JAXA has a plan to send a spacecraft to the Martian moons in the 2020s. The spacecraft needs to transfer to the orbit of one of the Martian moons. In this study, we investigate the maneuver design method and the characteristics of orbit injection that minimizes the velocity increment Δv for the JAXA's Martian exploration mission. In the previous study, a transfer trajectory with three impulsive maneuvers has been demonstrated and it has been shown that setting the initial argument of periapsis to be zero is effective to reduce the Δv for the transfer. In this paper, we proposed the optimal transfer trajectory with four impulsive maneuvers based on the grid searching of maneuver position. First, it is shown that when the declination of v-infinity vector is large relative to the Mars equatorial plane, there exists a region where adopting four impulsive maneuvers reduces the total Δv compared to the three impulsive maneuver transfer. Also, we show that optimal value of the initial argument of periapsis slightly deviates from zero when the declination is large. Finally, we investigate which transfer type has the smaller Δv for each of the possible launch windows of JAXA's Martian exploration mission.

Grid convergence characteristics of the spectral volume (SV) method is explored for hybrid unstructured meshes. In this study, the SV method originally developed for tetrahedral computational cells is extended to use prismatic and hexahedral cells. The extended SV method (hereafter SV⁺) for hybrid unstructured mesh is first applied to solve the linear advection and diffusion problems. The formal spatial accuracy is rigorously kept in both problems. Then, the SV⁺ is applied to solve the turbulent boundary layer flow over a flat plate. It is shown that the SV⁺ can obtain grid convergence even for using tetrahedral cells in the turbulent boundary layer, where the mesh aspect ratio becomes as large as 20,000. Convergence toward steady solutions using prismatic layers with the same number of DOFs becomes about two times faster than that using tetrahedral cells in the turbulent boundary layer. Finally, the SV⁺ is applied to compute the flowfield around the NASA-CRM. It is shown that the present SV⁺ gives mesh converged solutions which quantitatively agree with the corresponding experimental data regardless of mesh type.

Many space vehicles are powered by liquid hydrogen and liquid oxygen. Such fuel are cryogenic fluids, so they are easy to boil and become gas-liquid two phase flow. The LE-5B-3 engine has the capability of the idle mode firing same as the LE-5B-2 engine. Assessment of flow condition at the inlet of fuel turbo pump is important to operate the engine, because the fuel may flow in saturated condition under the idle mode in principle. In a two-phase flow state, void fraction is one of the most important parameters to assess the flow. Although many types of void fraction sensors were proposed, the capacitive technique has advantages to mount on the engine from the viewpoint of size, weight, toughness. In this study, plural circular electrodes capacitive void fraction sensor is developed for LE-5B-3 engines' ground firing test. The sensor was designed based on electric field analysis, and the specification was assessed prior to the ground test. The sensor was used in qualification test, and it was succeeded in achieving stable measurement and it helped to understand the fluid state during the engine operation. The sensor design technique, the assessment results and the ground test results are discussed in this paper.

The chemical propulsion system of Hayabusa-2 consists of 12 bipropellant thrusters whose thrust is 20N. The communication with Hayabusa was lost due to the fuel leakage just after touchdown in 2005. Akatsuki failed to enter orbit around Venus in 2010. The chemical propulsion system of Hayabusa-2 took measure to prevent these accidents. It satisfied the requirements of continuous injection for SCI (Small Carry-on Impactor) operation. The short injection impulse was estimated using flight data. It was changed by the thruster temperature and the frequency of use. The approximation of the long injection impulse was improved using TCM (Trajectory Correction Maneuver) data and VIC (Velocity Increment Cut) test data. This paper presents development and outward operation of the chemical propulsion system of Hayabusa-2.

