The purpose of this study is to validate the utility of a forward-swept wing in significant reducing the acoustic level of the sonic boom during a supersonic cruise. This was achieved by investigating the aerodynamic performance of wings with different leading edge angles. A total of 13 wing-body-engine nacelle configurations were defined with varying outer wing planforms for a supersonic business jet model. The resulting aerodynamic drag was evaluated using the three-dimensional compressible Euler equation, and the acoustic level of the sonic boom was evaluated using the augmented Burgers equation. The calculated results for supersonic cruise conditions indicated that the maximum acoustic level of the sonic boom of a forward-swept wing was approximately 4.8 PLdB lower than that of the original backward-swept wing. Furthermore, a forward-swept wing reduced the aerodynamic drag as effectively as a backward-swept wing during supersonic cruising. The sonic boom signature at the ground level indicated that forward-sweeping of the wing caused obvious multi-peaks on the bottom peak of the trailing sonic boom, which demonstrated that the forward-swept wing could reduce the intensity of trailing sonic boom.

Impulsive vibration during rocket launch has the risk of causing indentations inside the bearings of reaction wheel mounted on a satellite. The indentations induce the vibration disturbance during the bearing rotation, which is harmful to precision equipment. In order to investigate the mechanism of the indentation formation, we carried out the experiments and finite element simulation for simplified reaction wheel. Based on the results of simulation, we estimated the depth of plastic deformation and wear using Palmgren’s formula and PV value, respectively. The indentation depth of simulation shows good agreements with that of experiment. The results also show that the wear between balls and rings is dominant factor of indentation formation, as compared with the plastic deformation. The amount of wear is greatly affected by the location of ball. In the location where the contact direction between ring and ball is parallel to the vibration direction, contact pressure contributes the wear as compared with slip velocity. Inversely, in the location where the contact direction is perpendicular to the vibration direction, the slip velocity contributes the wear and induces the larger wear region.

To inexpensive and quickly launch microsatellites, ultra-small hybrid rockets have been developed. The current goal to us is to reach an altitude of 100 km in 2020. In order to develop a rocket that reaches a higher altitude in the future, it is necessary to develop the oxidizer tank to be lightweight, large-sized, and low-cost. For this reason, we started in house developed oxidizer tanks from fiscal 2017 and recorded the highest altitude of 4.8 km in the same year. This research reported the results of the 2018 oxidizer tank development.

It is well known that proportional navigation (PN) is used as a terminal guidance law when a missile intercepts its target. But PN does not always provide“optimal”guidance, especially during the boost phase of the missile guidance. Modified proportional navigation (MPN) was proposed for an accelerating missile. However, it assumes the axial acceleration of the missile as constant. On the other hand, in a practical scenario, the missile would be accelerated non-linearly and drastically by a solid rocket motor. In our study, the MPN law is revisited and evaluated with the aim of optimization. We propose a new method which improves the MPN. In our method, a missile and a target are considered as point masses. We assume the missile as a short-range air-to-air missile with a solid rocket motor, and its acceleration is a non-linear function of multiple variables. To maintain the missile speed at high levels as much as possible at the end of engagement (near collision), it is critical that kinetic energy be conserved. In order to achieve this objective, we tried to make the missile lift coefficient as small as possible. Under these conditions, we conducted numerical analyses, and successfully realized a more optimal trajectory than that of MPN. On the other hand, we also successfully reduced the area of the inner launch boundary of the missile.

At bipropellant chemical propulsion systems, iron in oxidizer is one of the topics to be noted and controlled. Actually, to avoid the influence of the iron ingredient, several long term mission spacecraft such as interplanetary probes flush the oxidizer piping periodically. The major influence of steel contaminated oxidizer is known to be flow decay. Still, operation data of the HAYABUSA2 spacecraft indicates possibility of other drawbacks. This paper discusses its impact to the reactivity of the hypergolic propellant, especially for the propellant combination of N₂H₄/MON3, which is the case of HAYABUSA2. A series of open cup firing tests show that the reactivity deteriorates when the oxidizer is contaminated with iron.

