嶋田 徹

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. We conducted a three-dimensional numerical simulation to ascertain the luminous flame shape around an ignited aluminum particle near the burning surface of composite propellant. The nu- merical simulations were performed with changing pressure. To simulate the luminous flame shape around the ignited aluminum particle, we incorporated vaporized aluminum ejected from the par- ticle surface and simulated the CO2 and H2O gas flow around the particle. Results of numerical simulations show that the cloud of vaporized aluminum ejected from the aluminum particle surface spread around the particle. The cloud shape was streamlined, resembling a raindrop. The cloud shape changed by the pressure and the gas flow around the aluminum particle. The luminous flame diameter estimated from the cloud, and the diameter decreased with increasing pressure.

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In practical usage of conventional hybrid rocket engines, the oxidizer-to-fuel ratio (O/F) shift occurs by either the fuel port diameter increase or throttling because the fuel regression rate is not proportional to the oxidizer mass flux. As a promising technique to eliminate the O/F shift in a wide throttling range, Altering-intensity-Swirling-Oxidizer-Flow-Type (A-SOFT) hybrid rocket engines are proposed. A-SOFTs control O/F, independently of thrust, with the swirl intensity of oxidizer from the injector, as well as the mass flow rate of the oxidizer. In this paper, the increase rates of engine performance caused by O/F shift eliminating technique are evaluated with a vertical launch simulation for single stage sounding rockets. This simulation includes the throat erosion and c* efficiency models which can be affected by O/F shifts. The statistical uncertainty of fuel regression model is also included to evaluate the robustness of A-SOFTs and SOFTs. The increase rates of total impulse and maximum altitude of A-SOFTs compared to SOFTs depends on maximum oxidizer mass flow rate and are about 2% and 4% respectively. The most effective indicators in this evaluation to the flight performance are residuals of propellants and c* efficiency. Owing to the sensitivity of the flight performances to residuals, the fuel regression errors can cause risks of large losses of the highest altitude in SOFTs, and it is found that the feedback control of A-SOFTs have robustness to the fuel regression errors to some extent. c* efficiency dependent on L* is also sensitive to O/F shifts because O/F shifts affect combustion chamber volume and increase of throat area.

© 2016 by Authors. A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view, i.e., problem definition, optimization, and data mining. The primary objective of the present design is that the downrange and the duration time in the lower thermosphere are sufficiently secured for the aurora scientific observation, whereas the initial gross weight is held down to the extent possible. The multidisciplinary design optimization was performed by using a hybrid evolutionary computation. Data mining was also implemented by using a scatter plot matrix. Polypropylene and liquid oxygen with swirling flow are adopted as solid fuel and liquid oxidizer, respectively. The condition of two-time ignitions is assumed in fight sequence on the equation of motion for the three degree of freedom rigid body. Consequently, the design information regarding the tradeoffs, the behaviors of the design variables in the design space to become the nondominated solutions, and the implication of the design variables for the objective functions have been obtained quantitatively. The structurization and visualization of the design space has been implemented in order to observe the effectiveness of the local regions of each design variable. The advantage of extinction-reignition has been indicated.

© 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. In order to clarify combustion phenomena of hybrid rocket engines with multi-section swirl injection method, visualization tests of combustion flames has been conducted. In the present paper, paraffin fuel whose regression rate is high was used, and several types of placement of ports which inject oxygen into combustion chamber were compared. The number of the ports in each section had marginal effect on a combustion phenomenon. On the other hand the distance between each cross-section affected performance and combustion phenomena. In multi-section opposite injection method, the flow toward downstream of combustion chamber was observed. In both methods, enlarging the surface area that high temperature gas flows along was very important to increase regression rate.

© 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Hybrid rocket engine with multi-section swirl injection method was designed and tested for the flight experiments of subscale space plane. Combustion experiments were carried out with high density polyethylene (HDPE) and paraffin fuel and gaseous oxygen with several combustion conditions. Fundamental data for flyable hybrid rocket including usefulness of high pressure CFRP oxygen tank have been revealed. Flight simulation using the measured thrust data clarified the reachable altitude and required thrust to future flight experiments.