Shock wave curvature effect due to the interaction between the DC glow discharged field and the shock wave was investigated by the experiment and the numerical analysis. From the experiment, in the case without the discharged filed, the normal shock wave propagation was observed. On the other hand, in the case with the discharge, the shock wave curvature was observed. From the numerical analysis by using the temperature modulated field model which models the discharged filed, the similar results from the experiment can be observed. In the numerical analysis, in the case with the temperature modulated field with the maximum temperature of 1000 K, the flow field behind the shock wave was investigated, at the timing after the shock wave passing through the maximum temperature position of x=0.025 m corresponding to the timing after passing through the electrode center in the experiment (t = 42.3 μsec, the relative time elapsed from the shock wave passing through x=0 m), for the initial shock wave Mach number of 2. The pressure in the region behind the shock wave was 32 % lower compared to the case without that field, and also the y-direction velocity induced by the baroclinic effect was 94 m/sec at the maximum value in the region just behind the shock wave. This y-direction velocity (94 m/sec) caused to incline the shock wave about 15 degrees from the normal shock wave. From this study, it was suggested that the vortex field mainly due to the baroclinic effect plays an important role for the shock wave curvature effect.

This study discusses a pseudo periodic orbit that is a nearly periodic orbit around a small asteroid such as the Martian moon Deimos. By taking differences of orbital elements between the spacecraft and the asteroid as state variables, basic equations of the relative motion are derived based on Gauss' variational equations. From the basic equations, the initial state of the spacecraft to fly on the pseudo periodic orbit and its period are approximately derived. Furthermore, the stability of the pseudo periodic orbit is analytically examined from the state transition matrix of the linearized equation around the nominal orbit. Numerical studies are executed to verify the analytical results of the conditions and stability of the pseudo periodic orbit.

The purpose of this research is to analyze the characteristics of a limit cycle occurring in the spacecraft attitude when CMGs are used as attitude control actuators. A limit cycle represents the oscillations that may arise from a nonlinear element in a nonlinear system. In a spacecraft system equipped with CMGs, the gimbal friction is a nonlinear element that can induce a limit cycle. In this study, the characteristics of the limit cycle are analyzed 1) by modeling the gimbal friction as static and dynamic frictions, and 2) as based on a describing function of the gimbal friction. The validity of the analytical solution was subsequently verified via numerical simulation and hardware experimentation.

A supersonic parachute is one of effective and promising deceleration processes at entry into a planet. However, fundamental researches on its aerodynamics, particularly at its opening phase, have not been fully carried out. In this study, flow fields in/around simple configurations that modeled a supersonic parachute at its opening are discussed using computed fluid dynamics. Results indicate that pressure inside the parachute is exposed to twice (at the maximum) as large as the pitot pressure at Mach 2. In addition, the so-called “breathing,” which is a more characteristic unsteady phenomenon, appears due to the internal pressure fluctuation, which changes the position of the detached shock wave periodically. Finally, at a higher Reynolds number, the boundary-layer inside the parachute became thinner, leading to more complex shock/boundary-layer interactions.

The membrane of a spinning solar sail IKAROS is considered to be deformed toward the Sun. The deformation was kept even under low spin rate. Previous studies suggest that curvature of thin-film solar cells on the membrane increases the out-of-plane stiffness. For a solar power sail OKEANOS, the shape, out-of-plane stiffness, and natural frequency have to be predicted. In this paper, the effect of a curved thin-film device on the natural frequency of a rectangle membrane under uniaxial tension is investigated. Three types of membranes are evaluated: a membrane with a curved thin-film device, a membrane with a flat thin-film device, a plane membrane. Geometric nonlinear finite element analysis and eigenvalue analysis are conducted to investigate the natural frequencies under varying tension. The simulations are verified by vibration experiments. It is found that under low tension, the natural frequency of the membrane with the curved thin-film device is significantly higher than that of the others and that under high tension, the natural frequency of the membrane with the thin-film device is slightly lower than that of the plane membrane. The results show that curved thin-film devices have a significant impact on the out-of-plane stiffness of a membrane under low tension.

