This paper addresses a 3-impulse Lunar transfer orbit design method using a real-coded genetic algorithm (real-coded GA). The real-coded GA is applied to the transfer orbit design taking into account the constraint of argument of perigee induced by a launch vehicle. SLIM is a Lunar lander which is developed in JAXA. SLIM is planned to be launched by the solid motor launch vehicle, Epsilon. Since the third motor of Epsilon employs spin stability, the argument of perigee of the injected orbit is constrained to maximize the launch capability, which makes it difficult to reach Moon directly from the injected orbit. Considering such unique characteristics by the launch vehicle, a mid-course maneuver is applied after the trans-Lunar injection to connect to the lunarcentric orbit. The proposed orbit design method using the real-coded GA generates a suboptimal orbit with sufficiently small computational burden, and can be widely applied to an orbit design including mid-course orbit control maneuvers.

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Current control of arrival flights at Tokyo International Airport is studied using real traffic data in order to predict performances which could be realized by a future arrival management system. The current control procedure employs radar vectoring which uses relatively wide airspace including en-route cruise phase before the descent. A stochastic model presenting the queueing mechanism of the procedure is constructed. Parameters in the model are derived from real traffic data. Various cases of different parameters are analyzed by Monte Carlo simulation in order to examine influences of various conditions upon flight time delay. It explains that the randomness of the entry time as well as the controlled time of separation inevitably derives some flight time delay even if the traffic volume is light, and the real system selects the traffic volume in order to make the delay at an acceptable level. It indicates that the traffic volume could be increased by 20 % if the error of controlled time for separation is reduced to the half.

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Laminar flow control is next-generation technology which is expected to improve aerodynamic performance greatly. As for an application of laminar flow control in the aircraft design, it is important to consider comprehensive effects of laminar flow control system on design results. In this paper, we include the effects caused by additional laminar flow control system into the aircraft conceptual design tool and consider interactive effects on aerodynamic, propulsion and secondary power systems. Two types of laminar flow control technology have been considered and were applied to 200 passenger class transport aircraft. They are natural laminar flow (NLF) and hybrid laminar flow control (HLFC). Results indicated that NLF can improve lift / drag ratio and fuel consumption, and hence decreases the aircraft empty weight. Applying the HLFC improves lift / drag ratio and fuel consumption further. However, because of the additional weight of HLFC system, the maximum take-off weight of HLFC aircraft is equivalent to that of the NLF.

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IKAROS is a solar power sail demonstrator launched by JAXA in 2010. IKAROS successfully deployed its large sail by a centrifugal force due to the spin motion of the spacecraft body and obtained a solar power sail navigation. In the evaluation of IKAROS membrane shape, some unexpected phenomena were observed; (1) tether connecting IKAROS membrane and main body got loose in spite of normal spin rate, and (2) the membrane was not warped by photon pressure in spite of low spin rate. The purpose of this paper is to understand generation mechanisms about these unexpected phenomena. Because thin-film devices of thin-film solar cell and reflectivity control device are multilayer film structures, curves occur. The bending stiffness of IKAROS membrane is increased due to the curve. This paper presents the multi-particle model of the membrane considering curve and bending moment of thin-film device. The membrane shape is estimated by numerical simulations and the influence of the thin-film device is made clear.

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SLIM (Smart Lander for Investigating Moon) is the Lunar Landing Demonstrator which is under development at ISAS/JAXA. SLIM demonstrates not only so-called Pin-Point Landing Technique to the lunar surface, but also demonstrates the design to make the explorer small and lightweight. Realizing the compact explorer is one of the key points to achieve the frequent lunar and planetary explorations. This paper summarizes the preliminary system design of SLIM, especially the way to reduce the size.

