The creation and sustainable development of the commercial aircraft original equipment manufacturer (OEM) business need to overcome not only the development of aircraft itself, but also various issues that involve changes in organization, various management processes, human resources, and so on. With respect to these issues, in reviewing past studies, we were not able to find systematic research in the aerospace engineering field or a model based on experiences in the commercial aircraft OEM business. We need to integrate research in new engineering fields and proceed with a comprehensive approach. From the perspective of creating new engineering fields, a comprehensive study covering the characteristics, establishment issues, sustainable development of a commercial aircraft OEM business and the technical development features for commercial aircraft, which is one of the most complex systems that human beings can build, is needed. This paper focuses on definition and how to proceed with system integration, and the relationship between system integration, program, and project management of commercial aircraft development. An integrated product team (IPT) matrix organization has been introduced into many aerospace and defense industries to achieve system integration, but further research is expected to improve the IPT matrix organization, which is balanced by performance and cost.

In this paper, we propose the design of a double fail-safe attitude control system (ACS) architecture for safe artificial meteor release and show evaluation results verified by a hardware-in-the-loop-simulator (HILS). ALE Co., Ltd. and the Space Robotics Laboratory of Tohoku University developed a 60-kg-class microsatellite ALE-1. ALE-1 can create on-demand artificial meteors using a gas-pressurized mass driver to release small balls, which are the source pellets of artificial meteors, for commercial entertainment events and scientific experiments in the upper atmosphere. Released pellets re-enter the upper atmosphere around an altitude of 60 km and burn up after approximately 1/5 of the orbit travel. Since the mass driver is fixed to the satellite body, ACS is responsible for the control of the pellet release direction and satellite position at the time of release. To avoid the insertion of pellets into an unintended orbit at the pellet release, we designed a double fail-safe ACS which consists of three independent CPUs and attitude sensors dedicated to each CPU. The final releasing decision is made by a mechanical unanimity system using solenoid valves which have to be opened in the proper order. Therefore, if any combination of two or fewer breakdowns occur simultaneously, the ACS can immediately abort releasing pellets. We could confirm the correct operation of the pellet release control using our HILS evaluation environment.

Numerical simulations of discharge process in two types of actuator modules were conducted to obtain better understanding of the process of electrohydrodynamic (EHD) force generation toward the low-voltage operation. The charge accumulated on the dielectric surface can be absorbed by the additional downstream-side exposed electrode, which is a component of the adjacent actuator module in a multi-module plasma actuator, when positively-DC-biased negative polarity pulses are applied. The elimination of the surface charge results in EHD force enhancement. The enhanced EHD force is also obtained when a negatively-DC-biased positive polarity pulses are applied. The applied voltage affects the surface charge profile and EHD force enhancement by the downstream-side exposed electrode. Under our simulation condition, the positively-DC-biased negative pulses obtain larger enhancement of EHD force compared to that obtained by the negatively-DC-biased positive pulses.

This study explores the minimum-time attitude maneuver of satellites using Control Moment Gyros (CMG). The CMG system is able to generate high torque and run for longer term than other actuators. However, it is known that the singular input that cannot generate effective torque appears in the optimal control for the minimum-time maneuver of a rigid spacecraft with a pyramid type SGCMG system. In previous studies, a new system equipped with variable speed CMG (VSCMG) to the CMG where the singular input appeared has been proposed. It is shown that the new system can rotate faster than the normal pyramid type SGCMG system in the simulation based on the optimal control law. However, for applications in actual satellites, numerical optimization cannot be used due to the high calculation cost. The objective of this study is to make the new semi-optimal control law for the minimum time attitude maneuver of the satellite equipped with the SGCMG with the variable speed wheel. The new control law is designed using a rule-based feedforward control combined with feedback control. The result by numerical simulation has shown that the proposed control law enables the satellite to achieve faster maneuver without any initial assumptions.

