Kenneth K. Kuo

Hybrid Rocket Investigations at Penn State University's High Pressure Combustion Laboratory: Overview and Recent Results

The Pennsylvania State University has actively pursued hybrid rocket fuel regression rate and combustion research for 15 years. Initial work focused on developing and testing a high-pressure slab-burning hybrid motor with X-ray diagnostics for characterizing the local, real-time regression rates of hydroxyl terminated p olybutadiene (HTPB) with gaseous oxygen. Fuel decomposition and pyrolysis investiga tions were also carried out. Follow-on work is continuing to investigate high regression r ate hybrid fuels with various metal additives in center-perforated hybrid motors using both HTPB and paraffin binders. The addition of aluminum powders to paraffin-based solid-fuel formulations was shown to increase the regression rates by a factor of 4 comp ared to neat HTPB. This regression rate increase corresponds to a mass-burning rate increase ~7 times that of HTPB when the increase in fuel density is considered. This artic le reviews some of the more significant results from previous investigations and presents r ecent data from on-going test programs. Nomenclature

Characterization of the Performance of Paraffin / LiAlH4 Solid Fuels in a Hybrid Rocket System

This investigation examined the burning characteristics of paraffin-based solid-fuel grains doped with various additive percentages (up to 28%) of lithium aluminum hydride (LiAlH4). In addition, the test sequence included examination of a paraffin-wax based fuel formulation containing 10% triethylaluminum and another formulation containing 10% diisobutylaluminum hydride. The fuel grains were cast into paper phenolic tubes and then tested in a cartridge-loaded hybrid rocket system. It was found that under similar test conditions, increased LiAlH4 additive increased the overall chamber pressure throughout the duration of the test, caused by an increase in the ratio of flame temperature to the molecular weight of the products. Due to deposits of unburned and unreacted fuel in downstream sections of the hybrid rocket motor, an accurate correlation between increased additive percentage and regression rate was not able to be found. It was determined that a new set of fuel grain formulations with changes to the overall fuel matrix (e.g., higher melting point wax) and/or changes to the energetic additive particles (e.g., reduced particle size) will allow for more accurate regression rate calculations and more favorable combustion characteristics. Despite the necessary modifications to the fuel formulations, the results from this series of tests showed that nearly all these solid-fuel formulations burned similarly. Qualitative comparisons of each type of fuel formulation proved to be a beneficial method for improving the solid-fuel formulations for future tests for hybrid rocket motor applications.

Instantaneous Regression Rate Determination of a Cylindrical X-Ray Transparent Hybrid Rocket Motor

The determination of the solid fuel regression rate is one of the key steps in hybrid rocket combustion studies. Historically, there is lack of direct regression rate measurements for validation of theoretical models. In practice most mass-burning rates were determined by the net burned mass divided by the test duration that yields an average rate. However, this method does not capture the instantaneous regression behavior. To achieve this, a newly designed X-Ray Transparent Center-perforated (XTC) hybrid rocket motor system has been fabricated and tested to provide the ability to measure the instantaneous solid fuel regression rate using a high-powered real-time X-ray radiography system. Tests have been conducted using hydroxyl-terminated polybutadiene (HTPB) as the baseline solid-fuel formulation. The solid-fuel regression rate can be enhanced by the addition of energetic metal powders. Tests have been conducted using a 13% Silberline aluminum flakes solid fuel formulation in order to evaluate a metalized fuel with the X-ray radiography system. The capability of the visual analysis system to capture instantaneous X-ray radiography images has been demonstrated. Time variations of port dimensions have shown good comparison with calculated regression results from developed numerical code. Differences in recovered fuel burning surfaces were observed from SEM photographs. Large surface roughness, exhibited on the burned surfaces of fuels containing nano-sized aluminum particles, indicates high potential for introducing stronger heat feedback to substrate of solid-fuel and enhance burning rate.