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Thanks to recent lunar exploration missions, high-resolution lunar surface observation data was obtained. In future lunar exploration, landing is being requested at a specific point having higher scientific interest than other areas. The SLIM project is demonstrating pinpoint landing technology, which entails a combination of “autonomous image-based high-precision navigation technology” and “autonomous guidance technology intended to generate a fuel-optimum landing trajectory.” This paper presents powered descending trajectory design in terms of trajectory optimization. As usually considered in general space mission development, an optimal solution in terms of minimum fuel consumption is the basis of investigation. This study addresses trajectory optimization considering specific objective functions derived from practical constraints regarding mission design, such as altitude, downrange length, and visibility from ground stations. In this paper, nominal trajectory design considering minimum fuel consumption is first presented, followed by parametric studies to identify the sensitivity to changes in initial conditions under which powered descending starts. Finally, trajectory optimization results with various types of objective functions are presented.

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This paper is on optimal trajectory of future lunar lander with coasting in powered descending phase. For the light weight/low cost lunar lander, optical navigation using onboard cameras to identify their current state is one of very few techniques available to achieve the pin-point landing. The optical navigation is to be operated between the powered descending phases, when the orbital maneuvering engine (OME) is turned off. This paper shows the series of different coasting conditions and discusses the effect of the coasting to the trajectory and fuel consumption. The results give some ideas for future gravitational planetary missions, which uses coasting during their powered descending phase. In addition, optimal trajectory with double coasting for the SLIM project is shown in this paper.

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SLIM project which aim for pinpoint landing on the moon surface. For achieving this plan, it is necessary to estimate the flight position of the space probe. The estimation is performed by matching the detected craters with database. This paper introduces a crater detection method using Principal Component Analysis (PCA) and its evaluation. This method is capable of real-time processing under low computational resources such as Field-Programmable Gate Array (FPGA). In this research, we report improvement of robustness at detection and high accuracy of crater size measurement.

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This paper focuses on the Evolutional Triangle Similarity Matching (ETSM) method for estimating spacecraft location in Smart Lander for Investigating Moon (SLIM) mission and improves it by adding the functions of elimination of line symmetric triangles between crater map and camera shot image, comparison of rotation relationship of triangles and triangle formation method using Delaunay triangulation and introducing point group matching as a coordinate calculation function. To evaluate the robustness of the improved method, we conduct simulation experiments using the crater map and camera shot images in six situations. This experiments have revealed the following implications: (1) this method improved accuracy of location estimation within 5.1 pixels by the functions of elimination of line symmetric triangles between crater map and camera shot image, (2) this method slight got worse accuracy at low or high altitude of spacecraft, however, this method successfully reduced incorrect spacecraft location estimation by comparison of rotation relationship of triangles, (3) this method improved accuracy of location estimation by triangle formation method using Delaunay triangulation, but possibility of incorrect spacecraft location estimation is slight increased, and (4) integration method of these three mechanism can estimate spacecraft location within 5 pixels without being affected altitude difference and rotation of camera shot image.

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Next generation moon landing mission will require autonomous pinpoint landing capability because of requirements for landing on specific terrains in a limited area. This capability requires precise absolute self-localization of the lander during braking descent phase. The purpose of this paper is to propose an algorithm to estimate the lander position and to evaluate its mountability to a space-grade FPGA. In this method, the position estimation is performed by matching crater point patterns with database point patterns by finding topological correspondences using crater-based linear features. In addition, we confirmed the resource amount and the calculation time when this algorithm is implemented on the FPGA using high-level synthesis.

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The landing radar employs a pulse-type radar using 4.3 GHz C-band microwave radiation. It has a wide beam for measuring the altitude in vertical direction, as well as four narrow tilted beams for measuring the velocity in horizontal direction. In this paper, development of the Bread Board Model (BBM), a field experiment, and the design of SLIM loading Flight Model (FM) are introduced. Furthermore, the radar simulator required for FM development of a radar is explained.

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Ceramic/metal brazing was investigated to produce light-weight and highly-efficient ceramic thrusters. Silicon nitride ceramic and metal bars were brazed using an Ag-based brazing material. Four-point bend tests were conducted at room and high temperatures to evaluate the strength of the brazed joints. Computational fluid dynamics (CFD) and finite element method (FEM) analyses were also performed to investigate the effect of the construction and shape of the joints on the stress distribution around them. It was demonstrated that brazing was a great candidate as the joining technique, and a 20 N ceramics/metal brazed thruster was successfully produced.