Space debris has become a great concern within the field of space development. In particular, one active debris removal method involves the use of an electrodynamic tether system that exploits the interaction between the geomagnetic flux and plasma electrons. Among the various tethers, tape-shaped tethers show superior survivability during debris removal missions. However, a concentration of stress may be generated at the edge of a damage hole caused by debris collision, resulting in crack advancement and severance of the tethers. Here, we propose an aluminum glass cloth tape (ALGC) tether, in which stress is uniformly distributed. To simulate space debris collisions with tethers, hypervelocity-impact experiments are conducted using a two-stage light-gas gun at ISAS/JAXA. First, lengths of circular or elliptical damage holes are measured and compared between two types of tethers, ALGC tethers and standard aluminum tape tethers. Next, their severance characteristics are defined in terms of tensile force. Additionally, collision simulations are conducted for aluminum tape tethers to allow a detailed comprehension of debris collisions. It is confirmed that, even in the case of two tethers with nearly equal damage dimensions, ALGC tethers outperform aluminum tape tethers in terms of tolerance to tensile force. These versatile ALGC tethers overcome the disadvantages of aluminum tape tethers and should thus exhibit high survivability against debris collisions during removal missions.

ALE-1 is a collaborative microsatellite project between ALE Co., Ltd. and Tohoku University. The mission objective of ALE-1 is to perform an aeronomy study by generating artificial meteors. Due to the uniqueness of the mission and an extremely limited development time, the project posed many challenges in the structural subsystem point of view. To cope with this, several key design concepts were proposed in the preliminary design phase. Moreover, ALE-1 is the first microsatellite developed by Space Robotics Laboratory (SRL) of Tohoku University to be launched by JAXA Epsilon rocket. Being a secondary payload of this rideshare, ALE-1 was launched in the sideway configuration. This is an unconventional way of launching microsatellite and is necessary to consider in the design process. Verification of the structural design of ALE-1 was also conducted by software simulation and environmental tests. Finite element method was implemented in the design process. Environmental test such as shock test, vibration test, and mass property test were also conducted to verify the launch survivability of the satellite structure. This paper aims to use ALE-1 as a demonstration of a microsatellite development with unique mission for Epsilon rocket in the structural subsystem point of view.

Hayabusa2 is a Japanese asteroid explorer launched by the H-IIA rocket on Dec. 3, 2014. After the long cruising with the ion engine system, the spacecraft approached its target body, Ryugu, using optical navigation. It arrived at Home Position (HP), 20 km above the sub-Earth point, on Jun. 27, 2018. The spacecraft is now on its way back to Earth. It will reach the Earth by the end of 2020. During the 1.5-year asteroid proximity phase, we performed many kinds of science observations using the remote sensing instruments. We performed several numbers of descent operations to deploy and lander/rovers, to collect the surface samples, and to create an artificial crater using a new small kinetic impactor. Detailed operation planning of the asteroid proximity phase started after the launch, and lots of training for the proximity phase were conducted from one year before the arrival. This paper summarizes the overview of the asteroid proximity operation plan of Hayabusa2 and its preliminary results.

A deorbit device is required for some microsatellites to meet space debris mitigation guidelines, although very challenging in terms of limited resources and reliability. The authors are conducting research on post-mission disposal (PMD) devices using an electrodynamic tether due to its high efficiency and simplicity. This PMD device consists of a tether deployment system, electron emitter, and integrated control system. The effectiveness of PMD devices were investigated by numerical simulation and ground experiments. The effects on the performance of various parameters were evaluated by numerical simulation in considering such precise models as those used for plasma density and geomagnetic field. We confirmed the effectiveness of the deployment system and the brake system by ground experiments. In this paper, the results of the initial study on a PMD device for microsatellites are presented.

