嶋田 徹

Discussions are made on the localized erosion of charring ablators used in the expanding part of solid rocket motor nozzles. Such erosion pattern is sometimes seen over liner surface downstream the throat inserts after static firing tests. The major characteristic of the localized erosion is that its shape is groove-like, its erosion amount is very large compared to surrounding region, and its location of occurrence is not simply related to the upstream configuration, such as axial slots or fins of the solid propellant grain. The objective here is to consider the growth mechanism of the localized erosion by reviewing facts reported in the literature on the charring ablators, ablation patterns, and vortical three-dimensional flows in nozzles.

In order to understand the mechanism of the large roll-torque generation in the operating solid rocket motor with axially-slotted propellant grain and the narrow nozzle submergence region, fully three-dimensional Navier-Stokes numerical simulations have been conducted. The several grain configurations are computed and it is found that there are at least two groups of quasi-steady state solutions, one shows large roll torque, and the other shows small one. From the present simulation results, it is observed that the large roll torque is generated due to the interaction of the circling flow around the nozzle inlet with the slot jet exhausting out from the slot end into the aft-end cavity. Although the roll torque evaluated from the computation is one-order higher than that observed in the real fight, the present simulation serves the insight into the qualitative mechanism of the real roll torque generation.

Evaluation of ablation of nozzle wall in solid rocket motor is studied numerically on three solid rocket motors. A coupled analysis of fluid dynamics and surface recession simulates a total ablation amount. The analysis consists of the two-dimensional axisymmetric fluid analysis and the estimation of ablation amount using a correlation equation of surface recession rate. The simulation results of total surface recession amount agree qualitatively with experimental results in Nozzle-A case. The numerical simulations estimate the erosion rate on the safe-side. The effect of shield for reactant gas from nozzle surface reaction is estimated by a simple model. This model works very well for the agreement to experimental results. However, the parameter of the model has to be decided from experimental data. This model is not sufficient for prediction of unknown rocket motor. The coupled analysis was performed on Nozzle-B and C cases in order to understand flow field and erosion mechanism.

無毒で常温貯蔵可能な液体推進剤として亜酸化窒素(N_2O)/エタノールの組合せに着目し,それによる扱い易い液体推進系の実証研究を進めている.当面の目標として大気吸い込み式極超音速推進系の飛行試験に用いる加速用ロケットエンジンへの適用を目指しているが,その低温環境順応性を活かす衛星・探査機搭載推進系への応用も視野に入れている.これまでに,推力700N級の要素試験供試体を用いた燃焼試験を2シリーズ行って,エンジン噴射器設計のための有用なデータと運用特性を取得してきた.併せて,水冷式燃焼器による燃焼器壁面熱流束分布の測定や厚肉のシリカ繊維強化プラスチックSFRP製燃焼器を用いた燃焼試験によって燃焼器への耐熱複合材料適用の可能性も探っている.

Nitrous Oxide (N2O) / ethanol propulsion system is distinguished as the liquid propulsion with non-toxic, user-friendly, and storable bipropellant. The current target of the present study is to build a quick-response and maneuverable main engine of a sounding-rocket like flying test bed which will be applied to the hypersonic air-breathing propulsion researches in the near future. The application to the spacecraft propulsion is also considered due to its compatibility in low-temperature operation environment. Two series of static firing tests were performed with 700 N class gas generator models so far. Current test results showed that valuable design data were collected and operational procedure was verified. Potential of application of composite materials to the combustion chamber was also examined from the chamber wall heat flux data obtained and the result of firing test using a thick SFRP chamber.

A numerical framework is established to predict the erosion amount of nozzle wall in solid rocket motor. Understanding of the flow field and erosion mechanism is very important for the performance of solid rocket motor and the success of the launch. A coupled analysis of fluid dynamics and surface recession simulates total ablation amount. The analysis consists of two-dimensional axisymmetric fluid simulation and estimation of ablation amount using a one-dimensional heat conduction equation with thermal decomposition. The model includes the effect of the mechanical erosion. Three kinds of simulations are carried out to reveal the features of ablation and evaluate the erosion model. Features of nozzle shape can explain heat flux distributions and surface recession amount for each nozzle. The simulation results of total surface recession amount qualitatively recreate the experimental results. The numerical simulations estimate the erosion rate on the safe-side for all nozzles.