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In this paper, the authors propose a novel landing method named “Two-step Landing Method” for small lunar lander which is needed to be designed considering constrains from envelope area of rocket and the weight of the lander. The proposed method enforces intentional body tumbling at the contact of main leg. We analyzed its dynamics by three-dimensional simulations which consider lander’s attitude and lateral velocity and landing site’s slope angle. Numerical simulation models have been designed on Mechanical Dynamics Software “ADAMS”, and lander models refer to “SLIM” which is a small lander proposed by ISAS/JAXA. It is found that the proposed landing method can land on steep slope by tilting body attitude toward inclination direction of landing site. Especially in the case of landing with lateral residual velocity, the proposed method has higher landing stability than conventional landing method.

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Energy absorbing system for landing gears is an important on the SLIM project. Open cell porous aluminum manufactured through 3D selective laser melting (SLM) process has been applied on the energy absorbing system. Compressive tests for cylindrical and hemispherical shaped porous aluminum with different porosities revealed the high potential as an energy absorbing component. Heat treatment after SLM processing is effective to increase the energy absorbing potential of the porous aluminum.

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Summary of First Aerodynamics Prediction Challenge (APC-I) is presented. The APC-I is a domestic CFD prediction workshop that was held on July 3, 2015. The test cases include aerodynamic prediction of NASA-CRM with and without aeroelastic effects, and its wake flow prediction. We compare the CFD results with JAXA's wind tunnel measurements. There are 15 participants from government, academia, industry, and commercial. The CFD results submitted from the participants are compared and discussed.

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Summary of Second Aerodynamics Prediction Challenge (APC-II) is presented. The APC-II is a domestic CFD prediction workshop that was held on July 6, 2016. The test cases include aerodynamic prediction of NASA-CRM with and without support effects, and buffet prediction. We compare the CFD results with JAXA's wind tunnel measurements. There are 9 participants from national research agency, academia, industry, and commercial software vendor. The CFD results submitted from the participants are compared and discussed in the presentation.

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The computational grid dependency is an important problem for CFD. We have computed aerodynamics on NASA-Common Research Model (CRM) with FaSTAR and various grids to investigate the grid dependency. We employed four grids: two Cartesian-based unstructured grids, a tetrahedral unstructured grid, and a hexahedral structured grid. The computational conditions are based on the test cases of Aerodynamics Prediction Challenge (APC). First, the grid convergence at a fixed angle of attack and the trend of an angle-of-attack sweep are compared between the four grids. The lift coefficients computed with the two similar Cartesian-based grids are different, and this is caused by the grid difference around the leading edge. However, the overall trend of angle-of-attack sweep is almost same between the four grids. Next, we computed aerodynamics on NASA-CRM with a support device to investigate the support interference. It is found that the support interference on the drag and pitching moment is large and should be considered.

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In response to the First and Second Aerodynamics Prediction Challenges, held in Tokyo, July 2015 and in Kanazawa, July 2016, respectively, computational fluid dynamics simulations were performed for the NASA Common Research Model using the Tohoku University Aerodynamic Simulation (TAS) Code. Our results were summarized in this manuscript, with an emphasis on key computational techniques and mesh generation methods. Unstructured hybrid meshes were generated using the Mixed-Element Grid Generator in 3 Dimensions (MEGG3D), and were deformed based on wing deformation data obtained during wind tunnel testing. The effects of support system interference, of mesh density and of laminar to turbulent boundary layer transition are shown to discuss the validity of computational results. Aerodynamic coefficients were well predicted at low angles of attack when the support system interference effect was considered, while an accurate prediction of pitching moment at high angles of attack was challenging because mesh density affected the shock location on the wing and the size of side-of-body separation.

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In order to validate CFD results using wind tunnel test (WTT) data, it is necessary to consider what types of corrections are applied to the WTT data, such as wall interference correction, near-field support interference correction and so on. In this paper, a CFD tool called “Cflow” developed by Kawasaki Heavy Industries, Ltd. is validated using results of wind tunnel test conducted by JAXA with NASA-CRM (Common Research Model). At first, effects of near-field support interference and wing deformation on aerodynamic performance are discussed. Then it is confirmed that Cflow results are well consistent with experimental results, taking these two effects into account.