© 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. In this paper, first, a theoretical prediction method of regression rates and heat flux to solid fuels of uni-directional vortex injection hybrid rocket engines is developed by introducing a new swirl intensity decline model toward axial direction. Next, a linear propagative relation of heat flux to solid fuels with disturbances of oxidizer mass flux, fuel regression, and initial swirl intensity is derived. The couple of this response model and another unteady response model of solid fuel gasification amplifies oxidizer mass flux disturbance in the form of regression rate oscillation. This is the basic mechanism of low frequency instability unique to hybrid rocket engines. The linear stability analysis for uni-directional vortex types simulating both ILFI amplification source and main stream model is conducted. The result of this analysis shows that uni-directional vortex injection hybrid rocket engines have the same linear unstable mode as axial hybrid rocket engines.

© 2014 by the authors (Yuki Funami and Toru Shimada). Published by the American Institute of Aeronautics and Astronautics, Inc. In order to design hybrid rocket engines, we developed a numerical prediction method to the internal ballistics, such as the characteristic of fuel regression rate. Our model includes quasi-one-dimensional flowfield and one-dimensional thermal conduction into the solid fuel. Besides, the energy-flux balance equation at the solid fuel surface is solved to determine the regression rate. In our previous method, Karabeyoglu’s model was used when evaluating convective heat flux, and only the radiation from gas was considered when evaluating radiative energy flux. In this paper, the model for convective heat transfer is modified considering the velocity-profile peak at the flame location, and soot is also considered as a radiation source. We employ two method; (1) the method where the original convective-heattransfer model is used and where radiative heat transfer is ignored, (2) the method where the modified convective-heat-transfer model is used and where radiative heat transfer from gas and soot is considered. The calculation results are compared with the experimental data in an open literature. As the results, it is confirmed that the order of magnitude of estimated regression rate is the same order of the experimental data. Next, the parametric studies for hybrid rocket design parameters are demonstrated. The three design parameters, which are chamber scale, initial grain temperature and nozzle throat diameter, are employed in the parametric studies. Consequently, we conclude that this method is useful for estimating hybrid rocket internal ballistics.

© 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. For aluminized AP/AN composite propellant, the relation between the agglomerate diameter and pressure was investigated by observing aluminum particle agglomeration in the reaction zone near the burning surface with changing pressure. When the burning rate increased with increasing pressure in aluminized AP/AN composite propellant, the agglomerate diameter decreased with increasing burning rate. We assumed the agglomerate range as the area of the distributed aluminum particles before agglomeration around the burning surface. When the pressure increased, the burning rate increased. Then the agglomerate range decreased. The agglomerate range changes with the burning rate and varies with the temperature profile in the reaction zone. The agglomerate diameter depends on the burning rate and the agglomerate range with changing pressure.

A single-stage launch vehicle with hybrid rocket engine, which uses solid fuel and liquid oxidizer, has been being studied and developed as a next-generation rocket for scientific observation due to the advantages as low cost, safety, re-ignition, and reducing pollution. Thereupon, the knowledge regarding hybrid rocket system has been being gained through the forepart of the conceptual design using design informatics. In the present study, the practical problem defined by using three objective functions and seven design variables for aurora observation is treated so as to contribute the real world using evolutionary computation and data mining for the field of aerospace engineering. The primary objective of the design in the present study is that the down range and the duration time in the lower thermosphere are sufficiently obtained for the aurora scientific observation, whereas the initial gross weight is held down. Investigated solid fuels are five, while liquid oxidizer is considered as liquid oxygen. The condition of single-time ignition is assumed in flight sequence in order to quantitatively investigate the ascendancy of multitime ignition. A hybrid evolutionary computation between the differential evolution and the genetic algorithm is employed for the multidisciplinary design optimization. A self-organizing map is used for the data mining technique in order to extract global design information. Consequently, the design information regarding the tradeoffs among the objective functions, the behaviors of the design variables in the design space to become the nondominated solutions, and the implication of the design variables for the objective functions have been obtained in order to quantitatively differentiate the advantage of hybrid rocket engine in view of the five fuels. Moreover, the next assignments were also revealed.

