Jinkon Kim

Combustion Characteristics of the Cylindrical Multi-port Grain for Hybrid Rocket Motor

The purpose of this experimental research is to investigate combustion characteristics of cylindrical multi-port grain of a hybrid rocket motor. The physical model of concern includes an oxidizer supply system, a multi-port fuel grain, and a combustor with preand post chamber. Gaseous oxygen and polymer (PE and PMMA) are applied as the oxidizer and the fuel, respectively. In geometric points of view, the effects of the port number and the distance between ports on a regression rate are taken into account. Internal combustion performance was analyzed using smalland large scale motors: one is motor diameter of 50 mm and the other is 96 mm, with oxidizer mass flux ranging from 10 kg/m-sec to 300 kg/msec for both cases. The port number positively shifts the O/F ratio to near its optimum value as well as the oxidizer mass flux to the typical range influenced by pressure effects so that the fuel regression rate becomes fast as the port number increases to a typical number up to 4 ports. Emphasis was placed on merging effects among fuel gain ports having 4 and 5 port to analyze how port geometrical configuration may influence the chamber pressure and thrust drop.

A Study on Burning Rate with Mass Transfer Number (B Number) of Solid Fuel in a Single Port Hybrid Rocket

The widely used general regression rate models in hybrid combustion depend solely on oxidizer mass flux [1]. However, these regression rate models cannot adequately represent well important effect of the thermo-chemical properties of each solid fuel and oxidizer. In this study, the regression rate equation using both the mass flux and mass transfer number (B number), which reflects the thermo-chemistry effect of the propellant, has been examined to determine whether this equation can represent the regression behavior of some polymer based fuels in the absence of radiation. In addition, the calculation of the B number either using the uniform thermo-chemistry property of each propellant combination or using the thermo-chemistry property evaluated at each test case has been performed to observe their influences on the regression rate. Gaseous oxygen and polymer (HDPE, PMMA and PP) are applied as the oxidizer and the fuels, respectively.

Investigation on the Effect of Liquefying Diaphragm in Hybrid Rocket Motors using Paraffin-Based Fuel

Hybrid rocket propulsion experiments using gaseous oxygen as oxidizer and pure paraffin as fuel with a liquefying combustible diaphragm made from 10 wt% LDPE and 90 wt% paraffin were performed. Results of experiments were compared to those of pure paraffin and to diaphragm grain having different step height. Blended combustible diaphragm fuel is chosen for this study to highlight its combustion characteristics and advantages over the metallic diaphragm fuel. Though the segmented grain configuration of the combustible diaphragm is intricate compared to the classical single grain, the capability of combustible diaphragm showed better combustion characte ristics over the single grain in terms of efficiency and performance. In addition, the exposure type diaphragm showed a further improvement in efficiency and characteristic velocity due to the increased turbulent intensity and mixing past the diaphragm. Results of this study accent on the potential of the combustible diaphragm which may be an efficient mixing device for improving the low combustion efficiency of the paraffin based fuel for hybrid rocket motor.

Study on Design- and Operation- Parameters of Supersonic Exhaust Diffuser

A comprehensive approach (theoretical, numerical, and experimental approach) has been conducted to study the designand operationparameters of supersonic exhaust diffusers simulating high altitude condition on the ground. A physical model of concern includes a rocket motor, a vacuum chamber, and a diffuser, which have axisymmetric configurations, using nitrogen gas as a driving fluid. An analysis has been conducted to investigate operation characteristics of a supersonic exhaust diffuser from a flow-development point of view. Emphasis is placed on physical phenomena and several designand operationparameters of the diffuser such as the area ratio of the diffuser to the rocket nozzle, the vacuum chamber size, and the minimum starting pressure of the rocket motor to start the diffuser.

Mass Transfer Number Sensitivity on the Fuel Burning Rate in Hybrid Rockets

The simplest regression rate formula, which depends solely on oxidizer mass flux, originates from Marxman’s theory introduced in the 1960s. This commonly adopted model is still widely used, even though it cannot adequately represent the important effect of thermochemical properties associated to a given specific fuel. In this study, the space–time-averaged regression rate formula taking into account the mass transfer number B is reevaluated to highlight its relative sensitivity with respect to the commonly used simple formula even if B has been known to be a weak function on the regression rate. Polymers (High Density Polyethylene, Polymethylmethacrylate, and Polypropylene) are considered as fuel where theoretical mass flux exponent of 0.75 from the classical theory of Marxman is investigated when applied to the empirical regression rate. Besides, the effect of chamber pressure and use of motor oxidizer to fuel ratio on B sensitivity have been quantified and experimentally analyzed. It is shown that Marxm...