Numerical analysis was performed to figure out the characteristics of the supersonic rectangular inlet buzz, which is the oscillation phenomenon of the shock wave. We found the small oscillation occurs before the Big buzz due to the bleed from the throat. The frequency of the small oscillation corresponds well with the acoustic resonance frequency, which depends on the length of the inlet duct. It was also found that the mechanism of the Big buzz is different from the small oscillation and that the frequency should be estimated by considering the internal shock motion time in addition to the Helmholtz resonance frequency. Namely, both of the length and volume of the inlet duct affect the inlet buzz. When the duct volume is sufficiently small, the buzz can be suppressed. In addition, as bleed plenum volume is increased for the inlet that the buzz is occurring, pressure pulsation is suppressed.

In recent years, laminar flamelet model for compressible flows (compressible flamelet model) was proposed for investigating complicated combustion phenomena such as supersonic combustion and combustion oscillation, which require reproduction of pressure propagation. In this research, we reduce the computational cost of compressible flamelet model by making it possible to refer to thermochemical state quantities of multi-component gas and terms that include spatial gradients of chemical species’ mass fractions from the flamelet tables. Then, we also show that by using Artificial Neural Network (ANN) to develop these flamelet tables, calculation of gradient values of chemical species' mass fractions become easier. For evaluation of these methods in actual combustion field, numerical simulations of the hydrogen combustion in a scramjet test engine of the German Aerospace Center (DLR) is conducted.

An experiment of sequential data assimilation in a cantilevered beam model is presented. The sequential data assimilation is based on the Ensemble Kalman filter (EnKF), which can be easily incorporated in nonlinear finite element analysis. A simple system noise setting scheme is presented, which can set a variance of the system noise in accordance with the residual error in the EnKF. In the experiment, a tip load of the beam is considered as a model parameter to be estimated. The experimental results indicate that the model parameter is accurately estimated by using the EnKF even when the tip load is rapidly changed.

Quality is an important parameter on the gas-liquid two phase flow for organizing its heat transfer, pressure loss properties and flow regime. However, a method to measure quality has not been established yet. A new quality measurement method has been developed by using the mixers and capacitive void fraction sensor. The method is based on the experimental results that the slip ratio of the two phase flow homogenized by the mixers correlates closely with quality. A slip ratio model is created using the mass flux and void fraction after the mixer. Helical type and cross type mixers are arranged in series to homogenize the two phase flow in wide range of the flow regime. Air and silicon oil are used as the working fluids in this experiment. Several types of the flow condition are made by changing the mass flow rates in the horizontal, vertical flow passages. As a result, quality can be measured within ±50% of the error at 92% cases of the whole experiments. Quality meter shows less or comparable errors compared with the Smith's formula which is generally used. In addition, quality meter can be used both for the developed and undeveloped flows.

Flight efficiency consisting of flight time and fuel burn for jet airliners arriving at Tokyo International Airport is analyzed in order to provide reference data for the future development of automated arrival management system AMAN. A total performance index that comprehends the flight efficiency of each flight is optimized, and the result is compared with real flights. Surveillance data, numerical weather forecast data, and aircraft performance data are used in the analysis. Minimal time separation constraint at the merge point is imposed on the trajectory optimization to ensure that there is no interference among arrival flights before landing approach. The quantitative difference of the flight efficiency between the optimal and real flights suggests an ideal achievable benefit to be pursued in the future AMAN.

The reentry capsule is one of the key technologies to support the sample return mission from deep space. The optimization of heat shield design is essential to meet to the mass allocation of a reentry capsule, and one of approach to reduce its weight includes a precise prediction of the aerodynamic heating environment at the rear-side heat shield. As an observation of the rear heat shield of “Hayabusa” reentry capsule, overlapped part of single side aluminized polyimide tape remained after the capsule returned to the earth. This fact shows that the actual heat flux during reentry on the rear was much lower than the assumption in the development phase. In this paper, the heat flux on the rear was identified by the heating tests of the polyimide tape on the ablators using the inductively coupled plasma heater. The result indicates that the heat flux on the rear heat shield is less than 2% of the front stagnation point heat flux of “Hayabusa” capsule. The result shows the possibility that the thickness of ablator material for the rear heat shield can be reduced with sufficient margin against the actual reentry environment according to experimentally estimated heat flux.

