竹内 伸介

A detonation engine system is successfully demonstrated for the first time in space using sounding rocket S-520-31 of the Japan Aerospace Exploration Agency/Institute of Space Astronautical Science. Detailed flight results of an S-shaped pulse detonation engine (PDE) installed in the rocket are presented herein. The flight is conducted to confirm that the PDE and its system operate at scheduled sequences in space, confirm the reproducibility of the PDE cycle, and despin the rocket around its axis. It is confirmed that the PDE operated successfully for 14 cycles in space. The experimental plateau pressure of 2.0 0.1 MPais 80 3% of the calculated plateau pressure, which suggests that detonation occurred in 14 cycles. The pressure profiles of the cycles are similar, and the pressure integrals are 2.0 0.1 kN⋅ s∕m2, confirming the excellent reproducibility of the PDE cycle. A probability statistical approach assuming a Gaussian distribution is applied to determine the average angular acceleration difference between processes of the PDE operation, mixture supply, and oxygen supply. The results suggested that the PDE despun the rocket via the thrust produced via detonation combustion, which is consistent with a quasi-steady-state model with an accuracy of 101 15%.

To create a new flyable detonation propulsion system, a detonation engine system (DES) that can be stowed in sounding rocket S-520-31 has been developed. This paper focused on the first flight demonstration in the space environment of a DES-integrated rotating detonation engine (RDE) using S-520-31. The flight result was compared with ground-test data to validate its performance. In the flight experiment, the stable combustion of the annulus RDE with a plug-shaped inner nozzle was observed by onboard digital and analog cameras. With a time-averaged mass flow of 182 11 g∕s and an equivalence ratio of 1.2 0.2, the RDE generated a time-averaged thrust of 518 N and a specific impulse of 290 18 s, which is almost identical to the ideal value of constant pressure combustion. Due to the RDE combustion, the angular velocity increased by 5.7 deg ∕s in total, and the time-averaged torque from the rotational component of the exhaust during 6 s of operation was 0.26 N ⋅ m. The high-frequency sampling data identified the detonation frequency during the recorded time as 20 kHz in the flight, which was confirmed by the DES ground test through high-frequency sampling data analysis and high-speed video imaging.

2021年7月27日早朝5：30，JAXA内之浦宇宙空間観測所からデトネーションエンジンシステムを搭載した観測ロケットS-520-31号機が打ち上げられた．高度約200kmにてメタン–酸素推進剤による回転デトネーションエンジン（RDE）の6秒間作動およびパルスデトネーションエンジン（PDE）の2Hz作動を実施した．取得されたフライトデータから，RDE作動で時間平均推力518N，比推力290±18sおよび速度増速量8.0m/sを達成した．PDE作動では1サイクル当たりの圧力時間積分値が5％以内の高精度での繰り返しインパルス生成およびロケット機軸周りのスピンレート減少が確認された．本結果は，地上燃焼試験データとよく一致し，宇宙空間でのデトネーションエンジン作動が実証された．デトネーション波の判定に用いた圧力・加速度センサの高速サンプリングデータおよびRDEプルーム撮影用のデジタルカメラ画像は，JAXA/ISASで開発された再突入データ回収システムRATSにて回収することに成功した．

The application of superelastic alloys to microvibration isolators for a spacecraft was studied. The superelastic alloy in this study is characterized by a wide hysteresis loop in the stress-strain curve, which dissipates a large amount of energy even with a slight change in stress. Prototype struts with damping mechanisms utilizing such characteristics were developed, and an isolation system consisting of the damping struts was investigated. The isolation system was designed to satisfy the requirements of the Space Infrared Telescope for Cosmology and Astrophysics (SPICA), which has several mechanical cryocoolers to cool down the entire telescope to cryogenic temperature. Since the vibrations generated by the cryocoolers can be transmitted to the telescope and deteriorate the pointing stability, SPICA is designed to integrate all cryocoolers into cooler plates and isolate the plates from vibrations with the isolation system. The simulation method to design the isolation system was developed and validated by comparing the estimation with the test results of a bipod with a pair of the damping struts.

This paper presents the results of an S-shaped pulse detonation engine (PDE) ground firing test in the form of a detonation engine system. The world's first technology demonstration of PDE in space using a sounding rocket is planned, and the aim is to control the rocket spin rate in the axial direction using pulsed detonation. The PDE operation at full sequence was successful. The despin rate change of the rocket between continuous oxygen supply and successful PDE operation is expected to be 0.95 deg/s per run. This change in despin rate can be measured by an onboard gyro sensor, making the system flyable. The test results were compared with data from thrust measurement tests conducted in a laboratory, the results of which confirmed the thrust generation under an ambient pressure of 0.5 & PLUSMN;0.1 kPa. The average thrust values in the thrust measurement experiments showed good agreement of 101 & PLUSMN;3% with a quasi-steady-state model introduced to predict the PDE thrust. These results demonstrate the feasibility of the newly developed PDE and its system as the world's first technology demonstration of detonation propulsion in space.