A Study on the Combustion Instability of the Hybrid Rocket Motor with a Diaphragm

In this paper, the main cause on excitation of the combustion instability which may occur in the hybrid rocket motor with a diaphragm was studied. Hybrid rocket motor propulsion tests considering various experimental conditions such as with a diaphragm or not, a diameter of diaphragm, oxidizer mass flow rate, fuel length, etc were performed, and the combustion visualization for the inside of a hybrid rocket motor with a diaphragm was performed. With these experimental results, it was confirmed that the main cause of a large excitation was the hole-tone, and it was shown that the hole-tone model can be predicted experimental primary pressure oscillation frequency quite well.

Combustion Characteristics of Initial Port Diameter Variation of Solid Fuel in Hybrid Rocket Motor

The Combustion characteristic with respect to initial port diameter variation in hybrid rocket motor is investigated using single port cylindrical grain. Experimental studies are performed with GOX and GN2O as oxidizers, and HDPE as fuel. Four initial port diameters of 10 mm, 15 mm, 20 mm and 25 mm were tested to see the effect of initial diameter on the burning rate where the regression rate behaved differently depending on the type of oxidizer. With GOX as oxidizer, the regression rates for the four port diameters are laid closely to each other so that they can be represented by a single empirical correlation i.e., as if the effect of initial port diameter can be negligible on the burning rate. At the contrary with GN2O, the regression rates are hardly represented by a single empirical correlation, and a large scatter of O/F ratio is observed with different initial diameters implying that the effect of initial port diameter can be important on the burning rate. While the behavior of regression rate on the variation of initial port diameter appeared differently depending on the type of oxidizer, the burning rate of fuel at a given oxidizer mass flow rate is decreased with increasing initial port diameter due to the decrease of heat flux from the flame to the solid fuel surface for both the oxidizer. A simple relation between thrust at the nozzle throat and initial geometry characteristic which is composed of oxidizer mass flow rate, initial port diameter and fuel grain length is proposed for the design method of hybrid rocket motor.

Experimental Investigations of the Tapered Fuel Regression Rate of a Hybrid Rocket Motor

The purpose of this research is to investigate a way of improving the non uniformity of the local regression rate in a cylindrical or multiport grain which is an inherent disadvantage of hybrid rocket combustion. A lab scale Hybrid Rocket Motors (HRM) using tapered solid grain (convergent or divergent) having different taper angles is opted for our study. The non uniform local regression rate can not only cause the slivers of solid fuel, but also cause early port merging in the pre and aft section of the multiport grain during combustion. Experiments are conducted to confirm whether taper angles can relax the non uniformity of the local regression rate or influence the performance of lab scale HRMs. In this study, the steeping method is used to measure the accurate local regression rate of tapered grain with taper angles of 1 and 2 degrees. Combustion characteristics such as the overall regression rate and the characteristic velocities are compared with those of cylindrical grain to see taper angles effects on the performance. The possibility of improving the overall regression rate and the non uniformity of local regression rate is discussed.

Influences of B Number Effect on the Burning Rate of Solid Fuel in Single Port Hybrid Rocket

Most of burning rate models used in hybrid combustion depend solely on oxidizer flux. But this empirical relation can not represent well the important effect of the thermo-chemical properties of solid fuel and thereby requires different value of empirical exponent and constant for each fuel considered. In this study, a new burning rate correlation was proposed using the mass transfer number(B number) which encompasses the thermochemistry effect of solid fuel and the aerodynamic effect caused by the combustion on the solid fuel surface where the effect of aerodynamic property in the mass transfer number was studied. The PMMA, PP, and PE were chosen as fuel, and gas oxygen as oxidizer. The new empirical burning rate expression depending on both the oxidizer flux and the mass transfer number was able to predict the burning rate of each fuel with just a single exponent value and constant, and it was found that the aerodynamic effect on the blowing effect did show a minor effect on the burning rate correlation.

Trade-off Evaluation due to Application of Mixing Chamber for Hybrid Rocket-Propulsion System

The intermediate mixing chamber is one of various methods for improving the regression rate and combustion efficiency of the hybrid rocket. The mixing chamber with its non-combustible material makes the propulsion performance increase, but it leads to a low fuel-loading density in the combustion chamber; therefore, this performance-related trade-off between the mixing chamber and the low fuel-loading density was studied. In this study, the trade-off was conducted by comparing the intermediate-mixing-chamber case with a w/o-mixing-chamber case. The small hybrid-sounding rocket is designed with internal ballistics for comparing the rocket length to the weight. In addition, an external ballistic analysis was conducted for comparing the performances of the w/and w/o-mixing-chamber cases. As a result, the intermediate-mixing-chamber case shows that the length and the weight were decreased to 12 % and 8 %, respectively; furthermore, when compared with the normal cases, the estimated altitude result of the w/-mixing-chamber case was increased to approximately 75 m.