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2011 CFR

...Section 1507.10 Commercial Practices CONSUMER PRODUCT SAFETY COMMISSION FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall...

2011-01-01

429

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2012 CFR

...Section 1507.10 Commercial Practices CONSUMER PRODUCT SAFETY COMMISSION FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall...

2012-01-01

430

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2013 CFR

...Section 1507.10 Commercial Practices CONSUMER PRODUCT SAFETY COMMISSION FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall...

2013-01-01

431

16 CFR 1507.10 - Rockets with sticks.  

Code of Federal Regulations, 2010 CFR

...Section 1507.10 Commercial Practices CONSUMER PRODUCT SAFETY COMMISSION FEDERAL HAZARDOUS SUBSTANCES ACT REGULATIONS FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall...

2010-01-01

432

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

NASA Technical Reports Server (NTRS)

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

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

2007-01-01

433

Hyperspectral chemical plume quantification and temperature estimation  

NASA Astrophysics Data System (ADS)

Most hyperspectral chemical gaseous plume quantification algorithms assume a priori knowledge of the plume temperature either through direct measurement or an auxiliary temperature estimation approach. In this paper, we propose a new quantification algorithm that can simultaneously estimate the plume strength as well as its temperature. We impose only a mild spatial assumption, that at least one nearby pixel shares the same plume parameters as the target pixel, which we believe will be generally satisfied in practice. Simulations show that the performance loss incurred by estimating both the temperature and plume strength is small, as compared to the case when the plume temperature is known exactly.

Niu, Sidi; Golowich, Steven E.; Ingle, Vinay K.; Manolakis, Dimitris G.

2014-06-01

434

Rocket Engine Numerical Simulator (RENS)  

NASA Technical Reports Server (NTRS)

Work is being done at three universities to help today's NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines. Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming. In 1995, a 3year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews. The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida. In 1996, the Southern University/University of West Florida team conducted a 4day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The discussion was recorded on video and audio tape. Transcriptions of the entire proceedings and an abbreviated video presentation of the discussion highlights are under development. Also in 1996, two additional 3year grants were awarded to conduct parallel efforts that would complement the work being done by Southern University and the University of West Florida. Dr. Prem Bhalla of Jackson State University in Jackson, Mississippi, is developing the architectural framework for RENS. By employing the Rose Rational language and Booch Object Oriented Programming (OOP) technology, Dr. Bhalla is developing the basic structure of RENS by identifying and encoding propulsion system components, their individual characteristics, and cross-functionality and dependencies. Dr. Ruknet Cezzar of Hampton University, located in Hampton, Virginia, began working on the parallelization and objectification of rocket engine analysis and design codes. Dr. Cezzar will use the Turbo C++ OOP language to translate important liquid rocket engine computer codes from FORTRAN and permit their inclusion into the RENS framework being developed at Jackson State University. The Southern University/University of West Florida grant was extended by 1 year to coordinate the conclusion of all three efforts in 1999.

Davidian, Kenneth O.

1997-01-01

435

Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48  

NASA Technical Reports Server (NTRS)

The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

1959-01-01

436

Stationary Plasma Thruster Plume Characteristics  

NASA Technical Reports Server (NTRS)

Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 1.8 x 10 exp 17 cubic meters at a distance of 0.3 m to 1.8 X 10 exp 14 cubic meters at a distance of 4 m from the thruster. The electron temperature over the same region was between 1.7 and 3.5 eV. Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 2.6 within 22 degrees of centerline. The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V. The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties. In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteria

Myers, Roger M.; Manzella, David H.

1994-01-01

437

Behavior of mercury emissions from a commercial coal-fired power plant: the relationship between stack speciation and near-field plume measurements.  

PubMed

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

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

2014-11-18

438

Thermal Analysis for Orbiter and ISS Plume Impingement on International Space Station  

NASA Technical Reports Server (NTRS)

