An attitude control experiment of a spacecraft using six single-gimbal control moment gyros is described in this paper. The control moment gyros are arranged in a twin-triangular pyramid configuration, where internal singular states are all passable. By using the passable characteristics of the configuration, a steering law which calculates the gimbal angular rates from the attitude control torque is derived. A ground experiment by the steering law is executed with a three-axis motion simulator and is compared with that by the singular direction avoidance steering law. The usefulness of the proposed steering law is verified by the ground experiments as well as the numerical simulations.

SCRAM jet fuel sinking within compressible turbulent boundary layer was difficult to contribute to efficient thrust production. Then, the aerodynamic instability induced by a low-height wedge named Single-Wedge, which is composed of vertical sidewalls and a vertical backwall, installed on combustor wall was adopted for the passage height recovery of fuel tube being attached to the wall. It was attempted that the three-dimensional complicated flow induced around sidewall of the wedge raises the fuel tube toward supersonic airstream, immediately behind the wedge the mixing state of fuel was experimentally confirmed and the flow structure in the wake was numerically clarified. Since the relatively thick boundary layer inhibited the development of vortical motion in the airstream and confined the fuel tube to the wake flow of the wedge, the fuel-air diffusion progressed only in the horizontal direction with the appearance of two fuel tubes divided by an unique separated airstream from the top of the wedge but the achievement of vertical penetration was insufficient.

The aerodynamic instability induced by a low-height streamwise-symmetrical wedge named Double-Wedge installed on combustor wall was adopted for the passage height recovery of SCRAM jet fuel being attached to the wall. It was attempted that the three-dimensional complicated flow induced around sidewall of the wedge raises the fuel tube toward supersonic airstream, immediately behind the wedge the mixing state of fuel was experimentally confirmed and the flow structure in the wake was numerically clarified. A pair of vertically aligned counter-rotating streamwise vortices was formed and withdrew horizontally a branch part from the main fuel tube. The branch part of fuel holding a firm connection to the main tube extended into main-airstream and achieved successfully the continuous fuel transport from boundary layer to supersonic airstream. In addition, the double-wedge showed superiority to the single-wedge in the aspect of reduction in total-pressure loss.

Many papers have been published on air traffic management research. In those papers, BADA model which was developed by EUROCONTROL is often used, however, few papers have discussed accuracy of the BADA model. This paper investigates the accuracy of BADA model by using cargo flight data which includes weight data recorded more accurately than passenger aircraft. Fuel flow, its statistic tendency and fuel consumption with different Cost Index (CI) calculated by BADA model are compared with flight data. The results show that fuel flow is almost identical for every part of flight phase except for descent phase with flaps and landing gear extended. Furthermore, flight state estimation information is compared with the flight data. Flight state estimation only requires the position information such as surveillance radar data from the ground, numerical weather prediction data and aircraft weight. The results have revealed that flight state information including fuel flow can be estimated accurately in most parts of flight. Next, the optimal trajectories generated by BADA model and numerical weather prediction data are studied as another part of the model evaluation analysis. The optimal trajectory calculated so as to minimize fuel consumption and flight time plays key role in the future four-dimensional trajectory based operation as well as potential benefit analysis. The tendency of fuel consumption and flight time change due to CI is almost agree. The air speeds, however, are different. Therefore, it is difficult to estimate flight time even if the CI is given.

The direct simulation Monte Carlo (DSMC) method using the improved collision scheme (U-system) has been verified to be effective for one-dimensional normal shock wave that includes extreme non-equilibrium over the whole region of the wave, even with the use of a rough cell-network in the previous paper. The purpose of this study is to investigate a bad influence of a very small number of molecules in a cell on the DSMC analysis of the one-dimensional normal shock wave. The paper also studies the mean free path that restricts cell dimensions in the DSMC method. The effective mean free path measured by a moving distance between a series of collisions of one molecule is verified to become larger than the defined mean free path on the condition that the flow has a large velocity.

This paper illustrates a procedure for designing structured robust flight control systems based on Multiple-Delay-Model / Multiple-Design-Point (MDM/MDP) approach, which is one of the useful ways for designing robust flight control systems. The MDM/MDP approach is based on solving a simultaneous stabilization and optimization control problem which minimizes the sum of the stochastic linear quadratic cost indexes weightless on the control inputs. The main contribution of this work is proposing an efficient algorithm for solving the MDM/MDP design problem with less conservatism. Non-convex necessary and sufficient condition of the design problem with static output feedback is derived using an enhanced linear matrix inequality (LMI) characterization. The solution algorithm with guaranteed convergence at least to a local optimal solution is proposed as well, which iteratively solves the LMI sufficient condition of the original non-convex condition by fixing some of the decision variables. This design method is applied to structured roll and side-slip angle controller design. The simulation results show that the proposed method is effective and efficient for robust flight control design with the MDM/MDP approach.

A Control Moment Gyro (CMG) is an actuator suitable for agile spacecraft because it has large torque capability. However, CMG systems can fall into singular states where commanded torque can not be exactly generated and the spacecraft attitude control law becomes complicated. A control method which uses gimbal angle trajectories is effective for the singularity avoidance and the time optimization of the attitude change. In this study, focusing on the spacecraft equipped with four Single-Gimbal CMGs in a pyramid configuration, the procedure of time optimization using the Chebyshev-Pseudospectral method and a nonlinear optimization program is described. In addition, the tracking control method is investigated and the steering logic is demonstrated by conducting hardware experiments as well as numerical simulations.