A spectral volume method for hybrid unstructured mesh developed earlier is built in an aeroelastic code to determine transonic flutter boundary of AGARD 445.6 wing. It is shown that the computed transonic dip of flutter boundary reasonably agrees with the experimental data. In the present study, we focus on to develop a fast numerical method for evaluating the bottom of transonic dip. In order to reduce the number of computational cases to determine the flutter boundary, an interpolation method using the radial basis function is examined. The initial sampling points are sparsely distributed over the entire parameter plane to roughly determine the flutter boundary by interpolating damping ratio of wing oscillation. Then, based on the consideration of load distribution over the wing, some additional sampling points are introduced in the region where the transonic dip likely appears. It is demonstrated that the present approach allows us to evaluate the bottom of transonic dip with less computational efforts for the present problem. Discussions are included to examine the possible efforts of the arrangement of initial sampling points and interpolation method on the shape of flutter boundary near the dip.

The objective of this study is spacecraft pointing control using a Variable-speed Single-gimbal Control Moment Gyro(VSSGCMG). A VSSGCMG is an attitude control device that consists of a variable-speed rotor and one gimbal. A gyroscopic torque is produced by tilting a rotor and a reaction torque is produced by changing a rotational speed. Thus, a VSSGCMG has two degrees of freedom. Pointing control is to align a certain axis of a spacecraft in an arbitrary direction, and it needs two degrees of freedom. Therefore, the pointing control can be achieved by a VSSGCMG. In this paper, we have proposed a pointing control law by using new parameters to express pointing errors. The validity of the proposed control law is verified by numerical simulations and a hardware experiment.

A morphing wing has been studied to improve aerodynamic performance by deforming the geometrical shape seamlessly and continuously. A compliant mechanism has been proposed as the internal structure to achieve the desired deformation of the morphing shape. This study proposes the multi-layered compliant mechanisms as the internal structure of the morphing flap for improving the design flexibility of the deformation shape. Each layer has an independent structural configuration, but is connected at the wing skin to deform as a morphing flap under an independent applied load. The optimum configuration is obtained by the level set-based topology optimization. In order to perform the optimization with reasonable computational effort, the two-dimensional structural model is adopted to each layer, where the deformation shape at the airfoil skin generated by each layer is regularized to coincide with each other. Through numerical examples, the optimum configurations of the multi-layered rib to achieve the desired deformation under the prescribed load condition are illustrated. Then, validity of the two-dimensional structural model in the topology optimization is demonstrated by comparing with the deformation analysis by using the three-dimensional model.

Numerical simulation results for the optimum design are often different from the experimental results because of simulation errors resulting from the idealization and modeling errors. Less accurate simulation will make the optimization lower confidence. On the natural frequency test of displacement magnifying structure (DMS) which we have obtained by the optimum design, the experimental result had a large gap from the simulation results. In this study, the statistical model is constructed to quantify the gap that is considered as uncertainties. Full factorial numerical integration method that is one of the techniques of uncertainty propagation is adopted to construct the statistical model. The statistical model is used to quantify the effect of the uncertainty on the structural performance such as the reliability concerning the natural frequency of DMS. That differs from the deterministic model parameter estimation method such as finite element updating, that is difficult to quantify the effect of the uncertainties. The proposed method is adopted to construct the statistical model of DMS.

This paper addresses the steering law for a control moment gyro(CMG) system with only three units using an inverse kinematics logic. A CMG system, which usually consists of four or more CMG units for redundancy, is required accurate attitude control even when one of the CMG units stops its function. In this paper, a suitable steering law for such the CMG system with only three units is proposed. In order to utilize the full angular momentum workspace of the CMG system, we propose the control method with the inverse kinematics logic, instead of derivation of the inverse of the Jacobian matrix. The proposed method is evaluated with numerical simulations and hardware experiments, and is concluded effective for such the CMG system with one failure.