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Inviscid and Reynolds-averaged Navier-Stokes (RANS) simulations of transonic flows around the NASA Common Research Model are conducted using the Cartesian flow solver UTCart. The immersed boundary method is used to represent the smooth geometry surfaces on the Cartesian grids. The wall function is combined with the immersed boundary method to reproduce the turbulent boundary layer on the geometry surface in the RANS simulations. In the inviscid calculations, the qualitative flow feature including the position on the shock-wave on the wing shows agreement with the reference result a body-fitted grid. In the RANS calculations, the trend of pitching moment and drag shows fair agreement with the reference result, while prediction of the flow separation at high angle of attack is still difficult. Compared with the reference result, the differences in the total drag coefficient at a moderate angle of attack on the medium grid (33 million cells) and the fine grid (99 million cells) are 31 drag counts (10%) and 20 drag counts (6.5%), respectively. Furthermore, each of the calculated aerodynamic coefficients shows a consistent trend of grid convergence toward the reference result.

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Aerodynamics coefficients of NASA common research model (NASA-CRM) are computationally investigated by revisiting some test cases appeared in aerodynamics prediction challenge (APC) using high-order discontinuous Galerkin (DG) methods. While the employed high-order DG methods reasonably well predict the overall aerodynamics for the NASA-CRM, some discrepancies appear between the CFD and experimental data, especially in a lift and a pitching moment at low angle of attack. Although those discrepancies still exist, installing a sting enhances the reproduction of experimental data of pitching moment. Additionally, fluid-structure coupled simulations used to consider aero-elastic deformations give better agreements of the pitching moment with experimental data.

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The conventional spectral volume (SV) method for three-dimensional tetrahedral unstructured mesh is extended to use hexahedral mesh. In the test calculations of linear scalar advection problem and diffusion problem, the formal spatial order of accuracy is achieved even for skewed computational meshes. The Spalart Allmaras turbulence model implemented in the present code is verified by calculating the grid-converged skin friction of turbulent boundary layer on a flat plate. Then verified code is applied to compute the flowfield around the NASA-CRM. In this study, we examine the agreement of the computed aerodynamic coefficients with the corresponding experimental data for different angles of attack. We also examine the change of aerodynamic coefficients with varying Reynolds number, although the experimental data is not available. It is shown that the present SV code can predict aerodynamic coefficients around the cruise angle of attack conditions fairly well. When higher Reynolds number is assumed, the computed lift increases while the viscous drag is reduced, as is expected. On the other hand, the pressure drag is increased with increasing Reynolds number due to the shock wave on the wing which moves toward downstream.

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In this study, the aerodynamic performance of NASA Common Research Model (wing-body configuration) was analyzed by coupling a Cartesian mesh CFD solver of Building-Cube Method (BCM) and an unstructured mesh CFD solver of Tohoku University Aerodynamic Simulation (TAS) codes. The thin boundary layer was handled by the unstructured body-fitted mesh near wall, while the vortical wake was effectively resolved by the multi-level Cartesian mesh of BCM. The computational results were compared with those of the transonic wind tunnel tests for validation. The lift and drag coefficients as well as pressure coefficient around wing sections were comparable with the experimental and numerical results by other participants. In addition, the advantage of the multi-level Cartesian mesh was presented by the sharply captured wake in the present simulation. It is confirmed that results of the BCM-TAS Coupling Flow Solver is generally agreed well with the experimental data in aerodynamic predictions and wake analyses.

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For the aerodynamic design of aircraft, CFD has made a remarkable progress to provide lift, drag and surface pressure distribution collaborating with wing tunnel experiment. However, not much information is known about the flow phenomena in wakes of airplanes and wings. In the article, we concentrate on wake flow analysis as a post work of computations and experiments done for APC-I workshops. Through the analysis, we have found that wing tip vortices and boundary layers of a wing affect velocity distribution on a wake, while only the tip vortices do pressure. The precise prediction of wake physics deeply depends on grid resolution quality in subspaces such as cross-sectional planes in a wake away from an airplane surface, for both of experiment and computation.

