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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 757 records · Page 42
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Recent results of the Goddard rocket program for observing stars.

Rockets flown by A. M. Smith have yielded spectra of Alpha Vir, Bl V, Zeta Puppis, O 5f, Zeta Ori, O9.5 Ib, Zeta Oph, O9.5 V and Zeta Per, Bl Ib, at about 0.5 A resolution. Many weak subordinate lines are visible in the range 924 to 1700 A as well as strong resonance lines, some of which have P Cygni profiles. Sharp interstellar lines from C I, C II, O I and Si II are seen. Stecher and Smith have confirmed the presence of CO in the interstellar medium from lines in the ultraviolet spectra of Zeta Oph and Zeta Per. Absolute energy calibrations of the ultraviolet energy distribution of Alpha Leo, Alpha CMa, Gamma Ori and Kappa Ori have been derived with accuracy to about plus or minus 15% over the range 1350-3100 A from rocket spectra at 10 A resolution obtained by D. C. Evans.

Underhill, A. B.↗
NASA NTRS1972

The Viking Orbiter 1975 beryllium INTEREGEN rocket engine assembly.

Description of the conversion of the Mariner 9 rocket engine for Viking Orbiter use. Engine conversion consists of replacing the 40:1 expansion area ratio nozzle with a 60:1 nozzle of the internal regeneratively (INTEREGEN) cooled rocket engine. Five converted engines using nitrogen tetroxide and monomethylhydrazine demonstrated thermal stability during the nominal 2730-sec burn, but experienced difficulty at operating extremes. The thermal stability characteristic was treated in two ways. The first treatment consisted of mapping the operating regime of the engine to determine its safest operating boundaries as regards thermal equilibrium. Six engines were used for this purpose. Two of the six engines were then modified to effect the second approach - i.e., extend the operating regime. The engines were modified by permitting fuel injection into the acoustic cavity.

Martinez, R. S.↗
NASA NTRS1972

A fluidic sounding rocket motor ignition system.

Fluidic sounding rocket motor ignition has been found to be feasible using a system without stored energy and with the complete absence of electrical energy and wiring. The fluidic ignitor is based on a two component aerodynamic resonance heating device called the pneumatic match. Temperatures in excess of 600 C were generated in closed resonance tubes which were excited by a free air jet from a simple convergent nozzle. Using nitrocellulose interface material, ignition of boron potassium nitrate was accomplished with air supply pressures as low as 45 psi. This paper describes an analytical and experimental program which established a fluidic rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source.

Marchese, V. P.↗
NASA NTRS1972

Solid rocket motors for the Space Shuttle booster.

The evolution of the space shuttle booster system is reviewed from its initial concepts based on liquid-propellant reusable boosters to the final selection of recoverable, solid-fuel rocket motors. The rationale associated with each of the several major decisions in the evolution process is discussed. It is shown that the external tank orbiter configuration emerging from the latest studies takes maximum advantage of the solid rocket motor development experience and promises to be the optimum configuration for fulfilling the paramount shuttle program requirements of minimum total development risk within acceptable costs.

Odom, J. B.↗
NASA NTRS1973

Preliminary guided rocket feasibility study.

The feasibility of actively guiding sounding rockets to reduce impact dispersion has been investigated. The theoretical probability of Range Safety thrust termination for several high performance rockets was combined with the cost of acquiring the extended range at White Sands Missile Range (WSMR) to establish a guidance system price ceiling of $20K per flight. Guiding the Black Brant VC for the first five seconds of flight results in sufficient dispersion reduction to impact within the standard range boundaries at WSMR. The guidance system thrust level required to statically control the vehicle to a nominal-wind weighted trajectory for five seconds is between 150-200 pounds. A six-degree-of-freedom trajectory program with guidance simulation capability has been developed and the equations are delineated in this paper.

Nolan, M. B.↗
NASA NTRS1973

Electron density profiles from ionograms - Comparisons with rocket profiles.

From a series of ten rocket flights at Wallops Island conducted by the University of Illinois, detailed electron density profiles are available in the altitude range 60-120 km. Concurrent ionograms from a nearby conventional ionosonde of good sensitivity and precision, have been reduced to N(h) profiles by the NOAA methods. Special attention is given to the needs for accurate ionosonde calibration, proper ionogram interpretation, and corrections for underlying ionization, 'valley' ionization, and effects of lateral-gradients. Proper corrections in the present cases yield profiles by ionogram inversion which agree to within a few percent in density, and a few hundred meters or less in height, with the rocket results.

Wright, J. W.↗
NASA NTRS1973

Investigation of mid-latitude ionospheric currents by combined rocket techniques.

A comprehensive range of neutral atmospheric and simultaneous ionospheric structure measurements were obtained from a Skylark rocket launched at Woomera during evening twilight on a quiet magnetic day. Above 150 km, good agreement was found between the results obtained from three different methods of electric field measurement despite the low field strength (about 1.5 mV/m). In the immediate vicinity of an intense sporadic E layer, the probe measurements indicated large excursions of the electrostatic field amounting to 6.5 mV/m at 105.3 km and 2.7 mV/m at 104.8 km. The calculated ionospheric current system in the vicinity of the rocket trajectory was of similar magnitude to that indicated by local ground-based magnetometers, but was diametrically opposed in direction.

