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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.