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The paper reports numerical simulations of aerodynamic characteristics of a commercial aircraft model known as the NASA-CRM. Specific test cases and detailed experimental results have been provided by JAXA for the CFD workshop “Aerodynamic Prediction Challenge (APC).” We employed the cell-vertex finite-volume method on unstructured grids and carried out the RANS simulations with the k-ω SST turbulence model to present results in the APC-I. After the verification of the model in a benchmark problem, both alpha-sweep simulations from low- to high-angles of attack and the grid-convergence study at a cruise conditions are carried out. The simulations well reproduced the experimental results at moderate angles of attack by both types of computational grids used in the study but the results showed the dependency on the grid type and grid resolutions at the highest angle of attack of 5.72 degrees due to different predictions of flow separation near the root chord of the main wing.

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The moving surface method (Wang Y., et al., J. Jpn. Soc. Aero. Space Sci., Vol. 58, 2010, pp. 239-244, in Japanese) is applied to low Reynolds number flows over thin airfoils as a flow separation control method for further aerodynamic enhancement. The moving surface method is based on Couette Flow-type momentum addition and is expected as a promising means in order to achieve lift enhancement. In order to evaluate aerodynamic effects of the moving surface method at a low Reynolds number, we employed NACA0006, Ishii airfoil, and Owl-like airfoil, and performed two dimensional CFD simulations. The results demonstrate that by applying the moving surface method, the leeward flow separation is successfully suppressed and lift-to-drag ratio is increased, particularly for the Owl-like airfoil. In addition, interestingly, a further drag reduction has been observed by increasing the surface-moving velocity, in which pressure suction at the leading edge is promoted. As a consequence, the maximum lift-to-drag ratio reached 44 (Owl-like airfoil with 9 degrees of attack angle), 2.2 times the case without the moving surface.

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Takano Laboratory of Kanagawa University is working on the development and production of an ultra-small rocket for inexpensive and rapid launch of ultra-small satellites using a hybrid rocket engine. In the past, Canadian HyperTEK Hybrid Engine was used. However, because this has a limitation on engine performance, Takano Laboratory started development of its own hybrid engine from FY 2015. The goal is to develop a new engine that is more lightweight and equivalent in performance to the HyperTEK L type. Weight reduction was succeeded by using carbon fiber composite material and flight experiment was also performed in FY 2016. In this paper, the results of a series of engine developments are reported.

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The assessment of the simple temperature estimation theory based on the shock wave propagation velocity was conducted based on the comparison with the CFD simulation of the shock wave propagation through the temperature modulated field at the region from y=0 m to y=0.018m. The temperature estimated by the theory using the shock wave propagation velocity calculated from the CFD was compared with the initial temperature distribution given in CFD. This temperature estimation theory can predict the temperature distribution within at most 30 % error, for the present temperature field. The assessment suggests that the estimation error due mainly to the two dimensional effect should be paid attention for applying the theory.

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Four types of flexible parachute in different shape (Hemisphere, Disk Gap Band (DGB), Conical Gap Band (CGB), 10Slit) are tested in the JAXA-ISAS supersonic wind tunnel. Oscillation characteristics of the swinging motion and the canopy deformation are investigated. As a result, the canopy deformation, which was evaluated by using the shape-complexity, could be characterized by the Gamma-type probability-distribution-functions (pdfs), while the amplitude of swinging oscillation of the canopy can be characterized by the Gaussian-type pdfs. It was found that the CGB-shape features the largest canopy deformation and the lowest oscillation amplitude among tested shapes.

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Scaling characteristics of an electric multi-copter powered by internal combustion engine are investigated. A 1-kW-class internal combustion electric generator is developed, and the energy efficiency of the internal combustion engine is investigated experimentally, and the characteristics of brushless motor are modeled theoretically. The relation between the flight time and the Multi-copter size is investigated on the basis of the statistical modeling of the existing data of the commercial engines and the motors. As a result, the flight time is found maximized at the multi-copter weight of 8 kg.