Rees, D.↗
NASA NTRS1973

Reliability and effective thermal conductivity of three metallic-ceramic composite insulating coatings on cooled hydrogen-oxygen rockets

An experimental investigation of the structural integrity and effective thermal conductivity of three metallic-ceramic composite coatings was conducted. These coatings were plasma sprayed onto the combustion side of water-cooled, 12.7-centimeter throat diameter, hydrogen-oxygen rocket thrust chambers operating at 2.07 to 4.14 meganewtons per square meter chamber pressure. The metallic-ceramic composites functioned for six to 17 cycles and for as long as 213 seconds of rocket operations and could have probably provided their insulating properties for many additional cycles. The effective thermal conductivity of all the coatings was in the range of 0.7472 to 4.483 w/(m)(K), which makes the coatings a very effective thermal barrier. Photomicrographic studies of cross-sectioned coolant tubes seem to indicate that the effective thermal conductivity of the coatings is controlled by contact resistance between the particles, as a result of the spraying process, and not the thermal conductivity of the bulk materials.

Price, H. G., Jr.↗
NASA NTRS1973

Space shuttle solid rocket booster recovery system definition, volume 1

The performance requirements, preliminary designs, and development program plans for an airborne recovery system for the space shuttle solid rocket booster are discussed. The analyses performed during the study phase of the program are presented. The basic considerations which established the system configuration are defined. A Monte Carlo statistical technique using random sampling of the probability distribution for the critical water impact parameters was used to determine the failure probability of each solid rocket booster component as functions of impact velocity and component strength capability.

Source record↗
NASA NTRS1973

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

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.↗
NASA NTRS1974

Determination of the availability of appropriate aged flight rocket motors

A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

Martin, P. J.↗
NASA NTRS1973

Extension of a simplified computer program for analysis of solid-propellant rocket motors

A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

Sforzini, R. H.↗
NASA NTRS1973

Prediction of explosive yield and other characteristics of liquid rocket propellant explosions

Work which has been done at the University of Florida in arriving at credible explosive yield values for liquid rocket propellants is presented. The results are based upon logical methods which have been well worked out theoretically and verified through experimental procedures. Three independent methods to predict explosive yield values for liquid rocket propellants are described. All three give the same end result even though they utilize different parameters and procedures. They are: (1) mathematical model; (2) seven chart approach; and (3) critical mass method. A brief description of the methods, how they were derived, how they were applied, and the results which they produced are given. The experimental work used to support and verify the above methods both in the laboratory and in the field with actually explosive mixtures are presented. The methods developed are used and their value demonstrated in analyzing real problems, among them the destruct system of the Saturn 5, and the early configurations of the space shuttle.

Farber, E. A.↗
NASA NTRS1974

MSFC Skylab ATM calibration rocket project

The Apollo Telescope Mount (ATM) Calibration Rocket (CALROC) performances and anomalies encountered are discussed. The performance period included six CALROC flights during the Skylab 2, 3 and 4 missions as well as those rocket flights prior to the Skylab mission which carried CALROC hardware for test purposes. Background material such as project purpose, management, launch facilities, booster and payload configuration is included for better understanding of the CALROC payload and its mission objectives.

Source record↗
NASA NTRS1973

Rocket investigations of the auroral electrojet

Five Nike-Tomahawk rockets were flown to measure perturbations in the magnitude of the geomagnetic field due to auroral electrojets. The dates and locations of the rocket launches are given along with a brief explanation of payloads and instrumentation. Papers published as a result of the project are listed. An abstract is included which outlines the scientific results from one of the flights.

Davis, T. N.↗
NASA NTRS1974

Method of electroforming a rocket chamber

A transpiration cooled rocket chamber is made by forming a porous metal wall on a suitably shaped mandrel. The porous wall may be made of sintered powdered metal, metal fibers sintered on the mandrel or wires woven onto the mandrel and then sintered to bond the interfaces of the wires. Intersecting annular and longitudinal ribs are then electroformed on the porous wall. An interchamber wall having orifices therein is then electroformed over the annular and longitudinal ribs. Parallel longitudinal ribs are then formed on the outside surface of the interchamber wall after which an annular jacket is electroformed over the parallel ribs to form distribution passages therewith. A feed manifold communicating with the distribution passages may be fabricated and welded to the rocket chamber or the feed manifold may be electroformed in place.

Fortini, A.↗
NASA NTRS1973

Hybrid 3-D rocket trajectory program. Part 1: Formulation and analysis. Part 2: Computer programming and user's instruction

Models utilizing various sub-sets of the six degrees of freedom are used in trajectory simulation. A 3-D model with only linear degrees of freedom is especially attractive, since the coefficients for the angular degrees of freedom are the most difficult to determine and the angular equations are the most time consuming for the computer to evaluate. A computer program is developed that uses three separate subsections to predict trajectories. A launch rail subsection is used until the rocket has left its launcher. The program then switches to a special 3-D section which computes motions in two linear and one angular degrees of freedom. When the rocket trims out, the program switches to the standard, three linear degrees of freedom model.

Huang, L. C. P.↗
NASA NTRS1974

Design and calibration of a rocket-borne electron spectrometer for investigation of particle ionization in the nighttime midlatitude E region

An explanation was developed for the formation, near midnight at midlatitudes, of a broad electron density layer extending approximately from 120 to 180 km and usually referred to as the intermediate E layer. The responsible mechanism is believed to be the converging vertical ion drifts resulting from winds of the solar semidiurnal tide. Numerical solutions of the continuity equation appropriate to the intermediate layer is described for particular models of ion drift, diffusion coefficents, and ionization production. Analysis of rocket observations of the layer show that the ionization rate is highly correlated with the planetary geomagnetic index, K sub p. Particle flux measurements support the idea that energetic electrons are the principal source of this ionization. A semiconductor spectrometer experiment for investigation of the particle flux, spectrum, and angular properties was designed and successfully flown on a Nike Apache rocket. A detailed description of the theory, design, and calibration of the experiment and some preliminary results presented.

Voss, H. D.↗