The electrodynamic flow control effect using an applied magnetic field has a potential of innovative technology for aerodynamic force control of atmospheric entry vehicle because of no requirement of mechanical devices. To control this effect and aerodynamic force, we proposed a new approach using exposed electrodes on the vehicle surface. If the induced electric current is bypassed via exposed electrodes, the electromagnetic field could be changed. In this study, the influence of the partial electrical conducting wall, which is consisted of insulating and conducting walls, is numerically investigated. As a result of the MHD simulation, flow field could be changed by the vehicle surface condition. Especially, in the frontal region of the vehicle where strong Lorentz force acts, if the conducting wall is installed to electrically short to the shoulder position, the electric field within this region and the electrodynamic effect is significantly changed. Additionally, stepwise drag variation could be made by selectable shorted electrodes.

The grid design of ion thrusters is currently based on trial and error development and durability tests. These methods make it difficult to rapidly develop optimal thrusters for diverse missions. As a way to address this problem, this paper provides a method of designing grids using numerical simulation with a genetic algorithm. Screen grid thickness, screen grid hole diameter, accel grid thickness, accel grid hole diameter, accel grid potential, and grid gap are optimized for long-life grids. The JIEDI tool is a powerful tool to design grid systems; it is used here as a fitness function to evaluate the lifetime of the grid in an ion thruster. As a result of the optimization, the lifetime of the model developed using the genetic algorithm is about 6.5 times longer than that of the experimental model. The optimization of grid parameters using the Genetic Algorithm and the JIEDI tool shows promise for future design tasks.

Anomalous transport of electrons has considerable impact on the overall performance of Hall thrusters. Numerous studies have shown that high-frequency (on the order of MHz) fluctuations in both the electric field and plasma density are one of the causes of the anomalous transport. This phenomenon has yet to be measured in detail, however, because it is difficult to evaluate the density distribution of neutral particles and the dispersion relation of plasma oscillation in the several MHz range. It was therefore necessary to develop a device to measure the fluctuations in order to elucidate the cause of the anomalous transport. Here we focus on electric field fluctuations in the azimuthal direction inside a Hall thruster, and measure floating potential difference between two Langmuir probes. When a discharge current increases under a strong magnetic field, the azimuthal fluctuation in the frequency of 1-5 MHz is enhanced and fluctuations at ~100 kHz appear. These results suggest that azimuthal fluctuations in the electric field play an important role in the anomalous transport of electrons in the Hall thruster.

In this paper, the aerodynamic characteristics of natural, atmospheric entry objects at hypersonic speeds are numerically studied. As a typical example of such objects, asteroid 101955 Bennu is selected. Orbit simulation is conducted by solving the three-body problem to obtain the initial velocity and angle of atmospheric entry. Based on the constructed shape model, the surface pressure distribution is calculated by the Newtonian method, as well as the force and moment coefficients by pressure integration over the surface. To validate the results of Newtonian analysis, wind tunnel experiment using real asteroid model is conducted. Good agreement is obtained between numerical and experimental results. Then, the atmospheric entry trajectory is predicted. Assuming the initial velocity as 12.36 km/s with an angle of 28.34° to the horizon, the landing time from 120 km is estimated to be 9.86 s. During the landing process of an object with a large ballistic coefficient like Bennu (BC = 6.9 × 10⁵ kg/m²) , the aerodynamic drag force is proved to be small. For such large asteroids, the atmosphere of the Earth is not dense enough to cause significant mass loss and velocity decrease. But for smaller and much more common meteoroids with a small ballistic coefficient, the existence of atmosphere will significantly alter their trajectories and final masses. For such small meteoroids, the landing time required with a large entry velocity v₀ may be longer than that with a small v₀. Given an entry velocity beyond a certain value, the meteoroid will even burn up before it reaches the ground.

