In a standard molecular beam source the maximum attainable intensity in the collimated beam is limited first by the effusion rate through the oven slit, which must be made sufficiently narrow to attain free molecule flow, and second by unfavorable geometrical factors encountered in selecting a collimated beam from random initial velocities. This paper will propose that the first slit be placed in the flow from a miniature high velocity nozzle coaxial with the final beam. The nozzle converts part (∼¾ for a Mach number of 4 in the design for air to be presented) of the random translational and internal energy of the oven gas into directed mass motion. The mass motion provides an initial rough collimation, which improves both the effusion rate and the geometrical factors, indicating a considerable possible beam intensification (by a factor of ∼75 in the sample design). The velocities of the molecules in the final beam are grouped about the initial mass velocity which provides a partial velocity selection. In this paper a theoretical design study and estimates of performance of this type of source are given.

A straightforward theoretical formulation of the longitudinal high frequency rocket stability criteria is presented. Based on Crocco's sensitive combustion time lag theory, the derivation ii considerably simpler, although somewhat less rigorous, than previously published detailed treatments. A series of rocket motor experiments is described, demonstrating that there exists ar upper limit to the chamber length at which each mode of longitudinal high frequency pressure oscillations will occur, and that this limit is accurately predicted by the theory over a wide rang* of mixture ratios on two different injectors. The method used to compare experimental result! with the theoretical stability limit predictions (including a simple experimental technique fo] determination of the sensitive time lag) is described. Phenomenon of Rocket Combustion InstabilityE XCEPT for starting and cutoff transients, rocket motors are designed for steady operation. Two kinds of unsteadiness have nevertheless been observed to be present. First, the combustion in a rocket is always more or less "rough"; that is, random pressure fluctuations of larger or smaller amplitude cannot be avoided. Second, under proper conditions organized pressure oscillations at well-defined frequencies and distributions may appear. The latter form of unsteady operation is given the general title of "combustion instability."There are three recognized types of combustion instability. The first, the "low frequency instability" (1-3,6,7) 5 is due to interaction between the processes taking place in the combustion chamber and the propellant feed system; it is now sufficiently understood and comparitively easy to avoid or cure.The second, the "intermediate frequency instability" (4,5), results from a so-called "entropy wave" produced C3^clically in the chamber and interacting with the exhaust nozzle. Of the three, it is the least frequent.The third, the "high frequency instability," consists of the excitation of acoustic vibrational modes of the combustion chamber. This is by far the most destructive kind of instability and the hardest to control. The present paper deals with certain theoretical and experimental results concerning this third type of instability.A possible cause for high frequency instability was first advanced in 1950 (6). It was speculated that if pressure waves are produced in the chamber, they must produce an effect on Presented at the rates of the various physico-chemical processes leading tc combustion. These modified rates, in turn, by causing periodic intensification or weakening of the local gas evolutior. or heat release, could sustain the pressure waves under the appropriate conditions. A somewhat oversimplified model introduced to express quantitatively the sensitivity of the combustion rates to the pressure oscillations, was able tc demonstrate that high frequency instability could indeed be produced (6). This simple model was refined over the years in order to treat the problem in a more realistic manner. The results of these theoretical t...

A water-cooled gas sampling probe capable of steady-state operation at 15 000°K and one atmosphere has been developed. Temperature is measured by a calorimetric technique which can provide high-accuracy calibration under arcjet conditions. The probe also measures velocity and chemical composition and, because of its small size (∼⅛ in. o.d.), is capable of accurate local measurements in the presence of severe gas-stream property gradients. Calibrations in the partially ionized argon-helium environment of an arcjet exhaust indicated an average agreement with mass and energy balance criteria within ½%, and a standard deviation from the mean of less than 3%.

A theoretical model of a turbine meter operating in the high Reynolds number regime has been formulated to study the effects of retarding torques, inlet velocity profile, blade interference effects, meter geometry, and other factors. A computer program predicts actual rotor speed by numerical integration of the lift and drag forces on the rotor blade. Numerical sample calculations indicated substantial effects due to velocity profile and blade interference. Retarding torques were relatively unimportant in the high Reynolds number range.

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