Reliability and efficiency in the operation from the assembly of rocket vehicles to the launch of them are considered. Especially, the form of health-confirmation inspection/test after procedures of assembling operations with high efficiency and, at the same time, with reliability maintained is considered. Specifically, the risk when the health-confirmation is not made immediately after an assembling procedure is evaluated by the expectation of the total number of procedures for given failure probability. Copyright 2007 by the IAF or the IAA. All rights reserved.

The acoustic combustion instability of solid rocket motor (SRM) is investigated by computational fluid dynamics and compared with theoretical results. The quasi-one-dimensional (Q1D) Euler equations for the unsteady flow inside the combustion chamber and the equation for the thermal conduction inside the solid propellant are simultaneously solved with a quasi-steady flame model near the burning surface. The Runge-Kutta Discontinuous Galerkin (RKDG) method is used as the platform for the flow simulation and the numerical accuracy study is carried out. The conventional second-order Finite Volume Method is verified to give accurate results by the comparison with the third-order RKDG method. The growth rate versus the nozzle entrance Mach number for the attenuation case shows good agreement with the linear theory. For the growing case, it is shown that agreement is good for small Mach numbers. The results of the stability limit show good agreement with the theory for low Mach number. For higher Mach numbers, the stability-limit curve of the present simulation show the dependency on the imaginary part of the response function. Extension to the axisymmetric problem is straightforward and preliminary results have been obtained. © 2007 by the authors.

This paper describes our experiment and computation of roll torque caused by the internal flow of star-perforated solid rocket motor. The roll torque induced by motor internal flow is known from the early days but is not sufficiently understood among rocket scientists in academia and industry. In the background, there is complexity of a three-dimensional vortical flow in combustion chambers. The roll torque occurring in the launch of the Mu-V rocket was reported by the author in the previous paper (Shimada, IAC-06-C4.3.02, Oct.2006), in which the relation with the internal three-dimensional flow was considered. The roll torque was observed in every seven launches during the early operation period of M-14 motor and it was one-order high compared with that of the aerodynamic and/or of thrust misalignment. The cause of the roll torque was discussed on the possibility of Type-I of Knauber's classification, namely the combustion instability, but it was concluded that the possibility of Type-I was small because the mass efflux from the burning surface was relatively large in M-14 and at the same time, no strong sign of combustion instability existed. In this paper, first, the result of a static firing test of a small motor (diameter of 500mm, burning period of 30 seconds, combustion pressure of about 5 MPa, the maximum thrust of about 50 kN, AP/HTPB/Al+MgAl propellant) is described. In this experiment, the swirling component of exhaust plume and the roll torque acting on the motor have been measured. The swirling flow is measured by the lift force acting on the vane which is installed right downstream the nozzle exit. The result shows the swirling has increased for several seconds after the ignition and attenuated gradually after that. On the other hand, roll torque has been evaluated from the balances of the force and the moment among the gravitational force, the suspension force from the test stand, and the two peripheral loads measured at diametrically either side (right and left) of the motor. The results show that the maximum torque has been about 28 N-m at around several seconds after the ignition in the opposite direction of the swirling flow. The evaluated dimensionless torque coefficient is rather a big value of 1.1 × 10-3. Next, discussion is made on whether the roll torque of M-14 is caused by Type-II, i.e., the internal swirling flow due to the grain shape. The M-14 has seven axial slots in each two grain segments. Because the mass efflux from the slots is larger than the remaining parts of the circumference of the cross section, a jet will flow out from each slot into the central port region. At least two possibilities can be considered; one is symmetric and the other is asymmetric secondary flow field in the cross section. It is only the symmetric case that no torque is generated; in which seven pairs of longitudinal vortices should steadily exist. On the other hand, if the symmetric flow is unstable, these jets might merge into one swirling flow which is supposed to be stabler than the symmetric flow. In this paper verification is sought concerning this supposition employing computational fluid dynamics simulations of the three-dimensional internal flow.

