Fuel efficient flight is demanded due to air traffic increase, environmental concerns and skyrocketing fuel prices. To achieve an effective futuristic air traffic management system, realization of TBO (Trajectory Based Operations) optimized from departure to arrival is aimed in CARATS (Collaborative Actions for Renovation of Air Traffic Systems). In this paper, 4-D trajectory of passenger aircraft is optimized for fuel minimization on the assumption that aircraft is not bounded by any air traffic control constraints. BADA (Base of Aircraft Data) model is adopted for aircraft performance model, and dynamic programming is used to calculate the optimal trajectory. The optimal trajectory is compared with airliner’s flight trajectory by evaluating fuel consumption, flight time and flight range. Examples show that considerable amounts of reduction in fuel consumption could be achieved by trajectory optimization compared to the estimated values of actual flights.

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We aim to investigate radius of water drop (rocket clouds) and mass fraction of the liquid phase in transonic wet air flow around a rocket fairing, and the aerodynamic characteristics of the fairing. In this study, a two-dimensional or axisymmetrical two-dimensional CFD code considering the heterogenous condensation (wet air flow) is developed. The numerical simulation is validated by a solving flow in channel with circular-arc bump. The dry and the wet air flow around the rocket fairing (simulated 4S-type for H-IIA, Length of 12 m, M_inf = 0.85, T_inf = 300 K, P_inf = 1000 hPa) are solved by the numerical simulation. In the case of the wet air flow around the rocket fairing, the Mach number and the mass fraction of the liquid phase are 1.45 and 1.08% respectively near the shoulder of the fairing z = 66.1 m (the fairing tip is located at z = 60 m). The maximum radius of a water drop is 0.27 μm. Water drops, which their radius is over 0.11 μm, are generated near the shoulder of the fairing. The water drops are likely of being observed as the rocket clouds like umbrella. Compared to the case of the dry air flow, the shock position on the rocket fairing wall is shifted to downstream in the case of the wet air flow due to the condensing heat in the wet flow. The difference of the static pressure between the dry and wet air flow is observed on the side wall of the fairing from z = 66.1 m to 68.5 m. In particular, the static pressure of the wet air flow is 7.20 kPa higher than that of the dry air flow near the shoulder of the fairing.

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In this study, the effect of Non-Condensable Gas (NCG) in Oscillating Heat Pipe (OHP) was evaluated by using Closed Loop OHP (CLOHP) made of copper capillary tube. The working fluid was ethanol and the NCG was argon. Temperature and pressure were measured over the heat input range from 40 W to 180 W. The temperature amplitude, the thermal resistance, the pressure amplitude and frequency were calculated. From the results, the NCG affects the thermal performance of OHP.

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In this paper, novel guidance method in which the path planning and following are integrated is proposed.The path planner provides the optimal trajectory minimizing the integral square curvature and the optimal feedback input.The input is used for path following with the PD controller for the cross-track error.The simple kinematic model which is so called unicycle is assumed.Simulations where the steady wind disturbance and initial position errors are considered is conducted to demonstrate adaptive capability of the proposed method.

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The purpose of this paper is creating a method to determine the locations of components applying Genetic Programming (GP) – one of evolutionary computation algorithm which is often applied to solve symbolic regression problems. In the aircraft conceptual design, first, mission requirements are given. Secondly, the parameters of main components such as wing, tail, fuselage and engine are determined to meet these requirements. Finally, the locations of the components are fixed considering the center of gravity, stability and gear layout. Then an initial three-dimensional view is drawn. The wings and the gear locations are mainly determined by the designers' sense based on their experience as there is no well-defined method during the preliminary phase. Therefore, GP is applied to the problem relating to the aircraft conceptual design and the functions between the locations of components and constraints such as stability and gear layout are obtained. These results indicated potential abilities of GP to create a model which would help the aircraft designer to determine the locations of the components during the design process.