In this paper, we proposed Inter-satellite laser power transmission service and indicated its feasibility from the viewpoint of both technology and business. Solar-pumped laser allows for power transmission between 24 satellites at an altitude of 300 km and 24 satellites at an altitude of 1800 km assuming that some energy conversion efficiency and beam efficiency. If 24 power transmission satellites could be developed with an initial cost of 7.97 billion yen, an operational one of 10 thousand yen/hour, and a development period of five years, initial cost could be recovered in the 14th year from launch and a benefit of 55 billion yen can be achieved during 15 years which is a lifetime of the power transmission satellites.

In this paper, we evaluate a mobility system using skids for a small lander. The small lander named Smart Lander for Investigating Moon (SLIM) has been developed by ISAS/JAXA. In the small lander mission such as SLIM, planetary surface exploration after the landing will become difficult due to the restriction of the lander's payload weight for a rover and observation equipment. To solve this problem, we propose the mobility system using skid-sliding to improve exploration capability of the small lander mission. We confirmed the skids' sliding potential of the small lander using numerical simulation. The simulation showed that the small lander can travel to the desired direction using skids and thruster control.

A thin-film device mounted on development type membrane structure is limited to arrangement and deployment on the plane. In this paper, we propose a three-dimensional film structure (3D film) that can stereoscopically place and deploy thin-film devices. This method can be widely applied to non-spin type or spin type developable structures having a mast structure. In this paper, in order to apply to non-spin / spin type developed structures, we evaluate extension shapes of 3D film by crease and condition of existence / nonexistence of centrifugal force, and placement and arrangement of thin-film devices.

In this study, we proposed an empirical noise evaluation parameter for a slat in a three-element high-lift airfoil to design a low-noise high-element airfoil with minimal evaluation cost. In our proposal, we focused on the distance from the trailing edge to the reattachment position of a shear layer emanating from the slat cusp, which can be estimated using Reynolds-averaged Navier Stokes (RANS) simulations and represents the size of the circulation region in a slat cove. To investigate this assumption and to evaluate the noise characteristics, Ffowcs-Williams and Hawkings' far-field and delayed-detached eddy simulations were conducted. We found that the lower the noise observed in real scale, the longer was the distance. In addition, we also found that a remarkable relationship exists between this distance and broad-band noise. We then attempted a multi-objective optimization using a Kriging model based on an evolutionary algorithm. The objective was to maximize the distance to reduce the broadband noise and maximum lift. Based on the optimization, 27 cases in which the slat position was changed were evaluated through RANS simulations to construct the Kriging model. The optimization results reconfirmed that slat overlap has the largest effect on the distance from the trailing edge to the reattachment position. These results suggest that the optimum slat position to achieve noise reduction can be achieved without reducing maximum lift.

This paper proposes an optimal design method of a trajectory of spacecraft using multi-flyby. In recent years, many deep space explorations have been conducted, and it is necessary to minimize the fuel consumption for these missions. Therefore, a trajectory of spacecraft that minimizes the fuel consumption with the mission constraints satisfied is determined by numerical optimization. In order to cope with complicated orbits with many flybys and severe constraints, this paper extends the multiple gravity assist (MGA) problem, and proposes a new optimal trajectory design algorithm. We newly add two discrete optimization variables to the MGA problem: the number of revolutions before flyby, and selection of multiple solutions to the Lambert problem. In the proposed algorithm, trajectory optimization is recursively executed by adding a flyby target one by one, instead of optimizing the entire trajectory with multi-flybys at once. The validity of the proposed algorithm is demonstrated by numerical experiments with two scenarios of Parker Solar Probe and Solar Orbiter. In addition, the relation between the number of flybys and fuel consumption is also examined, which shows the trade-off relationship between the two.