© 2014, American Institute of Aeronautics and Astronautics Inc. All rights reserved. The numerical analysis about the instability of liquid/dense fluid films under supercritical operating condition is performed on me than fuel. A numerical code for compressible fluid flows accommodated for van der Waals equation of state is developed in order to deal with supercritical fluid and dense fluid layers and has shown good convergence even at a very low Reynolds number flow that is often seen in actual hybrid rocket engines. Linear instability analysis is conducted and shows that an amplification rate has a peak at a certain wave number of initial perturbations. The pertubation becomes unstable as Reynolds number and chamber pressure increase and the instability region of wave number is enlarged when an acceleration body force in the stream wise direction is imposed. A limit cycyle of the amplitude of perturbations is observed at low Reynolds number flows and the instability of dense fluid layers leads to entrainment phenomena at high Reynolds number flows. It is thought that the perturbation which has a peak value of amplification rate is dominating in an actual hybrid rocket engine.

© 2015, American Institute of Aeronautics and Astronautics Inc. All rights reserved. The objective of this study is to clarify inner state of a chamber of the Swirling-Oxidizer-Flow-Type Hybrid Rocket which is one of the types of a Hybrid Rocket by means of numerical fluid analysis. In this study, a numerical code which uses the Large Eddy Simulation as a turbulent modeling and the Flamelet approach as combustion modeling is constructed, and the code is applied to the analysis of the swirling chamber. On this occasion, in order to guarantee an applicability of the results, an experiment of the diffusion flame swirling burner is simulated by the code, and it is confirmed that results of the simulation are well corresponding qualitatively and partially quantitatively to experimental data. Then, a simulation for the chamber of the Swirling-Oxidizer-Flow-Type Hybrid Rocket is done by the numerical code, and it can be obtained that the numerical results are well corresponds qualitatively to visualized data of the experiment. Due to the analysis using this numerical code, structure of flow and flame, distributions of physical quantities and chemical species and state of turbulent eddies are clarified in the chamber of the hybrid rocket.

In this paper, we develop a systematical method for the reduction of chemistry model of hydro-carbon oxygen/air reaction in order to compute the ignition process of boundary layer combustion with a proposed dynamic load balance strategy for the parallel computation of unsteady non-equilibrium chemically reacting flows. Firstly, the reduction method is achieved by omitting Zhu's chemical species determination process, which makes it possible to perform the reduction systematically. By the proposed method the necessary times for chemical reactions of propylene/oxygen and methane/air are reduced to 1/50 and 1/5 of the original each other. Secondly, it is found that, by the dynamic load balance strategy, we can compute the problem 18 times faster than simple load allocation, conventional, approach. Finally, an ignition process of boundary layer combustion of methane/air is calculated by applying the model and computational schemes. We set the initial flow field by using the converged cold-flow solution of air over a methane-injecting porous wall. Injecting high temperature methane gas from a part of the porous wall sets out the ignition simulation. As a result, the first hot spot has appeared at t=0.12 sec near the line of stoichiometry in the boundary layer. Propagation of flame is seen from the hot spot along the line of stoichiometry. The burning speeds are evaluated as 25 and 38 cm/s for the forward moving one and the backward moving one, respectively. They are very close to experimental data (45±5 cm/s). Simulation results also show that the phenomenon occurs under almost constant pressure and enthalpy conditions, and furthermore, the reaction is promoted mainly by the diffusion of radical species. Copyright © 2013 by the International Astronautical Federation. All rights reserved.

