Andrew C. Cortopassi

Comparison of Nozzle Throat Erosion Behavior in a Solid- Propellant Rocket Motor and a Simulator

The performance deterioration of solid-rocket motors caused by nozzle throat erosion becomes more severe with increased operating pressure from higher rates of heat and mass transfer from the core flow to the nozzle surface. Understanding of the rocket nozzle throat erosion processes and developing methods for mitigation of erosion rate can allow motor operation pressures to be substantially higher than those of the existing propulsion systems. Two test rigs have been utilized in the study of nozzle throat erosion phenomena for G-90 grade graphite; an instrumented solid propellant motor (ISPM) and a solid-propellant rocket motor simulator (RMS). The X-ray translucent nozzle assembly used for the RMS and ISPM allows the real-time imaging of the nozzle-throat station. It also has the feature for incorporating a nozzle boundary-layer control system (NBLCS) to mitigate nozzle-throat erosion rates. The RMS is a gaseous reactant combustor, allows for control of product species compositions, their flow rates, and combustor operating pressure. The erosion process of G-90 graphite was also evaluated in the ISPM using both non-metallized and metallized composite solid propellants. Tests conducted at operating pressures around 21 MPa showed greatly reduced nozzle throat erosion rate when the NBLCS was utilized. A dimensionless nozzle-throat erosion rate correlation was developed in terms of the effective oxidizer mass fraction, chamber pressure, Reynolds number, and relative boundary layer thickness. The correlation equation accurately predicts erosion rate data measured in the RMS and the ISPM for both non-metallized and metallized propellants over a wide range of operating conditions. The calculated erosion rates from the correlation showed agreement within ± 0.05 mm/s of the experimentally determined values.

Design of a Solid Rocket Motor for Characterization of Submerged Nozzle Erosion

Current understanding of physical and chemical processes involved in the erosion of submerged nozzles by highly-aluminized solid propellants is limited. The ability to predict the surface erosion rate of a given carbon-cloth phenolic (CCP) nozzle material is very important for the future design or modification of large solid rocket boosters for space launch applications. Although current erosion codes provide engineering accuracy for nozzle throat erosion rates, calculated rates for the forward surfaces of the submerged nozzle can vary significantly from observed values. The overall objective of this research study under the NASA Constellation University Institutes Project (NASA-CUIP) is to improve the understanding of nozzle erosion and related phenomena. In this work, the design of subscale solid rocket motor was performed based upon engineering analysis of the interior ballistic process and a series of CFD simulations of the flow and heat transfer processes in the region of the submerged nozzle. This motor design allows for the use a realtime X-ray radiography with a high-resolution image intensifier system to obtain submerged nozzle erosion data. From the CFD simulations, the maximum accretion rate of liquid alumina droplets was found to have a level of ~10 kg/s-m 2 in the nose-cone region. Elevated accretion rates in the submerged section of the nozzle were calculated and attributed to the impact of larger particles with higher inertia. These large particles could not follow the combustion product stream to flow out of the nozzle. Development of thermal waves in both the liquid film and the CCP material was investigated. Results showed that their interface temperature can reach 3,000 K in about 1 s. Future test results from this newly designed rocket motor will be highly beneficial for model validation as well as attaining in-depth understanding of interactions between the liquid alumina and nozzle material.

Update: A Subscale Solid Rocket Motor for Characterization of Submerged Nozzle Erosion

Current understanding of physical and chemical processes involved in the erosion of submerged nozzles by highly-aluminized solid propellants is limited. The ability to predict the surface erosion rate of a given carbon-cloth phenolic (CCP) nozzle material is very important for the future design or modification of large solid rocket boosters for space launch applications. Although current erosion codes provide engineering accuracy for nozzle throat erosion rates, calculated rates for the forward surfaces of the submerged nozzle can vary significantly from observed values. The overall objective of this research study under the NASA Constellation University Institutes Project (NASA-CUIP) is to improve the understanding of nozzle erosion and related phenomena. In previous work, the design of a subscale solid rocket motor was performed based upon engineering analysis of the interior ballistic process and a series of CFD simulations of the flow field in both the region of the submerged nozzle and the entire subscale motor. This motor design allows for the use of real-time X-ray radiography with a high-resolution image intensifier system to obtain submerged nozzle erosion data. An update on the design and fabrication of this subscale solid rocket motor is presented in the present work. In addition to this, 3D simulation of the internal flow-field of the rocket motor was performed including the effects of liquid alumina droplets. The modeling of the nozzle surface erosion, coupled with the flow field structure, addresses scientific understanding and characterization of the influence of a liquid layer formed due to deposition of Al2O3/Al droplets on the surface of the converging section of the submerged nozzle. Calculations have been performed which compute the accretion rate of alumina onto the nozzle surface, with accretion rates on the order of 20 kg/m 2 -s. As a part of the overall study, we examine several physicochemical processes on the nozzle surface due to the presence of this molten liquid layer. Future test results from this newly designed rocket motor will be highly beneficial for model validation as well as attaining in-depth understanding of interactions between the liquid alumina and nozzle materials.

Characterization and Detailed Analysis of Regression Behavior for HTPB Solid Fuels Containing High Aluminum Loadings

A. Introduction NASA Marshall Space Flight Center’s Materials and Processes Department, with support from the Propulsion Systems Department, has renewed the development and maintenance of a hybrid test bed for exposing ablative thermal protection materials to an environment similar to that seen in solid rocket motors (SRM). The Solid Fuel Torch (SFT), operated during the Space Shuttle program, utilized gaseous oxygen for oxidizer and an aluminized hydroxylterminated polybutadiene (HTPB) fuel grain to expose a converging section of phenolic material to a 400 psi, 2-phase flow combustion environment. The configuration allows for up to a 2 foot long, 5 inch diameter fuel grain cartridge. Wanting to now test rubber insulation materials with a turn-back feature to mimic the geometry of an aft dome being impinged by alumina particles, the throat area has now been increased by several times to afford flow similarity. Combined with the desire to maintain a higher operating pressure, the oxidizer flow rate is being increased by a factor of 10. Out of these changes has arisen the need to characterize the fuel/oxidizer combination in a higher mass flux condition than has been previously tested at MSFC, and at which the literature has little to no reporting as well. For (especially) metalized fuels, hybrid references have pointed out possible dependence of fuel regression rate on a number of variables:  mass flux, G oxidizer only (G0), or total mass flux (Gtot)  Length, L  Pressure, P  Diameter, D In Sutton, Boardman suggests the form