Solid Rocket Motor Nozzle Research Articles

Two-phase flows in the nozzles of solid rocket motors

Physicomathematical models of multivelocity and multitemperature two-phase flows are suggested. The analyses take into account the particle rotation due to off-center collisions in solid rocket motor (SRM) nozzles and the effects of crossing the trajectories of other particles in the combustion chamber. The rotation of particles is treated by means of quasi-one-dim ensional approximation. The effect of the rotation on particle-size distribution and two-phase losses is shown. Three-dimensional two-phase flows in SRM nozzles with rectangular cross sections are considered. The influence of nozzle geometry on the gasdynamic structure of two-phase flow is determined. Axisymmetric calculations are performed for the flows in the combustion chambers and nozzles of SRM. The calculations show the importance of taking into account the intersection of particle trajectories in investigating flows in motors with charges of complex geometry. Nomenclature A = number of particle collisions a = sound velocity c = specific heat cp = specific heat of gas at constant pressure d = particle diameter d4:, = mass mean particle diameter E = total energy of unit volume

Motor and plume particle size measurements in solid propellant micromotors

Particle size distributions were measured in the chamber, nozzle, and plume of a subscale solid propellant rocket motor. A significant reduction in the mean size of the aluminum/aluminum oxide particles occurred within the motor chamber. The mass fraction of small particles (<2 ft) at the nozzle entrance was less than 10%. Also, most particles were smaller than 50 ft, although a few as large as 85 ft were present. In the converging and throat portions of the nozzle it appeared that particle breakup dominated over collision coalescence. Collision coalescence was observed to occur more dominantly in the supersonic nozzle flow. At the nozzle exit the particle mean size was usually less than 2.5 ft. The particle size distributions were bimodal or trimodal, with the larger particles concentrated near the plume centerline. Because of the short nozzle residence times it is not known whether or not these results are also applicable to full-scale motors. The mean size of the small A12O3 particles in the plume edges was less than 0.5 ft in diam, with an index of refraction of 1.64 ± 0.04 (apparently y-Al2O3), independent of propellant composition, motor operating conditions and nozzle geometry.

Space Transportation System Solid Rocket Booster Thrust Vector Control System

The Solid Rocket Booster, Thrust Vector Control (TVC) system was designed in accordance with the following requirements: self-contained power supply, failsafe operation, 20 flight uses after exposure to seawater landings, optimized cost, and component interchangeability. Trade studies were performed which led to the selection of a recirculating hydraulic system powered by Auxiliary Power Units (APU) which drive the hydraulic actuators and gimbal the solid rocket motor nozzle. Other approaches for the system design were studied in arriving at the recirculating hydraulic system powered by an APU. These systems must withstand the imposed environment and be usable for a minimum of 20 Space Transportation System flights with a minimum of refurbishment. The TVC system completed the required qualification and verification tests and is certified for the intended application. Substantiation data include analytical and test data.

The Interaction of the Efflux of Solid Propellants with Nozzle Materials

AbstractProgress in obtaining solid composite propellants with highest specific impulse caused severe problems with materials such as graphite used in the construction of insulation and nozzles in solid rockets. Refractory metals provide relief but at the expense of the desired and gained performance increase in propellants. On the basis of data obtained by experimental firings and thermodynamic analysis it was demonstrated that the primary cause of graphite removal in solid rocket motor nozzles is dependent on pressure and on chemical attack by certain species of the exhaust gas. The most damaging gases are those which react with carbon to form carbon monoxide. The chamber flame temperature does not directly affect graphite removal. Analysis of data indicates a good correlation of propellant composition and oxygen balance of the initial propellant from which the exhaust gases are generated. Application of these findings make it possible to select propellant formulations with minimum corrosive action on graphite.

Grinding the Refractory Metals

A BRIGHT future has been predicted for the refractory metals for aerospace applications. Typical uses include; solid propellent rocket motor nozzles, re‐entry surfaces for gliders and nuclear space power systems (liquid metal containers). Refractory metals are already employed on ram jet engines, hot structure sections capable of supporting aerodynamic and thermal stresses, or as heat shields protecting low temperature structural members.