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to about 7000 metric tons. Most of the cross-sectional area and mass (97% in low Earth orbit) is concentrated in about 4500 intact abandoned objects plus a further 1000 operational spacecraft. Analyses have shown that the most effective mitigation strategy should focus on the disposal of objects with larger cross-sectional area and mass from densely populated orbits. Recent NASA results have shown that the worldwide adoption of mitigation measures in conjunction with active yearly removal of approximately 0.2-0.5% of the abandoned objects would stabilize the debris population. Targets would have typical masses between 500 and 1000 kg in the case of spacecraft, and of more than 1000 kg for rocket upper stages. In the case of Cosmos-3M second stages, more than one object is located nearly in the same orbital plane. This provides the opportunity of multi-removal missions, more suitable for yearly removal rate and cost reduction needs. This paper deals with the feasibility study of a mission for the active removal of large abandoned objects in low Earth orbit. In particular, a mission is studied in which the removal of two Cosmos-3M second stages, that are numerous in low Earth orbit, is considered. The removal system relies on a Chaser spacecraft which performs rendezvous maneuvers with the two targets. The first Cosmos-3M stage is captured and an autonomous de-orbiting kit, carried by the Chaser, is attached to it. The de-orbiting kit consists of a Hybrid Propulsion Module, which is ignited to perform stage disposal and controlled reentry after Chaser separation. Then, the second Cosmos-3M stage is captured and, in this case, the primary propulsion system of the Chaser is used for the disposal of the mated configuration. Critical mission aspects and related technologies are investigated at a preliminary level. In particular, an innovative electro-adhesive system for target capture, mechanical systems for the hard docking with the target and a hybrid propulsion technology suitable for rendezvous, de-orbiting and controlled reentry operations are analyzed. This is performed on the basis of a preliminary mission profile, in which suitable rendezvous and disposal strategies have been considered and investigated by numerical analysis. A preliminary system mass budget is also performed, showing that the Chaser overall mass is about 1350 kg, including a primary propulsion system of about 300 kg, and a de-orbiting kit with a mass of about 200 kg. The system designed results suitable to be launched with VEGA, actually the cheapest European space launcher.

Theoretical regression rates and combustion response functions in hybrid rocket motors are obtained by using the heat-feedback law that describes the heat flux from flame to fuel surface. In our previous studies, we derived a heat-feedback law using the analogy between momentum and heat transfer within the turbulent boundary layer, and obtained a regression rate expression to be used when the thermal radiation from flame to the fuel surface is neglected. In this study, we attempted to obtain a new regression rate expression using the heat-feedback law when the effects of the radiation cannot be neglected. As a first step, a regression rate expression was obtained when the molecular and turbulent Prandtl numbers are equal to one and then examined the regression-rate characteristics.

It has been clarified by experimental investigations that the regression rate can be improved by swirling injection of gaseous oxidizers. Because it has not been enough that the analysis considering characteristics of Swirling-Oxidizer-Flow-Type Hybrid Rocket, it has been hard to mention that internal state of the rocket has been completely cleared. Therefore, in this study, combustion simulation using LES is performed in order to clarify internal state of Swirling-Oxidizer-Flow-Type Hybrid Rocket.

Low fuel regression rate is fatal disadvantage for hybrid rocket. To overcome this problem, a lot of methods have been proposed. In Kyushu university multi-section swirl injection method has been proposed to increase the fuel regression rate and combustion efficiency. This method generates swirling flow in combustion chamber through injector ports located on the some cross-sections over a fuel grain. High density polyethylene and gaseous oxygen were used as propellant. Multi-section swirl injection method shows twice higher fuel regression rate than that of the conventional method with no swirl in the previous study. In the present study, the effects of the number and the diameter of injector ports was investigated under constant total injector ports area condition. To decrease interference of oxidizer flow, the average regression rate increased at low oxidizer mass flux region.

Recently, hybrid rockets have attracted a lot of interests, because it has main some advantages of low cost, safety, and thrust throttling. On the other hand, launching practical satellites, hybrid rocket has technical problems to overcome, such as low fuel regression rate and low combustion effic iency. In order to improve fuel regression rate and combustion efficiency, a new method with multi-section swirl injection was proposed. In the previous study, it was proved that this method was very effective in increasing fuel regression rate, combustion efficiency, and thrust of hybrid rocket engines. Especially, the fuel regression rate for paraffin fuels with multi-section swirl injection method reaches to about 3 to 10 times higher than that of the no-swirl conventional method. Additionally, deep grooves like erosion are observed on the surface around injector ports of fuel grains after combustion tests. In this paper, combustion tests for several grain types were conducted to clarify influences of difference in the number and diameter of injector ports on the regression rate.