Rocket Exhaust Flow in Tube Launchers

HE structural design of a lightweight, cylindrical-tube rocket launcher requires the calculation of the maximum pressure exerted on the tube wall by the flow exhausting from the rocket nozzle. Thus, an understanding of the resultant flowfield in the launcher is needed. The demand for. more powerful rockets coupled with the requirement for lightweight launchers has recently emphasized the need for understanding the flowfield. Test programs have been conducted to demonstrate the structural integrity of the launcher but have not used the opportunity to obtain flowfield information. Contents A highly underexpanded, supersonic, short-duration jet exhausting from a conical nozzle into a short tube with an inside diameter slightly larger than the nozzle exit characterizes the flowfield of the tube-launched rockets studied. The experimental investigation included eight tests in which double-base, solid-propellant rockets were statically fired into the instrumented tubes and 48 tests in which dry nitrogen was expanded through a conical nozzle into instrumented cylindrical tubes of various lengths. Because the cold-gas tests formed the principal data base, the simulation of the actual rocket exhaust flows was an important consideration. A full-scale facility was designed to simulate the flowfield characteristics of the rocket exhaust. For both types of tests, the reservoir pressure varied rapidly with time, with the maximum value being as high as 350 atm. The test times were relatively short, e.g., approximately 10 msec for rocket firings and approximately 300 msec for the cold-gas tests. In addition, the tube-wall pressure distribution was obtained with a rocket accelerating through a tube launcher. To correlate the experimental data, a theoretical model based upon the method of characteristics was developed for the flowfield from the nozzle exit to a point just downstream of the impingement shock. A sketch of the over-all flowfield based on the theory and substantiated by the acquired data is presented in Fig. 1. The axially symmetric, isentropic flow region is bounded by the in viscid jet boundary and the impingement shock. In actuality, this region may not be isentropic as assumed, if a strong boundary shock develops. However the assumption was found to be valid for the cases considered in the present investigation. Once the jet boundary intersects the tube wall an impingement shock develops as a result of the turning of the gases along the wall. The flow aft of the impingement shock is highly rotational. From the launcher design viewpoint, this is the most important region because maximum wall pressure exists behind the impingement shock. The impingement shock projects to the tube axis as an apparent paraboloid surface. If the tube is sufficiently long, the impingement shock reflects from the axis back to the wall, producing a second pressure peak. Additional reflections take place depending upon the tube length. Steady flow is assumed. However, in reality, the flow is not steady during rocket motor ignition or chamber pressure tailoff. The study concentrates on those data obtained during the "quasi-steady" flow which exists during the interval when the stagnation pressure is a maximum. The flowfield during this time interval is of interest from an engineering standpoint, because the stresses in the tube wall are at a maximum.