<p>This paper proposes a new concept of a deployable nozzle made of superelastic titanium-alloy sheet. A concrete design of deployable nozzle is determined by FEM analyses based on the condition of a kick motor on a certain launch vehicle. A 1/4-scale model of the nozzle was produced in trial and its storing and deploying test was completed. </p>

This study investigated cumulative damage mechanisms of short fiber type C/SiC under compression. To measure mechanical properties (unloading modulus and permanent strain) before fracture, repeated loading-unloading tests were conducted using a strain gage. Damage was observed to assess characteristics of crack density, length, number, and propagation angle. Furthermore, relations between mechanical properties and damage characteristics were elucidated by application of Basista's equations and by substituting crack densities inferred from damage observations. Stress-strain relations revealed nonlinear behavior. The unloading modulus did not change, but the permanent strain increased. Cracks propagated mainly between fibers, without fiber fracture, connecting other cracks in the direction of orientation 0 deg to 30 deg to the compressive axis. We estimated permanent strain using Basista's equations and damage characteristics. Estimates roughly agreed with experiment results, suggesting that the permanent strain increase is attributable to closed crack sliding and friction caused by increased crack density.

We present an overview of the cryogenic system of the next-generation infrared observatory mission SPICA. One of the most critical requirements for the SPICA mission is to cool the whole science equipment, including the 2.5 m telescope, to below 8 K to reduce the thermal background and enable unprecedented sensitivity in the mid- and far-infrared region. Another requirement is to cool focal plane instruments to achieve superior sensitivity. We adopt the combination of effective radiative cooling and mechanical cryocoolers to accomplish the thermal requirements for SPICA. The radiative cooling system, which consists of a series of radiative shields, is designed to accommodate the telescope in the vertical configuration. We present thermal model analysis results that comply with the requirements to cool the telescope and focal plane instruments.

A novel deployable rocket nozzle utilizing superelasticity was proposed in this study. Ti-4.5Al-3V-2Fe-2Mo alloy (SP-700) sheets were heat-treated to have appropriate alpha/beta ratio so that the sheet shows superelasticity at room temperature. A miniature nozzle model was fabricated through thinning, cutting, and welding processes of the sheets. Folding-deployment tests of the model were conducted in addition to finite element analyses of its folding behavior. The feasibility of the new concept of superelastically-deployable sheet structure was successfully verified.

Alumina (Al2O3) is believed to be the first major condensate to form in the gas outflow from oxygen-rich evolved stars because of the refractoriness and that alpha-Al2O3. (corundum, most stable polymorph) is a potential origin of a 13 mu m feature that appears close to stars. However, no one has directly reproduced the 13 mu m feature experimentally, and it has remained as a noteworthy unidentified infrared band. Here, we report nucleation experiments on Al2O3 nanoparticles monitored by a specially designed infrared spectrometer in the microgravity environment of a sounding rocket. The conditions approximate to those around asymptotic giant branch (AGB) stars. The measured spectra of the nucleated Al2O3 show a sharp feature at a wavelength of 13.55 mu m and comparable in width to that observed near oxygen-rich AGB stars. Our finding that alpha-Al(2)O(3 )nucleates under certain condition provides a solid basis to elaborate condensation models of dust around oxygen-rich evolved stars.

The exploration of energization and radiation in geospace (ERG) satellite, nicknamed "Arase," is the second satellite in a series of small scientific satellites created by the Institute of Space and Astronautical Science of the Japan Aerospace Exploration Agency. It was launched on December 20, 2016, by the Epsilon launch vehicle. The purpose of the ERG project is to investigate how high-energy (over MeV) electrons in the radiation belts surrounding Earth are generated and lost by monitoring the interactions between plasma waves and electrically charged particles. To measure these physical processes in situ, the ERG satellite traverses the heart of the radiation belts. The orbit of the ERG is highly elliptical and varies due to the perturbation force: the apogee altitude is approximately 32,200-32,300 km, and the perigee altitude is 340-440 km. In this study, we introduce the scientific background for this project and four major challenges that need to be addressed to effectively carry out this scientific mission with a small satellite: (1) dealing with harsh environmental conditions in orbit and electromagnetic compatibility issues, (2) spin attitude stabilization and avoiding excitation of the libration by flexible structures, (3) attaining an appropriate balance between the mission requirements and the limited resources of the small satellite, and (4) the adaptation and use of a flexible standardized bus. In this context, we describe the development process and the flight operations for the satellite, which is currently working as designed and obtaining excellent data in its mission.