X 線撮影と画像解析を用いて,鉛玉トレーサを含む模擬固体推進薬スラリの二重円筒内部三次元流れ場を可視化した. X 線を互いに直角な二方向から供試体に投影し,透過X 線をフラットパネル検知器とX 線イメージインテンシファイアを用いてビデオに記録した. X 線の相互干渉を抑制することによって,二方向同時撮影が良好に行われた.二方向X 線像の時系列画像データから各トレーサ粒子の空間及び時間的な識別を行い,更に較正用マーカー情報を用いた座標変換を行うことで,トレーサ粒子の刻々の三次元実座標を算出した.これらの手順によって,通常では見ることのできないスラリ流内部の流れ場を可視化し,さらに速度場の推算を行った.

Simulated solid propellant slurry containing lead sphere tracers is experimentally cast into a double circular cylinder container. During the casting, the temperature and the pressure environment has been mimicked to an actual composite solid propellant casting of solid rocket motors. X-rays are projected on to the slurry flow from two directions and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the slurry flow, which is invisible usually, has been estimated.

Three-dimensional, single-phase (equilibrium two-phase) flows inside a solid rocket motor at three burn-back grain configurations are studied by computational fluid dynamics (CFD) analyses of the Reynolds-averaged Navier-Stokes equations (RANS). The major concern is the relationship between the flow field and the circumferentially periodic erosion pattern arising in the inlet region of the nozzle, which will be of help for the better understanding of the surface recession mechanism. Obtained results for the first two cases show that, since the mass flux of slot phase is notably large compared with that of fin phase, a remarkable inter-phase gap in the amount of convective heating appears either in the throat or the exit cone. The peak heating rate appears, commonly to all cases, azimuthally in the slot phase and axially at the expansion ratio of about 0.9 upstream of the throat. The flow, which comes out of a slot into a fin base region, spreads toward the fin central part under the influence of the pressure gradient in the circumferential direction, and forms vortical flow tube of opposite rotation mutually with the flow which swirls out of the next slot. At fin phase, since proportionality relation is accepted between the total mass recession per unit area and the total convective heat mass transfer per unit area, it is considered that corrosion is dominant ablation mechanism. On the other hand, in slot phase, there exists surface recession which cannot be explained only by corrosion around a nozzle inlet nose. This surface recession has a very high possibility of having occurred by abrasion by the aluminum/alumina particles contained in the flow which comes out of axial slot of grain and collides with the TPS surface. It is expected that periodic erosion pattern which synchronized with axial slots observed after static-firing test is the result of such a mechanism ruling over. In both the throat and the exit cone, it is thought irrespective of a phase that the effect of mechanical erosion is very small and corrosion or so-called "chemical attack" is the dominant mechanism of surface recession.

Evaluation of ablation of nozzle wall and natures of three-dimensional flow in solid rocket motor are studied numerically. Three kinds of simulations are carried out to reveal the features of ablation. At first, a coupled analysis of fluid dynamics and surface recession simulates a total ablation amount. The analysis consists of the two-dimensional axisymmetric fluid analysis and the estimation of ablation amount using a correlation equation of surface recession rate. The features of nozzle shape can explain heat flux distributions and surface recession amount for each nozzle. The simulation results of total surface recession amount agree well with experimental results. Most of solid rocket motors use the different materials in the throat and nozzle parts because the throat diameter should be unchanged as possible as it could throughout the whole burning period. Therefore, a backward-facing step is formed at the boundary between those materials because the ablation rates of walls are different. Next, three-dimensional fluid steady analysis in solid rocket motors is carried out. Three-dimensional grid is made based on the results of the axisymmetric coupled analysis. Behind the step, longitudinal vortices and streak of heat flux appear. An effect of shape irregularity at nozzle inlet nose Is considered. The shape irregularity makes the fluctuation with low frequency and high amplitude on the heat flux. Lastly, three-dimensional fluid unsteady analyses around step on the nozzle wall are carried out to reveal the relation between longitudinal vortices and streaks of heat flux. The artificial steps with various heights are set on the nozzle wall. The number of longitudinal vortices depends on the step height. The non-uniformity of shear layer in circumferential direction affects the generation of the longitudinal vortices. The collisions of longitudinal vortices to the wall make the high heat flux.