A concept study was performed on the spin-launch of a small orbital launch vehicle from a balloon. The small rocket is based on an existing sounding rocket SS-520, the smallest orbital rocket in the world. The increase in launch capacity was calculated by considering the velocity increment required to reach a low earth orbit. Three-dimensional trajectory analysis was performed to obtain the allowable elevation angle error to achieve an orbit insertion accuracy of 1 km in apogee altitude. A new attitude control system of the launcher hanging on a balloon, which uses a double gimbal control moment gyroscope (DCMG), was proposed to satisfy the launch safety requirement. The other issues to be addressed, including the combustion and dynamic behaviors of propellant at high altitude, the reusability of the launcher system, were also discussed.

The effect of aerodynamic drag on loss of the velocity of a microsatellite on LEO is investigated and compared with the on-orbit data of two microsatellites, Hodoyoshi-1 of the University of Tokyo and QSAT-EOS of Kyushu University, which were launched in 2014. A new model, which includes the influence of the solar activities on atmospheric density, is introduced in this paper. The results obtained from the model suggest that the effect of solar activity on the atmospheric density has to be considered with the deviation of F10.7 parameter due to the number of sunspots. The model is applied to a microsatellite composed of a cubic main body and a deorbit sail to estimate its altitude change due to aerodynamic drag. The rotation of the satellite is taken into consideration.

The DC power supply system for the DC discharge plasma generation was developed and the performance of the system was measured. This system can change the discharge input power between 27±3 W and 80±8 W by adjusting the slidac adjustor knob. Corresponding to this input power, the temperature was varied between 1500 K and 2200 K from the initial temperature of 300 K. Based on this performance test, this system is expected to be useful for the future parametric experiment for the comprehension of the interaction phenomena between the shock wave and the DC discharge plasma.

This paper proposes a design method for low-thrust transfer trajectories from an Earth parking orbit to L₁ and L₂ halo orbits in the Earth-Moon system using stable manifold and backward feedback controller design. The transfer trajectory is divided into two phases: a trajectory from the Earth parking orbit to a stable manifold, and one to the halo orbit along the stable manifold. The stable manifold is discretized by two parameters to describe the injection point. A backward feedback controller for the thruster steering in the first phase is employed. By designing the controller backward in time, the target orbital elements can be treated as constants of the Earth parking orbit. Then, the feedback gains and the manifold injection point are optimized using Differential Evolution to minimize ΔV. In numerical simulations, transfer trajectories from an Earth parking orbit to L₁ and L₂ halo orbits are designed by the proposed method to demonstrate its validity.

SHSI, Supersonic Hot Streak Ignition, as a new concept of ignition system for supersonic combustion of hydrocarbon fuel in SCRAM jet engine installed on the small size hypersonic vehicle has been proposed. SHSI fuel of slight amount of hydrogen supplied into compressible turbulent boundary layer at forebody of the vehicle fuselage starts exothermic chemical reaction induced by high temperature at viscous sublayer, then during passing through intake and isolator the cold streak of pilot fuel is transformed to the hot streak as heat source for ignition and flame-holding at main combustor. Analytical research have been conducted to examine the potential for practical operation of SHSI. Damkohler number defined by the ratio between characteristic times of chemical reaction and flow is used to assess the possibility of forming of the hot streak, and it is found that the flight condition of Mach number around 5 at the relatively low altitude in the stratosphere is the limit to achieve auto-ignition of the anterior fuel. Also, compound compressible flow analysis based on one-dimensional stream-tube theory is applied to the validation of capability to avoid the fatal risk of engine unstart at the convergent part of intake, and it is confirmed that the biased distribution towards the downstream of heat release by the hot streak is effective to successfully suppress occurrence of the danger compound choking.

Closed-form solutions for anisotropic cylinders with bending-in-plane, shear-stretching, and twisting-bending coupling stiffness under internal pressure were derived by neglecting transverse shear deformation. The solution provides more accurate bending stress on the bending boundary layer compared with solving a more precise equation including transverse shear deformation. Because the closed-form solution can be calculated directly without iteration, it is useful for optimization problems. The solution was applied to the optimization problem of maximizing the margin of safety (MS) of stress on the bending boundary layer by combining it with the Monte-Carlo method gradient descent method. The result shows that an asymmetric layup has a larger MS than a symmetric layup. This means the closed-form solution including the coupling stiffness through the asymmetric layup is useful.

