A preliminary report of results obtained concerning radiation intensities measured with single Geiger tubes carried by Explorer I and III is presented. Resonable cosmic-ray counting rates were obtained for altitudes below about 1000 km. A plot of omnidirectional intensity vs. height in the vicinity of California for the first two weeks in February was obtained, This curve, extrapolated down to altitudes previously reached by rockets, agrees with earlier data, At altitudes greater than about 1100 km, very high counting rates were obtained. (W.D.M,)
The author studied engineering _ science at Harvard University ^ t f t f f lH^ where he received an A.B. in 1946 J ^ B I ^ ^ J L and a Ph.D. in 1949. His thesis flB V | was on an experimental measureI p l l llfcll ment of vibrational relaxation ISfl l ^iTW times in gases. In 1950 he joined ^ ' * 8 U H Wr the shock wave laboratory of the ^ n H E ' ^^i^mi^tJt Physics Department at Princeton ^SK^SJ^L I I L Z ^ ^ 1 ^ University where he continued his l l l l ^ j l H H k ^ research on physical problems in ' l l L ^ ^ • I H H B L gas dynamics. In 1951 he was ^ B « . / | ^ ^ H | | | H u appointed assistant professor of physics; later, associate professor. The subjects on which he has worked extensively include transonic flow, thermal boundary layers, shock structure, chemical kinetics, diffraction of blast waves, and supersonic flow. As one of the editors of the Journal of Fluid Mechanics, he has an intimate and up-to-date knowledge of the problem on which he writes. Dr. Griffith is currently manager of the Flight Sciences Division in the Missile Systems Division of Lockheed Aircraft Corp.
Turbulent heat transfer rates on the aft portion and on the blunt base of a hemisphere cylinder were measured in the 2-7/8 x 2-7/8 GALCIT shock tube over a range of shock Mach numbers between 3. 25 and 5. 1 and initial pressures between 3 and 17 em. Hg. The local Reynolds numbers on the cylindrical afterbody varied between 3.5 x 10^4 and 3.0 x 10^5 per em. A side support was used for the model in order to eliminate the disturbing effect of a rear sting support. The measured turbulent heat transfer rates on the cylindrical portion agreed very well with previous flat plate measurements for small temperature differences, although the ratio of stagnation to surface enthalpy varied between 3 to 8 in the present tests. Only a slight effect of this large variation in h_s/h_w was detected in this range of local Mach numbers, i.e., 1.25 < M_e < 1.5. The measured heat transfer rate on the base indicated that at the center of the base the heat transfer rate is comparable to that on the surface just ahead of the base, while the heat transfer rate falls off to 1/2 to 1/3 of this value towards the rim of the base. This unexpected distribution of heat transfer rate over the base, and particularly the high value at the center, shows the necessity for a careful study of wake phenomena.
Several investigators have attempted to explain and evaluate turbulent burning rates by assuming the turbulent flame brush to be a zone traversed by a wrinkled continuous laminar flame front. Others have argued that the brush consists of a distributed reaction zone characterized by turbulent energy transport and diffusion of active species. A third point of view is that of visualizing the brush as a region containing disconnected flamelets of varying chemical composition. The purpose of the present paper is to state the case for the wrinkled wave concept and to place limits on the regime of its applicability. Experiments indicate that the distributed reaction zone model applies to conditions of incomplete burning within the flame brush. It is suggested that the transition from wrinkled wave to distributed reaction zone corresponds to the breakdown of full turbulent flames through the incidence of holes. New photometric evidence is presented.
After receiving a B.S. degree in Mathematics and Physics from the District of Columbia Teachers' College, the author was employed by the Aerodynamics Section of the National Bureau of Standards. In 1947, he joined the Aerodynamics Department of the Cornell Aeronautical Laboratory where he carried out research in propulsion and high temperature gasdynamics. He received his Ph.D. in Physics from the University of Buffalo in 1955, and is currently head of the Aerophysics Laboratory, Aero physics Department, Space Technology Laboratories, Inc.
The author is Assistant Director of Research of the College of Engineering, New York University. His interest in the minimum weight design of structures has been devoted recently to the efficient application of materials under elevated temperature conditions. In this connection, Dr. Gerard is currently serving on several technical panels of the Materials Advisory Board, National Academy of Sciences.
: The problem of calculating the motion of an incompressible inviscid fluid with a free surface is formulated as a variational principle. The result is in the form of an expansion in powers of the wave height. The theory is specialized to the linear case where it is demonstrated that the variational principle is of the free boundary condition type. A procedure for calculating the natural frequencies and modes of the shallow and the deep tank by applying the Rayleigh-Ritz procedure to solution of the variational formulation is discussed. An approximate procedure for calculating the sloshing modes and frequencies of a tank of intermediate depth in terms of the solutions for shallow and deep tanks is suggested. The results of numerical calculations for conical tank are presented in the form of mode shape diagrams and frequency charts. (Author)
Basic information concerning the free jet method of supersonic development testing is presented. Parameters necessary to specify test hardware, define test facility capabilities, and optimize facilities for this type of testing are discussed. The substantial gains obtainable in a facility's testable range by separately exhausting engine gas flow and free jet air spilled around the engine are defined, as are performance levels of various free jet diffuser types. A nomograph is included which permits rapid estimation of facility requirements and/or capabilities.
Equations were developed relating solid propellant grain geometry to the important ballistic parameters of crosssectional loading density, sliver or tail-off fraction, progressivity ratio, initial surface and web. The equations were then solved by means of high speed computers and the results graphed in a readily usable form. If the loading fraction, sliver, progressivity ratio and web are given, grain designs which meet or most nearly meet given requirements can be found from the graphs. The graphs are useful in determining the effect of slight variations in design parameters. Illustrations of the use of the graphs are given for the internal-burning star and the internal-burning wagon wheel configurations. Calculations have also been made on a modified form of the wagon wheel which permits one, two or three levels of thrust during burning.
Examinat ion of Figs. 3(a) and 3(b) reveals, as is to be expected, t ha t the choice of area ratio has little effect on missile performance when the nozzle weight is a very small port ion of the burnout weight. The importance of correct choice of area ratio increases as the nozzle weight becomes more significant. Nozzle weights, in conventional practice, seldom become large enough in relation to the rest of the inert weight to invalidate the approximation made in Equa t ion [7]. The gains in velocity increment realized by correct choice of area ratio are usually only a few per cent, which m a y be impor tan t in some applications. In other cases, nozzle design m a y still be dictated by structural , cooling, etc., requirements. 2490
