Toru Shimada

Today developments of nano satellites, whose weight is less than 100 kg, become quite active. As nano satellites are used commercial, inexpensive components, the cost of nano satellites becomes cheap and also the size of subsystems of nano satellites becomes smaller and smaller. The latest nano satellite for single mission becomes very useful for commercial use. Above situations on nano satellites begin to request a low-cost launcher because a combination of low cost nano satellites and low cost launcher can develop nano satellite business market. For this request hybrid rocket is one of the most promising propulsion systems. However, some problems still remain in hybrid rocket such as low fuel regression rate, optimum scale rule and combustion oscillation. The present authors proposed a new method for increase of the fuel regression rate of hybrid rocket. The new method is to introduce swirling flow at multi-sections along the fuel. The new method has been applied for high density polyethylene fuels and paraffin fuels with gaseous oxygen. The results show the new method is quite useful for the increase of the fuel regression rate of hybrid rocket engines. For high density polyethylene fuels the fuel regression rate with multisection swirl injection method shows about 2 to 3 times higher than that of the conventional no-swirl injection method. For paraffin fuels the fuel regression rate with multi-section swirl injection method shows about 3 to 10 times higher than that of the conventional no-swirl injection method with paraffin fuels. The results show the new method of multi-section swirl injection is quite useful both for high density polyethylene fuels and paraffin fuels in order to increase the fuel regression rate of hybrid rocket engines. Copyright © (2012) by the International Astronautical Federation.

It is necessary for designing hybrid rocket engines which use liquefying fuel to understand the behavior of liquid films on the surface of solid fuel. Although it is reported that there is supercritical region inside hybrid rocket engines using liquefying fuel, the process of entrainment phenomena under supercritical operating condition has not been well understood. The present work obtained the steady-state solution for instability analysis of liquid layer as preliminary step. The phenomena in hybrid rockets that use liquefying fuel are formulated and numerical method analysis for van der Waals fluid is shown. As an evaluation of numerical flux, SLAU scheme and Roe scheme for van der Waals gas are calculated. The appropriateness of SLAU scheme for van der Waals gas is discussed by the way of comparing the mass flux in SLAU and the one obtained from Roe scheme accommodated for van der Waals gas. A modification to Roe scheme for van der Waals gas in order to take the change of specific heat at constant volume into account is presented. The steady-state solution with no numerical error is necessary for instability analysis, the steady-state solution obtained from the calculations using SLAU scheme and modified Roe scheme for van der Waals gas are investigated in detail. Copyright © (2012) by the International Astronautical Federation.

The multi-objective genetic algorithm (MOGA) is applied to the multi-disciplinary conceptual design problem for a three-stage launch vehicle (LV) with a hybrid rocket engine (HRE). MOGA is an optimization tool used for multi-objective problems. The Kriging based analysis of variance (ANOVA) and Self-organizing map (SOM), which are data mining methods, are employed for design knowledge discovery. A rocket that can deliver observing microsatellites to the sun-synchronous orbit (SSO) is designed. It consists of an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank and a nozzle. The objective functions considered in this study are to minimize the total mass of the rocket and to maximize the ratio of the payload mass to the total mass. To calculate the thrust and the engine size, the regression rate is estimated based on an empirical model for a paraffin (FT-0070) propellant. Several non-dominated solutions are obtained using MOGA, and design knowledge is discovered for the present hybrid rocket design problem using ANOVA and SOM. As a result, substantial knowledge on the design of an LV with an HRE is obtained for use in space transportation.

A new method with multi-section swirl injection was proposed in order to improve the fuel regression rate of hybrid rockets. The new method was to introduce swirling flow through injector ports, which were placed at several cross-sections along the fuel grain. The key point of the method was to generate swirling flow in the cavity of the fuel grain and provide oxidizer at several cross-sections. In the present study four injector ports were located at each cross-section along the axis of the fuel grain. At each cross section of the fuel grain four injector ports were located at every 90 degrees. The method was applied for high density polyethylene fuels and paraffin fuels (FT-0070) with pressurized gaseous oxygen. The results show the average regression rate of the proposed method is about 2 to 3 times with high density polyethylene fuels and 10 times with paraffin fuels compared with that of the conventional no-swirl injection method. Moreover, some correlations in the multi-section swirl injection method were obtained in the present study. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

In order to design hybrid rocket engines, we have developed a numerical prediction approach to the internal ballistics. The key point is its cost performance. Therefore simple but efficient models are required. Fluid phenomenon and thermal conduction phenomenon in a solid fuel should be treated time-dependently, because characteristic times of these phenomena are longer than those of other phenomena. Besides, they are solved with the energy-flux balance equation at the solid fuel surface to determine the regression rate. It is confirmed that numerical evaluation of time- and space-averaged regression rate is the same order of magnitude as that in experiments. However, the factors n in ṙ=aḠox n differ between calculations and experiments. © 2012 American Institute of Physics.

Boundary-layer combustion, a major characteristic of a hybrid rocket engine, is a complex phenomenon of fluid dynamics and combustion. Its rate-limiting process is diffusion, whereas combustion reactions are generally very fast. One of numerical approaches for this is to solve simultaneously the Navier-Stokes equations with the transport equation for the mixture fraction. Chemical composition of the combustion gas can be determined by solving local chemical equilibrium for a given flow and mixture fraction fields. The governing equations for a diffusion-combustion flow with fast chemistry are characterized by the convective term, the diffusion term, and the chemical equilibrium calculation. As seen from the numerical methods for these, the convective-flux Jacobian and the numerical flux schemes, upwind higher precision approximation and limiter design, and chemical equilibrium calculation method. This study is focused especially on upwind higher precision approximation method. In this paper, by solving test problems such as quasi-one-dimensional hybrid rocket flow, assessment is made on a variety of numerical methods with respect to precision and convergence. © 2012 American Institute of Physics.

