The flat spin of a fixed-wing airplane is usually considered to be dangerous, and should be avoided by all means. However, an airplane in a flat spin descends almost vertically and relatively slowly. It can be used as a means of vertical landing, especially for a small fixed-wing unmanned aerial vehicle. Controlling the impact speed to the ground and touchdown position is necessary to realize the application of flat-spin landing. A large pitch-up moment generated by the combination of an all-flying tail and the gyro precession due to the yawing motion generated by the so-called P-factor of the propeller rotation was found to be effective not only for the rapid entry into flat spin, but also to control the pitch attitude and rate of descent during a flat spin. In this paper, the results regarding the relationships among the deflection angle of the all-flying tail, the rotational speed of the propeller, the spin rate, and the rate of descent obtained by flight tests using a small unmanned aerial vehicle are presented as fundamental knowledge towards the realization of practical flat-spin landing.

This paper aims to assess the risk of mission termination for a large constellation of satellites in a low-Earth orbit. Many large constellations will be deployed to provide broadband network services using thousands of satellites. There is concern that such large constellations will have a serious impact on the long-term sustainability of outer space activities due to the rapid increase in population. First, therefore, the authors conducted an assessment under nominal activities (referred to as “business-as-usual”) on the basis of a prediction by ESA’s MASTER-2009 and NASA standard breakup model 2001 revision. The assessment found that nearly one catastrophic collision may happen in a large constellation, generating more than two million fragments as small as 1 mm in size. Second, the authors conducted a further assessment assuming a hypothetical collision of a satellite in a large constellation using the NASA standard breakup model and a spherical finite element model adopted in ESA’s MASTER-2009. In consequence, another catastrophic collision may happen to a large constellation, generating approximately a half-million fragments as small as 1 mm in size. Therefore, such catastrophic collisions and resulting secondary collisions should be prevented for large constellations.

An efficient implicit time integration method for high-order CFD using the flux reconstruction (FR) approach is studied. The formulation is based on the method called matrix-free defect correction on sub-cells (DECS), which considerably reduces the computational cost of the preconditioned block LU-SGS (BLU-SGS). However, the present method uses the exact FR formulation for the simplified viscous terms in the left-hand-side Jacobian, and it does not explicitly use sub-cell divisions as a result of using the finite-difference formula. The performance of the proposed method is compared with BLU-SGS and DECS. Reductions of the computation time for the convergence are not remarkable, but the slightly improved stability and reduced memory storage suggest the present method as a reliable alternative to the existing methods.

As a post-ISS mission, the construction of Gateway in lunar orbit is planned, and a cargo supply mission to Gateway is also being considered by JAXA. NRHO is a candidate orbit for Gateway, and several methods for the transfer trajectory from Earth are being researched. This study proposes a new transfer technique, the Perilune Rendezvous Method (PRM). With this method, a deceleration maneuver during a lunar swing-by inserts a spacecraft into an elliptical lunar orbit. Then, the spacecraft can enter the NRHO with a small velocity increment by waiting for when the orbital plane coincides with the NRHO. Compared with the Lunar Fly-by Method (LFM) adopted by NASA for manned flight missions, the PRM can reduce the required ΔV, but the transition requires a longer period. This method expands the options for cargo supply and manned flight missions. In this paper, the transfer trajectory from Earth to the NRHO is designed using both methods. Furthermore, the results of the ΔV optimization analysis are compared with the velocity increment and the transition period as variables. This leads to a discussion of the Jacobi integrals of the analysis against the theoretical minimum velocity increment.