HEAT FLUX AND INTERNAL BALLISTIC CHARACTERIZATION OF A HYBRID ROCKET MOTOR ANALOG

The results of lab-scale hybrid rocket motor test firings and solid fuel pyrolysis experiments were combined with several analytical and numerical techniques to develop a semi-empirical method to analyze and correlate the experimental data with descriptive non-dimensional parameters. The experimental regression rate database was combined with a solid-fuel pyrolysis model and the energy flux balance equation to determine the total heat flux reaching the fuel surface. The component heat fluxes due to convection, radiation from the gas-phase combustion products, and radiation from soot were determined using various methods. The results of the analysis showed that under some conditions the magnitude of the overall radiant heat flux to the fuel surface was quite significant compared to the convective heat. Though CO2 represented the most important radiating gas-phase combustion product, radiation from soot was found to account for 80 to 90% of the total radiant heat flux. Dimensionless regression rate correlations indicated that the classical hybrid boundary layer correlation must be modified to account for the effects of radiation and variable fluid properties across the boundary layer. The Boltzmann number and velocity ratio between the flame and bulk flow were employed to make these corrections. In addition, it was found that the deduced Stanton number can sometimes be larger than the reference Stanton for turbulent flow over a flat plate due to various effects in the hybrid shear flow. The effects of variable transport properties were not found to be important in correlating the data, and no evidence for the existence of rate-limiting chemical kinetics was observed even for relatively large mass fluxes and low pressures. Finally, for the same percent weight mass addition, Alex powder (an ultra-fine aluminum powder, 0.05 to 0.1 |J.m) caused a substantially greater increase in regression rate than did conventional aluminum powder (5 to 15 urn).

Numerical Simulation of Graphite Nozzle Erosion with Parametric Analysis

Nozzle throat erosion is a major problem for solid rocket motors since it causes the degradation in the propulsive performance of solid rocket motors. The AP/HTPB composite propellants used in the rocket motors generate high concentrations of oxidizing species such as H2O, OH, and CO2 in the combustion products at temperatures ranging from 2,700 to 3,200 K for non-metalized propellants. Earlier, the authors utilized a comprehensive numerical program called graphite nozzle erosion minimization code for prediction of graphite nozzle throat erosion rates as a function of pressure and propellant composition. From these studies, it was established that various parameters affect the nozzle thermochemical erosion rate including oxidizing species concentrations, flame temperature, and operating pressure. In addition, the thermal properties of graphite could also affect the nozzle throat erosion rate since these are directly related to the surface temperature at the nozzle throat. In order to assess the relative importance of these parameters in terms of their impact on the nozzle throat erosion rate, a parametric analysis was performed in this study. Each of these parameters was systematically varied while keeping all the remaining parameters constant. Based upon this research, it is concluded that flame temperature can affect the thermochemical erosion rate most, followed by chamber pressure and major oxidizing species concentrations. The mechanisms associated with the influence of these parameters are explored and described. A comparison of predicted results with the available experimental data shows match within 20%. The parametric analysis performed in this research provides an in-depth understanding of the thermochemical erosion process and the controlling steps in the nozzle erosion phenomena.

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.

Characteri stics of Nitromethane for Propulsion Applications

Although hydrazine has been used for space thrusters for many years, recent environmental concerns h ave led to t he need for its replacement for monopropellant candidates with high -performance and reduced toxicity . This has led to study of nitromethane (NM, or CH 3NO 2), which offers potential to meet that need. Nitromethane is a readily -available industria l chemical and has a wide variety of applications, including use as a solvent for chemical processing and analysis. The energetic nature of th is compound also makes it useful as a liquid explosive and as a high -performance additive to fuel for internal co mbustion engines. As a storable high -energy monopropellant, it exhibits high specific impulse and has low toxicity. It also has a great potential to serve as the base for energetic gel propellants. To encourage and enable future propulsion studies, this p aper summarizes much of the previous and current work, and collects many properties necessary for combustion and propulsion studies.

Testing of Hybrid Rocket Fuel Grains at Elevated Temperatures with Swirl Patterns Fabricated Using Rapid Prototyping Technology