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A single-stage launch vehicle with hybrid rocket engine, which uses solid fuel and liquid oxidizer, has been being studied and developed as a next-generation rocket for scientific observation due to the advantages as low cost, safety, re-ignition, and reduced pollution. Therefore, the knowledge regarding hybrid rocket system has been being gained through the forepart of the conceptual design using design informatics. In the present study, the practical problem defined by using three objective functions and seven design variables for aurora observation is treated so as to contribute the real world using evolutionary computation and data mining for the field of aerospace engineering. The primary objective of the design in the present study is that the down range and the duration time in the lower thermosphere are sufficiently obtained for the aurora scientific observation, whereas the initial gross weight is held down. Investigated solid fuels are five, while liquid oxidizer is considered as liquid oxygen. The condition of single-time ignition is assumed in flight sequence in order to quantitatively investigate the ascendancy of multi-time ignition. A hybrid evolutionary computation between the differential evolution and the genetic algorithm is employed for the multidisciplinary design optimization. A self-organizing map is used for the data mining technique in order to extract global design information. Consequently, the design information regarding the tradeoffs among the objective functions, the behaviors of the design variables in the design space to become the nondominated solutions, and the implication of the design variables for the objective functions have been obtained in order to quantitatively differentiate the advantage of hybrid rocket engine in view of the five fuels. Moreover, the next assignments were also revealed.

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This paper proposes to apply probability density ratio estimation technique to stochastic robust control design. The approach has the potential to reduce computational load in stochastic optimization. It is applied to a simple example problem of flight control and the result is promising.

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The transition characteristics on candidate shapes were examined numerically and experimentally in order to design a natural laminar flow fuselage nose for supersonic transport. The candidate shapes were designed through axisymmetric sinusoidal deformations added on base shapes. Not only parameters of geometry which define the deformation, but also the conditions of uniform flow were varied to obtain the greatest effect of natural laminar flow. From the results, the natural laminar flow effect was found to be obtained in cases of the small ratio of pressure gradients in azimuthal and axial directions in the side area. The natural laminar flow effect was validated experimentally.

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To establish a precise surface shape measurement technique, we improve our three dimensional measurement system using grating projection method. Improved points are 1) using a higher resolution camera, 2) setting supporting jigs with higher stiffness and higher accuracy, and 3) adjusting measurement parameters. Results of measurement experiments confirm effectiveness of our improvements. On the other hand, measurement results using optical method, such as grating projection method, are affected by optical characteristics of measured objects in general. Then, we apply the method to a surface shape measurement of a honeycomb sandwich panel which is widely used in space engineering. In this paper, to verify our measurement method we compare the measurement result with that using laser measurement equipment, and a good agreement is confirmed.

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A solar power sail demonstrator “IKAROS” deployed its square large membrane successfully by centrifugal force. The deployment consisted of two stages. The flight data suggest that the membrane deployment in the second stage was remarkably non-synchronous. This paper reproduces the non-synchronous deployment behavior using a small experimental model deployed in a vacuum chamber. Then, two kinds of finite-element analyses are conducted to show that the effect of membrane creases may have caused the non-synchronous deployment.

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Simple, closed-form solutions, which can be easily calculated using spreadsheet software, for thin and moderately thick anisotropic cylinders under torsion, axial compression and these combined loads that includes the effect of transverse shear deformation are derived. Previously derived partial differential equations of equilibrium, which are including both layup anisotropy and transverse shear deformation, are separated two equation groups by neglecting coupling stiffnesses (tension-shear, bending-torsion and tension-torsion couplings). One is tension-bending equation group and the other is torsional equation group, and these equation groups are solved independently and the closed-form solutions are obtained. One of the two tension-bending solutions satisfies a simply supported condition, another a fully fixed condition and both of these can be applied to cylinders that are shorter and longer than the length of the bending boundary layer. In addition, it is shown that previously derived bending-boundary solution can be applied as the same form when the length of the cylinder is longer than the length of its bending boundary layer. Moreover, Donnell-type shallow-shell approximation is inappropriate for solving torsional deformation of moderately thick cylinder is also shown. Comparison between closed-form solution and precise solution including both layup anisotropy and transverse shear deformation shows that the difference can be negligible when cylinder is honeycomb-sandwich with symmetric facesheet but cannot be negligible when the symmetric cylinder has non-negligible tension-torsion coupling stiffnesses. Therefore, the estimation formula for the effect of tension-torsion coupling stiffnesses is also derived.

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