Hydrogen and Methane are two fluids that are either used or in discussion as propellants for upper and lower stage rocket engines. The conception of a regenerative cooling system is a crucial part in the design of a rocket engine and so is the detailed knowledge of the coolants behavior and the heat transfer capabilities. Hydrogen is a very efficient and well known cooling fluid whereas the properties of methane as a cooling fluid are intensively investigated nowadays. Experiments were performed with a subscale combustion chamber that is divided into four sectors around the circumference each containing rectangular cooling channels with different aspect ratios. Cryogenic hydrogen and liquid methane were used as cooling fluids. These experiments provide a broad data basis that is used for the validation of CFD simulations. The simulations are capable of predicting wall temperatures for high pressure conditions. Thermal stratification effects that are known to limit cooling properties in high aspect ratio cooling channels arise for both fluids, but the effects are much stronger for hydrogen compared to methane. However in the vicinity to the critical point, when it comes to heat transfer deterioration, the simulations show large deviations to the experimental values.

JAXA, CNES and DLR have decided to join their force to develop and fly a scaled reusable VTVL rocket stage called CALLISTO (Cooperative Action Leading to Launcher Innovation in Stage Toss - back Operations). This vehicle will pave the way for future reusable launch vehicles. One of its main goals is to demonstrate the mastering of key techniques and technologies that will be needed later by operational reusable launch vehicles. Indeed in the frame of CALLISTO numerous challenging technologies are being matured for the specific case of reusable launch vehicles. The challenges of a selection of key technologies developed for CALLISTO are discussed and the solutions retained for CALLISTO are presented.

Functionally graded ablative materials with density gradient were newly developed for the thermal protection system of future space exploration missions. The ablating surface was densified to reduce the amount of surface recession, while the density inside the ablator was reduced with expectation of weight reduction and high heat insulation. Typical Bulk specific gravity was found to be about 0.8. Basic thermal characteristics of the developed ablative material were obtained by conducting heating tests. The heating tests were carried out in the arcjet wind tunnel facility in the Chofu Aerospace Center in JAXA for heat flux of 0.9~4.5 MW/m² and impact pressure of 8.6~31 kPa, and in the Sagamihara campus in JAXA for heat flux of 3.6~14.3 MW/m² and impact pressure of 12~54.3 kPa. The amount of surface recession of ablator was successfully obtained during the heating tests. According to X-ray CT inspection conducted after the heating tests, delamination between layers was not observed inside the test piece. In addition, the present study showed the developed ablative material has a potential to reduce the TPS weight by 33.3 % compared with the Hayabusa ablator, although the recession rate of the present ablator is almost same level as the Hayabusa ablator. From a different point of view, the developed ablative material showed a potential to reduce the recession rate in a relatively wide heating environment compared with a medium density ablator, which has the same specific gravity.

This paper reports the activities resulting from satellite communication lectures conducted by the KOSEN space group in January and August of 2018. Since 2014 the KOSEN space group has been developing two CubeSats to educate students and to complete scientific missions. Consequently, the aim of our lectures is to enable KOSEN students to monitor and operate the CubeSats by satellite communication. In the January lecture, students learned satellite communication theories, and practical training on the assembly and operation of simple receivers. In August, the students learned the theories of digital satellite communication and participated in advanced training using the simple receivers. A survey indicated the participants are interested in receiving signals from a real satellite in space, using the simple self-assembled receivers.

This paper describes the “2017 and 2018 KOSEN Space Camps.” These camps offer introductory space technology programs for students enrolled in the various KOSEN technology colleges in Japan. Since 2015, we have hosted this camp at the Marine Park Niihama. Over the last two years, we evolved and refined the camp's contents to develop a more effective space technology and science education program. In 2017, we successfully held two CanSat development courses: One for beginners to learn how to use onboard computers and sensors and one that is driven more by the participants' imaginations. We tried to make our camps stand out in terms of content variations and load balances including competition. In 2018, we attempted to introduce advanced content such as model CubeSat development and competition. This new trial included online preparation materials from KOSEN's space academia. We will continue to improve the Space Camp content so that in the future, this program may be established as an annual space event in Niihama City.