We present an overview of the thermal and mechanical design of the Payload Module (PLM) of the next- generation infrared astronomy mission Space Infrared Telescope for Cosmology and Astrophysics (SPICA). The primary design goal of PLM is to cool the whole science assembly including a 2.5 m telescope and focal-plane instruments below 8 K. SPICA is thereby expected to have very low background conditions so that it can achieve unprecedented sensitivity in the mid- and far-infrared. PLM also provides the instruments with the 4.8 K and 1.8 K stages to cool their detectors. The SPICA cryogenic system combines passive, effective radiative cooling by multiple thermal shields and active cooling by a series of mechanical cryocoolers. The mechanical cryocoolers are required to provide 40 mW cooling power at 4.8 K and 10 mW at 1.8 K at End-of-Life (EoL). End-to-end performance of the SPICA cryocooler-chain from 300 K to 50 mK was demonstrated under the framework of the ESA CryoChain Core Technology Program (CC-CTP). In this paper, we focus on the recent progress of the thermal and mechanical design of SPICA PLM which is based on the SPICA mission proposal to ESA.

<p>観測ロケットおよび航空機で得られる微小重力環境を利用し、ガス中蒸発法によりアルミナナノ粒子を作製した。干渉計や赤外分光光度計で、過飽和ガスからアルミナのナノ粒子が生成する過程をその場観察した結果について報告する。</p>

The abundant forms in which the major elements in the universe exist have been determined from numerous astronomical observations and meteoritic analyses. Iron (Fe) is an exception, in that only depletion of gaseous Fe has been detected in the interstellar medium, suggesting that Fe is condensed into a solid, possibly the astronomically invisible metal. To determine the primary form of Fe, we replicated the formation of Fe grains in gaseous ejecta of evolved stars by means of microgravity experiments. We found that the sticking probability for the formation of Fe grains is extremely small; only a few atoms will stick per hundred thousand collisions so that homogeneous nucleation of metallic Fe grains is highly ineffective, even in the Fe-rich ejecta of type Ia supernovae. This implies that most Fe is locked up as grains of Fe compounds or as impurities accreted onto other grains in the interstellar medium.

We present the new design of the cryogenic system of the next-generation infrared astronomy mission SPICA under the new framework. The new design employs the V-groove design for radiators, making the best use of the Planck heritage. The new design is based on the ESA-JAXA CDF study (NG-CryoIRTel, CDF-152(A)) with a 2 m telescope, and we modified the CDF design to accommodate the 2.5 m telescope to meet the science requirements of SPICA. The basic design concept of the SPICA cryogenic system is to cool the Science Instrument Assembly (SIA, which is the combination of the telescope and focal-plane instruments) below 8K by the combination of the radiative cooling system and mechanical cryocoolers without any cryogen.

The infrared space telescope SPICA (Space Infrared Telescope for Cosmology and Astrophysics) is a next-generation astronomical project of the Japan Aerospace Exploration Agency, which features a 3 m class and 6 K cryogenically cooled space telescope. This paper outlines the current status for the preliminary structural design of the SPICA payload module. Dedicated studies were conducted for key technologies to enhance the design accuracy of the SPICA cryogenic assembly and mitigate the development risk. One of the results is described for the concept of the on-orbit truss separation mechanisms, which aim to both reduce the heat load from the main truss assembly and isolate the microvibration by changing the natural frequency of the spacecraft. (C) 2015 Society of Photo-Optical Instrumentation Engineers (SPIE)

A new method of strengthening adhesive bonding structures, introducing a slit near the bonding layer, is proposed in order to solve a problem affecting composite pressure vessels; that is, delamination of the bonding layer between the composite and metal, especially around a mouthpiece. The effect of the slit is studied on the basis of an analytical solution, and evaluated using actual double cantilever beam experiments and finite element method calculations. An example of composite pressure vessel design is presented to illustrate the effect of the slit.

The infrared space telescope SPICA, Space Infrared Telescope for Cosmology and Astrophysics, is a next-generation astronomical project of the Japanese Aerospace Exploration Agency (JAXA), which features a 3m-class and 6K cryogenically- cooled space telescope. This paper outlines the current status for the preliminary structural design of the SPICA payload module. Dedicated studies were conducted for key technologies to enhance the design accuracy of the SPICA cryogenic assembly and mitigate the development risk. One of the results is described in this paper for the concept of the on-orbit truss separation mechanisms, which aim to both reduce the heat load from the main truss assembly and isolate the micro-vibration by changing the natural frequency of the spacecraft.