In recent years, many advanced space missions, such as on-orbit construction and on-orbit repair, have been proposed, and it is expected that free-flying space robots will need to be more intelligent, versatile, and dexterous in the future. One of the technologies required for such advanced missions is attitude control that actively uses the nonholonomic properties of attitude motion. Although some control methods have been proposed in previous studies, their applicability is limited. Therefore, it is necessary to develop an analytical tool that can be applied to advanced space robots. In the present study, analytical investigations of attitude motion are conducted for future free-flying space robots that actively use nonholonomic properties. In order to derive the solution in a general form, we use a rotational matrix kinematics equation and solve this equation using Magnus expansion. Owing to this analysis, the derived solution preserves the structure of the Lie group of three-dimensional attitude motion. As a practical solution, we derive the solution for "rectilinear actuation," which contributes to concise formulation. We also apply the derived solution to the analysis of the nonholonomy of attitude motion and confirm that the analytical solution can be useful to comprehensively understand the nonholonomic system.

Air traffic flies from Asia to North America via the North Pacific (NOPAC) route system in oceanic airspace. The cruise altitudes of NOPAC routes are assigned on a first-come first-served basis. As a result, when an overflight and Japan departure flight compete for a cruise altitude, the former tends to receive its requested altitude, while the latter is either delayed to receive its requested altitude or allocated a less favorable altitude. This arguably reduces the fairness and overall efficiency of the NOPAC route system. To balance the benefits between Japan departure flights and overflights, alternative strategies are proposed for allocating the cruise altitudes and departure delays more equitably. An optimization method is presented for evaluating such strategies in terms of minimizing departure delays and the difference between requested and assigned altitudes for both Japan departure flights and overflights without creating conflicts. Simulation experiments showed that the proposed strategies can optimize altitude assignment and revealed the tradeoffs involved with different strategies. The proposed strategies can contribute to realizing collaborative decision-making for air traffic management of the NOPAC route system by balancing the potential benefits for both overflights and Japan departure flights.

This paper focuses on the deployment dynamics of space membrane structures, placing an emphasis on membranes folded and deployed in a manner in which inter-membrane interaction must be taken into account in order to produce accurate results. This is motivated by the ongoing development of the HELIOS mission; a spacecraft component scheduled to launch in 2022, which is expected to have high membrane entanglements during its deployment. Contact mechanics are added to a multi-particle method-based model to be able to simulate HELIOS-like membrane deployment configurations. The resulting deployment simulation is analyzed using the work rate of the deployment structure and membrane stress during deployment. Several deployment configurations are tested, and factors that amplify or suppress entanglement phenomena are identified.

“Nose-first entry” flight has been proposed as one of the methods of return flight of a vertical take-off and vertical landing reusable rocket using engine thrust for vertical landing. In this flight method, the engine exhaust jet opposes the free-stream during the turnover maneuver and landing, causing concern that the aerodynamic force acting on the vehicle changes due to a complicated flow field made by the interaction between the exhaust plume and free-stream. A slender-body model that can eject a supersonic jet was studied in a low-speed wind tunnel to characterize the flow. The aerodynamic forces were measured, and the flow pattern on the surface of the model was visualized using the oil-flow technique. The results indicate that the axial force decreases in the low angle-of-attack region, and the rate of change in the axial force is much smaller compared with previous studies in which the jet is ejected from a blunt configuration. The surface flow pattern is also changed by the jet ejection. However, the normal force and pitching moment do not change. Therefore, the influence of the jet strongly depends on the vehicle shape, and the aerodynamic characteristics are restrictive in slender-body shape rockets.

Many satellites are currently in orbit around the Earth and will remain so for years, even though they have completed their intended missions owing to their long natural de-orbiting times. Many methods have been proposed to increase the de-orbiting speed of such satellites, of which one uses atmospheric drag sails. The micro-satellite ALE-1 is equipped with a drag sail that is deployed at a distance from the satellite through the use of a boom element to ensure continuous communication and solar charging. The drag sail is also capable of separating itself from the micro-satellite if necessary to decrease the de-orbiting speed. This paper discusses an in-orbit demonstration where the aforementioned boom element is extended, as well as the deployment of the aforementioned drag sail and com-pares the space and ground experiment results.