The main drawback of hybrid rocket is low fuel regression rate. One of the solutions is usage of fuel grain with a complicated geometry port made with additive manufacturing technology. It is difficult to evaluate time- and space-averaged regression rate from experimental data in the case of complicated geometry port, compared to the case of circular port. This reason is that the solid fuel surface of complicated geometry port regresses in intricate manner. In this research, the numerical method to evaluate the time- and space-averaged value was hence developed under the assumption that the solid fuel surface regresses with a uniform and constant speed. Combustion experiments were performed and these experimental data were evaluated with this numerical method. As the results, it was suggested that, except for at low oxidizer mass flux region, the time- and space-averaged regression rates of the star fractal port are almost comparable to those of the circular port and that the fuel mass flow rates of the star fractal port are higher than those of the circular port. These implies the advantage of the star fractal port compared to the circular port in terms of thrust generation.

Solar sail technique was demonstrated in the IKAROS mission. However, the membrane of IKAROS has deformed to a shape that was not flat. The whole membrane changes greatly like an umbrella shape or a saddle shape depending on the warping direction of the thin-film devices on the membrane, such as thin-film solar cells, reflectivity control devices and dust counters. Objection of this study is to clarify mechanism of influence on solar radiation pressure (SRP) torque due to warp of thin-film device and its solution method. Therefore, the shape of the overall membrane is clarified by using a simple finite element model and the SRP torque with respect to the shape is calculated, and the mechanism of the overall shape change in warpage and its influence is clarified. As a result, the influence on SRP is related to membrane stiffness and warped direction and it is found that the suitable conditions to minimize SRP torque are as follows: a) the sail is warped in the radial direction, b) the base film of the sail has low stiffness, c) the outermost part of the sail has high stiffness, and d) the length of bridges between petals is short.

We investigated the unsteady aerodynamic characteristics of NACA0012 airfoil undergoing constant-pitch-rate motions at low Reynolds numbers (Re = 2.5 ×10⁴ ∼ 1 ×10⁵). In the wind tunnel testing, the lift, drag, and pitching moment were measured at constant pitch rates. The experimental results show that unsteady aerodynamic characteristics are equivalent to near-steady ones when the nondimensional pitch rate r is smaller than 4.5×10^-4. When Reynolds number equals to 5 ×10⁴ or less, the nonlinearity of C_l near the pitch angle of 0° is a significant feature at low Reynolds numbers, even in unsteady cases. It is also found that, in the low-pitch-angle region, large changes in C_l and C_m occur due to the unsteady effect. These changes occur at r smaller than in the cases of higher Reynolds numbers. The hysteresis in C_l significantly depends on pitch angle and nondimensional pitch rate.

This paper proposes a deterministic approach to decide the feasibility of a merging problem of an unmanned aerial vehicle (UAV) with a cluster of UAVs that has uncertainty in its velocity. We model the UAVs system yielding decentralized control laws as a mixed logical dynamical system whereby the merging problem is formulated as a mixed-integer convex program to calculate the optimal initial velocity for maximizing the allowable uncertainty. By communalizing certain 0—1 variables in the formulated problem, our approach guarantees the robust feasibility of merging within the calculated uncertainty. The effectiveness of the proposed approach is demonstrated through numerical simulations.

A runway for the operation of a fixed-wind UAV cannot be always available in practical missions and there is no handy method for landing in confined space while a catapult can be used for launch. The purpose of this paper is to develop a new vertical landing method using the flat spin. Flat spin is the motion that an aircraft rotates horizontally while it is descending vertically. In order to control its touchdown position to a desired position, it is necessary to generate centripetal force to control UAV's descent path to target direction while descending vertically. This paper shows the method to change the path during vertical descent, as well as the result of flight tests and wind tunnel tests.