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In future space transportation technology, reentry-systems are required, which is useful, safe and low-cost. MAAC (Membrane Aeroshell for Atmospheric-entry Capsule) R&D group have been developed the flexible structure atmosphere reentry system. The paper investigates the buckling strength of the flexible membrane reentry structure under static aerodynamic pressures in terms of non-linear finite element method. Numerical results are compared with experimental data obtained from a series of low speed wind tunnel tests using scale models of the proposed membrane aeroshell system. The comparison shows that the numerical results qualitatively agree with the experimental data, although the results clearly overestimate the buckling load of the membrane aeroshell. We conjecture that the discrepancies are originated from geometrical imperfections of the scale models not accounted for in the numerical analysis and introduce plausible patterns of the imperfections into our finite element models. Numerical analysis for the modified finite element models more precisely predicts the buckling loads, indicating that the geometrical imperfections significantly affect the structural strength of the membrane aeroshell.

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Sand erosion is a phenomenon wherein solid sand particles impact on materials causing damage to the material surface. Aircraft jet engine is exposed to many particles from takeoff to landing, because sucking in a considerable amount of air and sand. Aircraft jet engine operating under such an environment suffer from degraded performance and have a shorter lifespan. Therefore, it is necessary to estimate the amount of damage caused due to the shape and material of the aircraft engine airfoil at the design stage. In this study, a sand erosion prediction method for evaluating the lifespan of engine structural material was proposed. First, sand erosion experiments in which the erosion rate was calculated by varying the impact velocity and impact angle for the A7075-T7351 and Ti-6Al-4V plates were conducted. Second, the erosion rate on the surface was predicted by applying the erosion rate equation to a two-dimensional airfoil. Subsequently, erosion rates of the calculated A7075-T7351 and Ti-6Al-4V two-dimensional airfoil were examined and compared. Finally, the validity of the prediction was confirmed by a verification experiment.

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A new Continuous Wave (CW) laser propulsion system using porous carbon as a heat exchanger is experimentally studied. The propellant receives the heat from heated porous carbon absorbing CW laser and becomes high pressure in the chamber of the thruster. The thruster obtains propulsive force by emitting hot and high-pressure gas through its nozzle. In this experiment, we used the 4 kW CW fiber laser, helium gas as propellant gas. We found that pressure in the chamber and thrust increased by laser heating and it was possible to raise the stagnation temperature to 1523 K.

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Foreign object damage (FOD), e.g. a bird strike, is seriously damaging event during flight operations of aviation. For simulating bird strike, soft-body impact test is investigated and gelatin is used as the projectile material traditionally. However, the effect of projectile shape on the soft-body impact test results has not been sufficiently established. In this study, a cantilever target made from aluminum alloy and carbon fiber/epoxy composite were used, and experiments and FEM simulations were conducted to investigate the effect of the projectile shape. The projectile shape should affect impulse and/or work immediately after impact, and it affects the plastic deformation around the impact point of aluminum alloy specimen. And the projectile shape should affect the interlaminar shear stress around the impact point and edge of specimen, and affects the delamination of carbon/epoxy composites specimen.

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For satellite, it is necessary to reduce the vibration of rotating bearings in reaction wheel. In order to investigate the mechanism of the small vibration caused by bearings and its control scheme, we have conducted the multi-body dynamics analysis for angular ball bearing and reaction wheel. We focus on the effect of mechanical factors such as non-circularities of the inner and outer rings and the angle misalignment between them to the rotating secondary vibration component in the axial direction. It was found that the angle-misalignment enhances the rotating secondary vibration component when the inner ring involves second-order waviness. Also, axial-direction vibration is enhanced in the case that the waviness orders of inner and outer rings are coincident each other and angular-direction vibration is enhanced in the case that the difference in waviness orders of inner and outer rings is equal to 1 when the inner and outer rings involve the waviness. These vibrations are caused by the non-linear elastic contact between ball and raceway. Finally, from the parameter design, it was found that the effect of angular misalignment to axial vibration is greater than that of non-circularity of outer rings and these mechanical factors independently affect to the vibration.

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