The static environment of the solid rocket motor chamber was simulated by the improved QPCB. This paper discussed about the combustion products collection method and the particle size distribution of Al/Alumina. Based on the combustion mechanism of the Al agglomeration particles, the shift of the aluminum/alumina particle size distribution in the combustion gas of the solid propellants depending on the traverse distance is estimated experimentally. Further, the correlation between the AP size distribution in the propellant and the particle size distribution of CCP was investigated.

There are unique flow-induced phenomena about solid rocket motors (SRM) whose mechanisms have not been fully understood. The generation of roll torque acting on SRM and peculiar ablation patterns of a nozzle liner surface are taken as examples. By reviewing the open literature, it is found that very few systematic prediction methods exist on these phenomena. Roll torque has been observed during the burning of the first-stage motor of the Mu-V rocket in all six flights since 1997. The cause of the roll torque is sought by evaluating the acoustic effect with mass efflux and combustion response, but sufficiently consistent results have not been obtained. The ablation pattern called striation and cross-hatching has been observed on many specimens in the ablation tests, on reentry , objects after recovery, and on the inner surface of SRM nozzle exit cone. The mechanism of the occurrence of these phenomena is discussed. The existence of the longitudinal vortices is essential for the striation, but for the cross-hatching, whether or not it is an indispensable condition is a pending issue.

The third stage propulsion system, M-34, of Mu-V launch vehicle of Japan Aerospace Exploration Agency (JAXA) is a high performance solid rocket motor which has composite filament winding case and extensible nozzle. In a pre-flight (assembling) operation of Mu-V6, which was launched in July 2005, we found the increase of drag in extension of the nozzle during the function test and made many basic examinations which evaluate the increase and investigate the reason. In the result we determined to exchange the nozzle with the one prepared for next Mu-V-8 in order to make product assurance doubly sure and continued examinations and evaluation. At last we found that the clearance and compression of the gasket, which prevents the leak of high temperature combustion gas through the slit between fixed and extended part of the nozzle, have large effect on the drag force in nozzle extension and modified the nozzle to reuse in the next Mu-V-8 instead of exchanging the nozzle. After modification we attached the nozzle to Mu-V-8 and launched it successfully in February 2006. This paper shows the process and result of examinations, the method to modify the nozzle, and availability for the reuse of the nozzle. This paper also includes the basic description and development and operational history of the extensible nozzle of M-34.

The characteristics of low frequency, bulk-mode combustion instability in solid rocket motors with high-speed internal flow were studied numerically by solving the nonlinear Euler's equation coupled with an unsteady nonlinear combustion equation of solid propellant. Then, the results were compared with those obtained by the so-called L* burner equations, which assumes the uniformity of all physical quantities in the combustion chamber. The linear stability boundaries obtained by both calculations agreed fairly well if the steady-state pressure difference between fore and aft ends in the combustion chamber is taken into account. When the steady-state operating point is near the stability boundary in the unstable region, a small oscillation grows first and then shifts to a limit cycle. On the other hand, when it is far away from the boundary, the oscillation grows until temporary extinction occurs. Frequencies of the L* oscillation obtained by the CFD agreed well with those of the L* burner equations.

M-V is a three-staged solid-propellant satellite launcher developed by the Institute of Space and Astronautical Science (ISAS) of Japan to support various Japanese space science missions. It launches a two-ton class payload into low earth orbit. Although the initial two launches were successful, the third launch of M-V failed to put an X-ray observatory satellite into the orbit because the attitude control was lost for 20 sec in the first-stage flight. From the various data, it turned out that the heat-resistive throat-insert of the nozzle of the first stage fractured and dropped-off, and the hot gas damaged the nozzle throat. And the equipments for attitude control neighboring the nozzle were damaged by the hot gas blew out from the side of the damaged nozzle. After careful review, ISAS and the Space Activity Commission concluded that the fracture of the insert was most likely caused by an embedded or surface crack in the throat insert made of brittle graphite. Because further study came to the conclusion that the performance of up-to-date non-destructive inspection is not enough to surely detect all the detrimental cracks embedded in the graphite billet of the size of this insert, ISAS determined to use inserts made of 3D carbon-carbon composite material. After the redesign and verification by static firing tests, M-V satellite launcher successfully returned to the flight in May 2003, injecting HAYABUSA spacecraft for asteroid sample return mission into a correct orbit.