The NASA Reaction Control System (RCS) Plume Model (RPM) is an exhaust plume flow field and impingement heating code that has been updated and applied to components of the International Space Station (ISS). The objective of this study was to use this code to determine if plume environments from either Orbiter PRCS jets or ISS reboost and Attitude Control System (ACS) jets cause thermal issues on ISS component surfaces. This impingement analysis becomes increasingly important as the ISS is being assembled with its first permanent crew scheduled to arrive by the end of fall 2000. By early summer 2001 , the ISS will have a number of major components installed such as the Unity (Node 1), Destiny (Lab Module), Zarya (Functional Cargo Block), and Zvezda (Service Module) along with the P6 solar arrays and radiators and the Z-1 truss. Plume heating to these components has been analyzed with the RPM code as well as additional components for missions beyond Flight 6A such as the Propulsion Module (PM), Mobile Servicing System, Space Station Remote Manipulator System, Node 2, and the Cupola. For the past several years NASA/JSC has been developing the methodology to predict plume heating on ISS components. The RPM code is a modified source flow code with capabilities for scarfed nozzles and intersecting plumes that was developed for the 44 Orbiter RCS jets. This code has been validated by comparison with Shuttle Plume Impingement Flight Experiment (SPIFEX) heat flux and pressure data and with CFD and Method of Characteristics solutions. Previous analyses of plume heating predictions to the ISS using RPM have been reported, but did not consider thermal analysis for the components nor jet-firing histories as the Orbiter approaches the ISS docking ports. The RPM code has since been modified to analyze surface temperatures with a lumped mass approach and also uses jet-firing histories to produce pulsed heating rates. In addition, RPM was modified to include plume heating from ISS jets to ISS components where the jet coordinates are specified, together with the engine cant angle. These latter studies have been focused on the PM with plumes from its reboost and ACS jets impinging on various ISS components and also focused on the Japanese H2 Transfer Vehicle (HTV) with the plumes from its reboost engines impinging on the Cupola window. This paper will present plume heating and surface temperature results on a number of ISS components with and without jet-firing histories, evaluate post-flight data, and describe any potential thermal issues

Rochelle, William C.; Reid, Ethan A.; Carl, Terry L.; Smith, Ries N.; Lumpkin, Forrest E.

2001-01-01

439

IR sensor design insight from missile-plume prediction models  

NASA Astrophysics Data System (ADS)

Modern anti-tank missiles and the requirement of rapid deployment have significantly reduced the use of passive armour in protecting land vehicles. Vehicle survivability is becoming more dependent on sensors, computers and countermeasures to detect and avoid threats. An analysis of missile propellants suggests that missile detection based on plume characteristics alone may be more difficult than anticipated. Currently, the passive detection of missiles depends on signatures with a significant ultraviolet component. This approach is effective in detecting anti-aircraft missiles that rely on powerful motors to pursue high-speed aircraft. The high temperature exhaust from these missiles contains significant levels of carbon dioxide, water and, often, metal oxides such as alumina. The plumes emits strongest in the infrared, 1 to 5micrometers , regions with a significant component of the signature extending into the ultraviolet domain. Many anti-tank missiles do not need the same level of propulsion and radiate significantly less. These low velocity missiles, relying on the destructive force of shaped-charge warhead, are more difficult to detect. There is virtually no ultraviolet component and detection based on UV sensors is impractical. The transition in missile detection from UV to IR is reasonable, based on trends in imaging technology, but from the analysis presented in this paper even IR imagers may have difficulty in detecting missile plumes. This suggests that the emphasis should be placed in the detection of the missile hard body in the longer wavelengths of 8 to 12micrometers . The analysis described in this paper is based on solution of the governing equations of plume physics and chemistry. These models will be used to develop better sensors and threat detection algorithms.

Rapanotti, John L.; Gilbert, Bruno; Richer, Guy; Stowe, Robert

2002-08-01

440

14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.  

Code of Federal Regulations, 2014 CFR

...TRANSPORTATION (CONTINUED) AIR TRAFFIC AND GENERAL OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class...

2014-01-01

441

14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.  

Code of Federal Regulations, 2012 CFR

...TRANSPORTATION (CONTINUED) AIR TRAFFIC AND GENERAL OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class...

2012-01-01

442

14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.  

Code of Federal Regulations, 2013 CFR

...TRANSPORTATION (CONTINUED) AIR TRAFFIC AND GENERAL OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class...

2013-01-01

443

Opacity meter for monitoring exhaust emissions from non-stationary sources  

SciTech Connect

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

Dec, J.E.

2000-02-15

444

Engine exhaust control system and method  

SciTech Connect

This patent describes an exhaust gas control apparatus for an internal combustion engine. It comprises: a rotary fan blade assembly having a hollow hub and plurality of hollow blades, each having a plurality of apertures in a trailing edge; drive means for driving the rotary fan blade assembly; feed means feeding exhaust gases from the engine into the hollow hub and hollow blades; air intake means for feeding intake air to the rotary fan blade assembly from a direction opposite to the direction of flow of the exhaust gases into the hollow hub of the rotary fan blade assembly; exhaust means for exhausting a mixture of air and the exhaust gases; whereby the flow of exhaust gases through the rotary fan blade assembly and out through the exhaust means reduces back-pressure, exhaust noise, exhaust temperature and exhaust pollutants.