For the hybrid engine, it has been proven by ground and flight experiments that the fuel regression characteristic can be improved by tangential injection of the oxidizer. The mechanism, however, of this enhancement has not yet been well-understood. The goal of this study is to establish most efficient way of this type of injection, as well as to better understand the physical mechanism of the effect, by means of Computational Fluid Dynamics. In the chamber, the fuel gas vaporized from fuel grain reacts in swirling flow with the oxidizer to from diffusion frame. For the analysis of the hybrid engine with swirling oxidizer injection, the objectives in this study are to construct the numerical code for diffusion frame in swirling flow and to validate it. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.All rights reserved.

In order to improve fuel regression rate of hybrid rockets, a new method with multisection swirl injection is proposed. The new method is to introduce swirling flow through multi-section swirl injector ports, which are placed at several locations along the fuel grain. The key point of the method is to generate swirling flow in the cavity of the fuel grain and provide oxidizer at several cross-sections. In the present study four injector ports are located at four cross-sections along the axis of the fuel grain. At each cross-section of the fuel grain four injector ports are located at every 90 degrees with deflected angle where injected oxidizer causes swirl at a cross-section in the fuel grain cavity. The method is applied for high density polyethylene fuel and paraffin fuel (FT-0070) with pressurized gaseous oxygen. The results show the average regression rate of the proposed method is about 2 - 3 times with high density polyethylene fuel and 10 times with paraffin fuel compared with that of the conventional no-swirl injection method. © 2011 by Shigeru Aso.

To investigate correlation between the orientation of sub-millimeter AP particles and the local burning rate in a propellant containing high amount of Al particles, the orientation angle data of sub-millimeter AP particles were obtain by using X-ray CT. The orientation data were compared with the local burning rate obtained in previous study. As a result, it is confirmed that if the orientation angle of the coarse AP particles against the burning direction is small, the burning rate become high. This result provides experimental evidence for the supposition that the orientation of AP particle affects the local burning rate. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

This paper describes the development of a numerical simulation system, "Advanced Computer Science on Solid-Rocket-Motor (SRM) Internal Ballistics (ACSSIB)". The objectives of this technology development consist of development of composite-propellant slurry casting-flow simulation, development of local burning-rate correlation with the slurry flow field characteristics, and development of the internal ballistics, i.e., combustion pressure time history, prediction. The ACSSIB have proved itself a promising technology for improvement of SRM reliability and drawn the following conclusions. (1) Hump effect of solid rocket motor combustion is verified by small-scaled motor firing tests and strand burner measurements. (2) Form microscopic observation by microfocus X-ray CT and data deduction by image processing, it is verified that there is a significant correlation between the orientation of coarse AP particles and the burning rate. (3) Development of propellant slurry casting simulation has been successfully conducted. From the casting simulations, it is verified that there is a significant correlation between the angle of the burning direction against the isochrone surface tangent (in plane with the normal) and the burning rate. (4) Development of simulation technique for internal ballistics has been successfully conducted. Simulation results are in good agreement with static firing test results of real motors. Finally, several future technical challenges are identified. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Three dimensional information of sub-millimeter AP particles and local burning rates of composite propellant have been obtained. In this study, to investigate correlation between the orientation of sub-millimeter AP particles and the local burning rate, the data of the local burning rate and the orientation data were arranged and evaluated. As a result, it is suggested that there is the correlation at the propellant that the midweb anomaly occurs. This result provides experimental evidence for the supposition that the orientation of AP particle affects the local burning rate. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

In rocket flights, ionized exhaust plumes from solid rocket motors may interfere with RF transmission under some conditions. In order to clarify the important physical process involved, microwave attenuation and phase delay due to rocket exhaust plumes were measured during sea-level static firing tests conducted on two types of full-scale solid propellant rocket motors. The measured data were analyzed by comparing them with numerical results such as flowfield simulations of exhaust plumes and by employing a detailed analysis of microwave transmission by using a frequency-dependent finite-difference time-domain (FD2TD) method. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma RF interaction. For the actual launch vehicle flight, the transmission process is dominated by the diffraction effect so that we applied a two-dimensional diffraction theory to analyze the communication between a vehicle and a ground station. The attenuation levels estimated using diffraction theory agree with the data recorded in-flight. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

The thrust performance of the solid propellant is improved by adding the metal particles which have good ignition and combustion characteristics. In the reaction zone which is a very thin zone near the burning surface, the metal particles ignite, burn and raise the temperature around them, and the burning rate increases. In order to clarify this mechanism the small solid propellants were combusted and CFD simulations around the metal particles were performed. The temperature histories in the reaction zone, velocity and temperature distributions around the metal particles were obtained. The behavior of the metal particles in the reaction zone was clarified. © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

It has been proven by ground and flight experiments that the fuel regression characteristic of hybrid rocket can be improved by tangential injection of the oxidizer. The mechanism, however, of this enhancement has not yet been well-understood. The objective of this study is to establish most efficient way of this type of injection, as well as to better understand the physical mechanism of the effect, by means of computational fluid dynamics. The report covers the first two phases of the study consisting of five phases. © 2009 by the American Institute of Aeronautics and Astronautics, Inc.