Hypersonic aeroheating is one of challenging tasks in the modern Computational Fluid Dynamics (CFD), partly due to difficulties in accurate capturing of shockwaves and thermal boundary-layers, and partly due to lack of the reference data. We have tackled this problem by computationally reproducing the heating created by a hypersonic-flow/cross-flow-jet aerodynamic interaction. In the first half of this work, we surveyed both numerical flux functions and computational grids, and reached the conclusion that both AUSM⁺-up2 and SLAU2 successfully yielded analytical heating values with symmetry profiles over a blunt-body at Mach 8.1 (without the jet), as long as the minimum cell size was set small enough (cell-Reynolds-number being 0.32 or smaller). Then the second half was dedicated to the cross-flow-jet interaction. Using AUSM⁺-up2, we again succeeded in simulating the jet-interaction-induced aeroheating and pressure augmentations and their distributions, as well as complicated three-dimensional flow structures. Furthermore, a new heating mechanism not tied to pressure increase has been identified. These pieces of information are expected to serve as one of guidelines for hypersonic CFD, new insights into hypersonic flow physics, and also additional data for validation purposes.

For the observation of microwave in space over 100 GHz by radio astronomical satellites, this study develops shape-deformable antenna reflectors. When a certain portion of the antenna reflector prototype is deformed, the entire reflector is affected by the partial deformation. The purpose of this study is firstly clarifying the effect of coupling in reflector displacement. Secondly, this study tries the performance evaluation of hysteresis elimination by a piezoelectric strain feedback system under the effect of the displacement coupling. To achieve these objectives, this study quantifies the displacement coupling, the relations between piezoelectric strain and reflector displacement, and the residual reflector displacement error under the feedback control when multiple actuators are driven simultaneously. As a result, this study clarifies that there is a non-negligible coupling effect and the stress history dependence in the reflector surface deformation, causing the limit in the accuracy of the feedback system. Based on these experimental observations, this paper presents three design improvement plans that will possibly solve these problems.

Study Group for the Future Air Traffic Systems in Ministry of Land, Infrastructure, Transport and Tourism, Japan compiled the ``Long-term Vision for the Future Air Traffic Systems'' and named CARATS. The CARATS pointed out that, weather information is not fully utilized for predicting air traffic capacity in current air traffic systems. In the CARATS Roadmap, conversion of weather information into air traffic operational constraints is listed as a measure. For the purpose of discussing what kind of phenomena can be visualized by Himawari-8 satellite data, the meteorological phenomena that lead to air traffic constraints were analyzed. In this paper, as for the meteorological phenomena that lead to aircraft operational constraints, the cases of typhoons, thunderstorms, fog and volcanic plume have examined. Furthermore, for the cases of typhoons and convective clouds, the differences in information obtained by Himawari-8 satellite data and meteorological radar were also investigated by overlapping the flight path data on the meteorological observation data.

In the present study, a novel wall-modeled large-eddy simulation (WMLES) based on lattice Boltzmann method (LBM) is proposed. WMLES is essential for calculating high Reynolds number wall-bounded turbulent flows such as an external flow around an aircraft. However, WMLES based on LBM is still at its early development stage and further development is needed to handle arbitrary geometries on a non-body-fitted Cartesian grid. Image Point (IP) is introduced to the reconstruction-based WMLES which is first proposed by Malaspinas et al. (2014). Wall function is activated at IP and macroscopic flow velocity derived from the wall function is used to reconstruct the distribution functions at the first off-wall cell. Turbulent channel flow is calculated for the verification and validation purposes and the results are compared with DNS database. The proposed method shows fair agreement with the DNS results in terms of the time-averaged streamwise velocity, Reynolds shear stress and streamwise Reynolds normal stress distributions.

An spin control of a spin-stabilized spacecraft using only magnetic torquers is proposed in this paper. It can simultaneously deal with both of the spin axis control and the angular momentum control. The proposed control law is obtained by applying the cross product law to the angular momentum error. The behavior and the stability under the control are studied analytically. As the result, the proposed control converges the axis direction and the rate of the spin to the desired ones under some assumptions. The analytical study is evaluated through numerical simulations, which show the good agreement between them.

A low-cost, lightweight, high-stiffness and low-shock separation mechanism using separation nuts was developed. This simple and easily manufactured mechanism, which consists of separation nuts, a polycarbonate band, nichrome burn wire, and connecting bolts, is non-explosive; thus, the separation shock is extremely small. The mechanism is intended for a non-pyro, non-explosive rocket using hybrid rocket engine, which does not use explosive fuel. In addition, by virtue of it being simple, low-cost, easily manufactured, lightweight, high stiffness and low-shock, the mechanism can also be used for micro- and nano-satellites. The design, development test, and launch test of the mechanism are reported.