A reliable and evolvable vibration suppression technique using a digital processor powered by energy harvested from a piezoelectric element was developed and investigated. This technique, called a self-powered digital vibration control, benefits from both the sophisticated control logic of the digital processor and a reliable self-powered semiactive element. A system embodying this technique consists of a circuit outline, CPU, and DC/DC converter. Two experiments emulating the synchronized switch damping on inductor (SSDI) technique demonstrated its effectiveness for steady inputs as well as transient ones. [DOI: 10.1115/1.4005027]

We propose a digital autonomous power scavenger with a microprocessor. The proposed system is a completely self-powered one that does not require any external power supply at all, and can thus be used portably at any site. Nevertheless, the digital approach enables the power scavenger to be programmable and thus, it affords some versatility with regard to control schemes. The proposed digitalautonomous system is much more advanced and progressive than clumsy analog-autonomous ones. It can be implemented in multiple-input multiple-output systems to scavenge electrical power from even complicated structural vibrations. We determined the value of the storage capacitance that gives the best balance between scavenging power and consumed power.

We present the SPICA Coronagraphic Instrument (SCI), which has been designed for a concentrated study of extra-solar planets (exoplanets). SPICA mission provides us with a unique opportunity to make high contrast observations because of its large telescope aperture, the simple pupil shape, and the capability for making infrared observations from space. The primary objectives for the SCI are the direct coronagraphic detection and spectroscopy of Jovian exoplanets in infrared, while the monitoring of transiting planets is another important target. The specification and an overview of the design of the instrument are shown. In the SCI, coronagraphic and non-coronagraphic modes are aplicable for both an imaging and a spectroscopy. The core wavelength range and the goal contrast of the coronagraphic mode are 3.5-27 μm, and 10 , respectively. Two complemental designs of binary shaped pupil mask coronagraph are presented. The SCI has capability of simultaneous observations of one target using two channels, a short channel with an InSb detector and a long wavelength channel with a Si:As detector. We also give a report on the current progress in the development of key technologies for the SCI. © 2011 COSPAR. Published by Elsevier Ltd. All rights reserved. 31. -6

An adhesive bonding structure around a metallic mouthpiece of a cryogenic composite tank was analyzed with fracture mechanics. The energy release rate was formulated analytically by considering difference in strain energies in tension and bending before and after crack growth based on a simplified mathematical model. The analytical results were compared by finite element method calculations for five example tanks; there was a fairly good match between the analytical and numerical results. This analysis provides a guideline for the initial optimal design of a cryogenic composite tank based on fracture mechanics.

This paper includes everything which is necessary to make a polymer-lined cryogenic composite tank, i.e. concept, design method, manufacturing process, and recent results of tests. They are also the history of development of the tank in our Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (ISAS/JAXA). At first the concept and production method of the tank is explained. In the production process it is the key that whole liner is formed at one time by use of heat melt bonding. Next the design method based on fracture mechanics is derived. Here an analytical solution of energy release rate for the bonding structure is formulated by considering difference in strain energies in tension and bending before and after crack growth based on a simplified mathematical model. In addition to this derivation, comparison with numerical results and one design optimization sample are also shown. After that the results of tests for actual tank is shown and current status of the tank is discussed. At last conclusion is shown.

This paper proposes a new aerodynamic device, which was designated multi-row-disk (MRD). This device has a cone and stabilizer disks being arranged in the axial direction. This device can arbitrarily change its aerodynamic characteristics by translating stabilizer disks. In the first part of this paper, the effect of several nose shape configurations including the MRD device on the aerodynamic characteristics is reported. By increasing the number of stabilizer disks, zero-lift drag and induced drag can be reduced. It was also found that putting cavities on the conical surface is effective for improving longitudinal static stability. In the second part, the effect of cavity flow instability on pressure and strain oscillation is reported. We drew the design criterion that the configuration of stabilizer disks should be determined not to couple the 1st mode with pressure oscillation frequency, which can be predicted with Rossiter's formula. (c) 2007 Elsevier Ltd. All rights reserved.