Billington, W.G.

1990-04-03

445

Bright polar mesospheric clouds formed by main engine exhaust from the space shuttle's final launch  

NASA Astrophysics Data System (ADS)

The space shuttle launched for the last time on 8 July 2011. As with most shuttle launches, the three main engines injected about 350 t of water vapor between 100 and 115 km off the east coast of the United States during its ascent to orbit. We follow the motion of this exhaust with a variety of satellite and ground-based data sets and find that (1) the shuttle water vapor plume spread out horizontally in all directions over a distance of 3000 to 4000 km in 18 h, (2) a portion of the plume reached northern Europe in 21 h to form polar mesospheric clouds (PMCs) that are brighter than over 99% of all PMCs observed in that region, and (3) the observed altitude dependence of the particle size is reversed with larger particles above smaller particles. We use a one-dimensional cloud formation model initialized with predictions of a plume diffusion model to simulate the unusually bright PMCs. We find that eddy mixing can move the plume water vapor down to the mesopause near 90 km where ice particles can form. If the eddy diffusion coefficient is 400 to 1000 m2/s, the predicted integrated cloud brightness is in agreement with both satellite and ground-based observations of the shuttle PMCs. The propellant mass of the shuttle is about 20% of that from all vehicles launched during the northern 2011 PMC season. We suggest that the brightest PMC population near 70°N is formed by space traffic exhaust.

Stevens, Michael H.; Lossow, Stefan; Fiedler, Jens; Baumgarten, Gerd; Lübken, Franz-Josef; Hallgren, Kristofer; Hartogh, Paul; Randall, Cora E.; Lumpe, Jerry; Bailey, Scott M.; Niciejewski, R.; Meier, R. R.; Plane, John M. C.; Kochenash, Andrew J.; Murtagh, Donal P.; Englert, Christoph R.

2012-10-01

446

Mobile Bay turbidity plume study  

NASA Technical Reports Server (NTRS)

Laboratory and field transmissometer studies on the effect of suspended particulate material upon the appearance of water are reported. Quantitative correlations were developed between remotely sensed image density, optical sea truth data, and actual sediment load. Evaluation of satellite image sea truth data for an offshore plume projects contours of transmissivity for two different tidal phases. Data clearly demonstrate the speed of change and movement of the optical plume for water patterns associated with the mouth of Mobile bay in which relatively clear Gulf of Mexico water enters the bay on the eastern side. Data show that wind stress in excess of 15 knots has a marked impact in producing suspended sediment loads.

Crozier, G. F.

1976-01-01

447

Simple models of tropical plumes  

E-print Network

. Tropical plumes and related phenomena. . . , . III MODEL AND METHOD IV SIMULATIONS AND DISCUSSION a. Overview 16 b. Simulation 0 ? zonally asymmetric initial state. . . . . c. Simulation 1 ? a tropical plume. . . d. Simulation 2 ? forcing 4' poleward... (dashed, interval 107 m2 s-1) at day 8. . 32 26 Simulation 2 divergence (interval 10-s s-1, dashed & 0) at day 8. . . . 32 27 Simulation 2 smoothed Rossby source (interval 10-~ s-1 day-1, dashed & 0) at day 8. . . . . 33 28 Simulation 2 height (solid...

Carrie, Gordon David, d 1960-

1994-01-01

448

Dynamics of laser ablated colliding plumes  

SciTech Connect

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

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

2013-01-15

449

A perfect launch for the Boeing Delta II rocket carrying Stardust  

NASA Technical Reports Server (NTRS)

Billows of exhaust fill Launch Pad 17-A, Cape Canaveral Air Station, as the Boeing Delta II rocket carrying the Stardust spacecraft launches on time. After a 24-hour postponement, the rocket lifted off at 4:04:15 p.m. EST. Stardust is destined for a close encounter with the comet Wild 2 in January 2004. Using a silicon-based substance called aerogel, Stardust will capture comet particles flying off the nucleus of the comet. The spacecraft also will bring back samples of interstellar dust. These materials consist of ancient pre-solar interstellar grains and other remnants left over from the formation of the solar system. Scientists expect their analysis to provide important insights into the evolution of the sun and planets and possibly into the origin of life itself. The collected samples will return to Earth in a sample return capsule to be jettisoned as Stardust swings by Earth in January 2006.

1999-01-01