Hybrid rocket fuel grains fabricated with rapid prototyping technology enable the use of complex internal structures and port geometries. Using rapid prototyping to print features that introduce flow disturbances and increased surface area can result in improved regression rate and combustion efficiency without the need for difficult machining and casting procedures. In some small-scale hybrid rocket applications, such as small satellites or CubeSats, a lack of robust environmental control might require the motor be used at elevated temperature. Additional increase in regression rate can result from firing the rocket motor with an elevated initial fuel grain temperature, however, due to slumping in liquefying hybrid rocket fuels this is also typically accompanied by a decrease in combustion efficiency. In order to characterize the performance of various fuel grains at elevated temperatures, printed fuel grains with a heterogeneous paraffin and acrylic matrix supplied by The Aerospace Corporation were compared with cast paraffin grains using the Long-Grain CenterPerforated hybrid rocket motor (LGCP) at the Pennsylvania State University’s High Pressure Combustion Laboratory (HPCL). Results from the LGCP testing showed the effects of initial temperature on regression rate and combustion efficiency. The calculated regression rate and combustion efficiency for each fuel grain was compared to previous testing at Penn State and a correlation previously developed for room temperature paraffin fuels. Regression rate increases of over 20% were found for the heated fuel grains, both printed and cast. As expected, the cast paraffin fuel grains experienced a decrease in combustion efficiency as unburned paraffin wax was expelled from the rocket. The printed fuel grains, however, maintained the combustion efficiency of a room temperature cast paraffin fuel grain. The addition of swept honeycomb cell structures utilizing rapid prototyping technology reduced paraffin slumping and allowed more complete combustion at elevated fuel grain temperatures.

Student Investigation of Rapid Prototyping Technology for Hybrid Rocket Motor Fuel Grains

This study reports on a class project designed to introduce students to applications of additive manufacturing in the rocket propulsion industry. Students from the Mechanical Engineering department at Alfred University (AU) have designed, built, and tested smallscale hybrid rocket motors. The project focused on design and fabrication of a test rig that could test many different grain geometries. The students used CAD and CAM software to design this hardware, and had the opportunity to work with a CNC mill and lathe to machine their designs. The primary additive manufacturing technology investigated in this project was 3D printing and was used to make unique hybrid grain designs. The novel ABS hybrid grains were made with additive manufacturing technology from two very different machines, one much more advanced than the other. Through small-scale propulsion testing at AU, we developed a reliable test to compare grains made with the inexpensive printer. To validate our testing and introduce students to a more advanced propulsion test lab, we tested several 3D printed grains at The Pennsylvania State University’s High Pressure Combustion Lab (HPCL). This project was coupled to a senior-level fluid mechanics course taught at AU, and students were introduced to nozzle sizing and simple hybrid rocket motor ballistic predictions and analysis. Difficulties in accurately predicting the performance of complex grain geometry was introduced as well by comparing to data from the various tests. The ultimate goal of this student-led research is to report on the future impact additive manufacturing technology could have on hybrid rocket performance, with suggestions for grain designs.

Comprehensive Three Dimensional Mortar Interior Ballistics Model for 120mm Mortar System with Experimental Validation

A three-dimensional mortar interior ballistic (3D-MIB) model and code have been developed and stage-wise validated with multiple sets of experimental data in close collaboration between The Pennsylvania State Univ. (PSU) and Army Research and Development Engineering Center. This newly developed MIB model and numerical code realistically simulates the combustion and pressurization processes in various components of the 120mm mortar system. Due to the complexity of the overall interior ballistic processes in the mortar propulsion system, the overall problem has been solved in a modular fashion, i.e., simulating each component of the mortar propulsion system separately. The physical processes in the mortar system are two-phase and were simulated by considering both phases as an interpenetrating continuum. Mass and energy fluxes from the flash tube into the granular bed of M1020 ignition cartridge were determined from a semi-empirical technique. For the tail-boom section, a transient one-dimensional two-phase numerical code based on method of characteristics (MOC) was developed and validated by experimental test results. The mortar tube combustion processes were modeled and solved by using a twophase Roe-Pike method with van Leer flux limiter, a fourth-order Runge-Kutta scheme, and an adaptive mesh generator to account for the projectile motion. For each component, the predicted pressure-time traces showed significant pressure wave phenomena, which closely simulated the measured pressure-time traces. The experimental data for the flash tube and ignition cartridge were obtained at PSU whereas the pressure-time traces at the breech-end of the mortar tube were obtained from the tests conducted at Yuma Proving Ground (YPG) and by using an instrumented mortar simulator at Aberdeen Test Center (ATC). The 3DMIB code was also used to simulate the effect of flash tube vent-hole pattern on the pressurewave phenomenon in the ignition cartridge. A comparison of the pressure difference between primer-end and projectile-end locations of the original and modified ignition cartridges with each other showed that the early-phase pressure-wave phenomenon can be significantly reduced with the modified pattern on the flash tube. The flow property distributions predicted by the 3D-MIB for a zero charge increment case are explained in details in this work.