This paper reports an experimental study on flow oscillation characteristics of an aerodynamic control device that we have proposed. The device can achieve an enhancement of the aerodynamic control ability and a reduction of the flow instability by adding multiple stabilizer disks to a conventional aerospike so as to divide the flow separation region into multiple cavities. In this device, several axisymmetric cavities are formed. It is well known that pressure oscillation is induced around cavities. In this study, the characteristics of the pressure oscillation of several cavities on a cone surface were investigated experimentally by unsteady pressure measurements in a wind tunnel testing. The conical-cavity pressure oscillation had a feature that the oscillation level is large in case that a length-to-depth ratio of the cavity is large; the oscillation frequency can be predicted by the famous Rossiter formula, which is reported in many previous researches on a single rectangular cavity. It was also found that adding thin disks into the large cavity is effective in the reduction of the pressure oscillation level downstream of the cavities. In addition, disk structural vibration measurements were conducted simultaneously with the unsteady pressure measurements, revealing that a flutterlike vibration could occur when the pressure oscillation frequency agrees with the disk eigenfrequency.

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.

Japan Aerospace Exploration Agency (JAXA) is currently studying "Solar Sail" populsion for future deep space explorations. One of the key technologies to realize a solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utlizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize a 7.5μm-thickness and 50m diameter polyimide membrane, combined with a thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In August 9, 2004, JAXA launched the S-310 sounding rocket, which tested two different shapes of membranes during the zero-gravity flight. The first type of membrane looked like a "clover-leaf", and another is like a "fan". These two membranes, both of them having 10m diameter, were unfolded sequentially during the zero-gravity flight under the free spin condition, and their behavior was observed by onboard cameras. This paper focuses on the "clover-leaf" solar sail, which was fully deployed successfully, and introduces the S-310-34 experiments, and then shows the flight results and postflight evaluations. © 2006 Lister Science. All rights reserved.

A cryogenic tank made of carbon fiber reinforced plastic (CFRP) shell with aluminum thin liner has been designed as a liquid hydrogen (LH2) tank for an ISAS reusable launch vehicle, and the function of it has been proven by repeated flights onboard the test vehicle called reusable vehicle testing (RVT) in October 2003. The liquid hydrogen tank has to be a pressure vessel, because the fuel of the engine of the test vehicle is supplied by fuel pressure. The pressure vessel of a combination of the outer shell of CFRP for strength element at a cryogenic temperature and the inner liner of aluminum for gas barrier has shown excellent weight merit for this purpose. Interfaces such as tank outline shape, bulk capacity, maximum expected operating pressure (MEOP), thermal insulation, pipe arrangement, and measurement of data are also designed to be ready onboard. This research has many aims, not only development of reusable cryogenic composite tank but also the demonstration of repeated operation including thermal cycle and stress cycle, familiarization with test techniques of operation of cryogenic composite tanks, and the accumulation of data for future design of tanks, vehicle structures, safety evaluation, and total operation systems. (c) 2005 Elsevier Ltd. All rights reserved.

Japan Aerospace Exploration Agency (JAXA) is currently studying on the "Solar Sail" propulsion for future deep space explorations. One of the key technologies to realize the solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utilizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize the 7.5μmthickness and 50m diameter polyimide membrane, combined with thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In the August 9, 2004, JAXA has launched the S-310 sounding rocket, which tested two different shapes of membranes during the zerogravity flight. The first type of the membrane looks like a "clover-leaf", and another is like a "fan". These two membranes, both of them have 10m diameter, were unfolded sequentially during the zerogravity flight under the free spin condition, and their behavior was observed by onboard cameras. This paper focuses on the "clover-leaf" solar sail, which was fully deployed successfully, and introduces the S-310-34 experiments, and then shows the flight results and postflight evaluations.

A solar sail is one of propulsion systems for near future interplanetary mission, which utilizes solar pressure for thrust by reflecting sunlight with very large membrane. A solar sail is proposed for an engineering experiment satellite. Although a deployment experiment is necessary in order to verify the solar sail design, it is difficult to perform an experiment of large membrane on the ground because of its lower stiffness. To solve this problem, we used a sounding rocket. On the last August we launched our sounding rocket S-310 No.34 to investigate two solar sail designs « clover type sail (previously it was called reverse umbrella type sail)» and «fan type sail». The rocket was successfully launched and we got some pictures during deployment of the sails. In this paper we will report the result of the experiment.

To prevent satellite failure during inter stage separation, the magnitude of separation shock must be estimated. The Marman Clamp System is a separation joint often used to mount satellites on a launch vehicle. This paper proposes a method of estimating separation shock for the Marman Clamp System. First, experimental results are discussed and the dominant parameters of separation shock are identified for the system. Next, a simple band-mass model is proposed to estimate the system's separation shock. The differences between the model and experimental results are discussed. Radial and axial, but not tangential, separation shocks were estimated well by this method.