It is sometimes difficult for aircraft designers to grasp the overall picture during the conceptual design process because the process is highly complicated. Various parameters related to an aircraft, such as design requirements and design variables, entangle each other with complexities. This paper aims to evaluate whether Principal Component Analysis (PCA), which is used to analyze multidimensional data, is applicable to clarify the relationships among design requirements and design variables when a conceptual design example is considered. First, an aircraft conceptual design tool, using a multi-objective genetic algorithm, is constructed. The validity of this tool is confirmed by comparing the design results with an existing aircraft. The design tool is used to create a data set of aircraft. From the data set, five important parameters are chosen and PCA is applied to them. They are classified into two groups based on their characteristics, and the relationships among them are discussed. The results indicate that PCA could be used to provide beneficial information for the aircraft conceptual designers.

This paper describes a testing method to measure acoustic and thrust characteristics of exhaust nozzles using a model engine and the uncertainty analysis of this method. The uncertainty in the measurement of gross thrust coefficient and 1/3 octave band sound pressure levels in the model engine test were estimated to be ±0.6 % and ±0.41 dB, respectively. Although there are a few differences between the reference data of jet noise and the obtained data, it was demonstrated that the proposed method provides both acoustic and thrust data with reasonably good repeatability and also allows us to conduct a simulated flight test of hot-jets. A demonstration using a noise reduction nozzle was also performed to demonstrate the usefulness of the proposed method, and it clearly emphasized the difference between a baseline nozzle and a noise reduction nozzle with the estimated uncertainty.

The dynamics of a spinning solar sail is affected by the vibration of its sail membrane because it does not have supporting structures to avoid the deformation of the sail. For example, attitude maneuver by impulsive thrusting causes coupled vibrations between the main body and the sail membrane. Previous research analyzed this attitude motion by use of a simple vibration model of a spinning sail membrane, called First Mode Model (FMM). They succeeded to simulate the actual flight data of IKAROS even though the FMM does not take higher vibration modes into account. This paper theoretically shows that the first mode vibration is dominant for the attitude motion of a spinning solar sail, based on a complete vibration model of a spinning sail membrane.

The relationship between the shock wave propagation velocity and the discharge power was investigated by applying the laser schrielen technique for the shock tube experiment at the shock Mach number of 2.17±0.03. The time interval of the shock wave arrival between the two laser beams was 160.5 μsec in the case without the discharged plasma. On the other hand, in the case with the discharged plasma, it was 157.8 μsec at 10.2 W and 149.2 μsec at 73.0 W, respectively. These results suggest that the shock wave propagation velocity can be accelerated with increasing the input power. This tendency agrees with the numerical simulation results, at least qualitatively.

The volume of air traffic is increasing year by year. Currently, Air Traffic Control (ATC) ensures safe separation when merging air traffic flows by stretching flight paths using RADAR vectoring. In the future under 4-dimensional trajectory-based operations, ATC will instead instruct an arrival time at a specific waypoint, and aircraft will achieve the arrival time using the Required Time of Arrival (RTA) function of the Flight Management System (FMS). The RTA function controls arrival time by changing airspeed, and with some FMS models it can operate during all flight phases, while others can change only cruise speed. When the arrival time is not achievable by RTA, ATC may provide RADAR vectors or even instruct holding. However, RADAR vectoring is often accompanied by speed instructions that cause an aircraft to operate at above its optimal speed. We therefore suppose that holding is more efficient than vectoring, and that holding procedures with 4-dimensional trajectory will be similar to current holding because holding is a kind of time-based operation. In this study, we compared the fuel consumption of actual flights under RADAR vectoring with simulated fuel optimal flights which instead executed holding. The results show that holding was able can reduce the fuel consumption of all aircraft, but we observed many conflicts at the merging point. We propose different holding altitudes in order to resolve these conflicts, and the proposed modification showed that average fuel consumption can be still reduced with a limited deterioration.

This paper presents a method to derive a minimum-time formation guidance of fixed-wing aircraft which is based on optimal solutions of single aircraft. This problem is conventionally formulated by treating the aircraft as one group which means centralized control. Moreover, the computational time is increased rapidly as the number of aircraft is increased. On the other hand, it is solved by each aircraft with small communication in the proposed method. Further the computational time is not sensitive to the number of aircraft by using homotopy method. Numerical results show almost no difference about the time for the minimum-time formation and the effectiveness of the computational time.

The concept of avoiding complete loss of thrust by a multiplexed motor system was proposed for the purpose of reducing the accidents due to engine failure of single engine aircraft. The developed multiplexed motor system consists of four motor elements and can provide totally 60 kW power. Even if one motor element loses the power during takeoff and climb, the aircraft can safely continue to climb due to the output of 45 kW from the remaining motor elements. The pilot can maneuver with the same feeling as in the conventional aircraft without any discomfort in the multiplexed motor system. The motor system and the pilot interface were mounted on a demonstrator and manned flight tests were conducted. This paper describes about the summary of flight tests in FEATHER (Flight-demonstration of Electric Aircraft Technology for Harmonized Ecological Revolution) project at JAXA and discusses the design method of the multiplexed motor system for aeropropulsion.