Vortices generated in a solid rocket motor may cause pressure oscillations, which cause problems such as deteriorated flight performance and payload damage. The building-cube method (BCM) solver is expected to accurately predict such vortex flows using high-order schemes as well as easily manage complex configurations. The aim of this study is to establish an analysis method for the internal flow of a solid rocket motor using the BCM. Computational results obtained in this study are compared with analysis data based on other solvers as well as experimental data for validation. The BCM solver is applied to the validate the converging-diverging verification nozzle, upper jet model, and propellant surface ejection model. Furthermore, it effectively captures the three-dimensional (3D) trailing vortex downstream of the inhibitor and the propellant surface injection flow. In general, the 3D BCM solver accurately reproduces the internal flow field of the solid rocket motor.

The benefit of Boundary Layer Ingestion (BLI) was evaluated by wind tunnel test and analysis. Using a simple engine/airframe model which is composed of an airfoil based on NACA0012, an electric fan and an electric motor, wind tunnel test was conducted. Net streamwise force and electric power were mainly measured. Based on the power balance method originally described by Drela, the concept of conversion efficiency was added to separate demerit of BLI. By using conversion efficiency, mechanical power was changed to electrical power and making it easy to understand benefit of BLI. This analytical method was fit with experimental data and well represent BLI benefit. At cruise condition in 40m/s, BLI configuration needs lower 7.66% electrical power than nonBLI configuration.

The characteristics of active-grid-generated flows are investigated. The streamwise flow velocities are measured using a hot-wire anemometry. The turbulence intensity, integral length scale, and power spectrum density are evaluated to understand the influence of active-grid-operating parameters such as the rotation rate of the grid's agitators and the duration of the agitator's constant-rate rotation. The effects of randomized rotation and the test-section type are also investigated. It is found that the randomization of the rotation rate improves the spectrum distribution of velocity fluctuation, while the statistical quantities of turbulence are not sensitive to the randomized rotation rate. The randomized duration does not affect the generated flow characteristics. The turbulence intensity and integral length scale can be correlated with a non-dimensional parameter expressed by the rotation rate, mesh size, and free-stream velocity. When the non-dimensional parameter is small, the test-section type shows no influence on the turbulence intensity and integral length scale.

The development of energy regeneration system and the result of flight demonstration were described in this paper. The regeneration system was designed and integrated to regenerate potential energy of aircraft to electric energy. The autors experimentally clarify the conditions to maximize regeneration efficiency. The display system and a power lever included as pilot interface in the system are configured to adjust the regenerated power without any increase of pilot workload. As the result of test flights in 2015, functions of the system are properly demonstrated and succeeded to realize stable `Regenerative Soaring' for the first time in the world.

A method of wind tunnel testing is proposed for generating streamwise velocity disturbance in low Reynolds number flows. When a model moves along the free stream at a constant speed, the effective turbulence intensity for the translating model can be adjusted through the change in the relative velocity between the free stream and the model. The feasibility of the proposed method is experimentally verified in a wind tunnel with a linear motion mechanism installed in the test section. The experiment shows that the effective turbulence intensity for the translating model increases when the model moves along the free stream. Agreement in the lift coefficients between the tests with and without using the linear-motion mechanism confirms the feasibility of the proposed method. The experimental range available in turbulence intensity can be extended using the model streamwise translation. Moreover, a stepwise gust can be generated while the turbulence intensity remains almost constant when the model moves against the free stream.

Wind tunnel experiments using several turbulence grids are conducted to investigate the effect of turbulence intensity on the nonlinearity of the aerodynamic characteristics of a NACA0012 wing at a low Reynolds number. A semi-span model with an aspect ratio of 3 is used. Lift and pitching moment curve slopes are evaluated to discuss the nonlinearity of the aerodynamic coefficients. The angles of attack at the local maximum and minimum of the pitching moment correspond to the ones at which the lift curve slope changes. Therefore, the stall angle can be defined by the local pitching moment maximum, even in the case that the change in the lift coefficient at the stall is obscure. For turbulence intensity of 1.9% and above, the change in the lift curve slope at around 0 deg is not observed, and a local maximum of the pitching moment coefficient at low angles of attack disappears. As the turbulence intensity increases, the constant-lift-curve-slope region enlarges. Hence the nonlinearity of the aerodynamic coefficient decreases in a more-turbulence environment. However, the nonlinearity of the aerodynamic coefficient does not disappear for turbulence intensity below 3.7%. The turbulence intensity significantly affects the aerodynamic coefficient and its nonlinearity at 1.0% or fewer turbulence intensity.

In sample return missions, particularly those to small bodies, landing operations have been performed by a single spacecraft. However, there is a potential risk of spacecraft breakdown in this approach. This paper proposes a new framework for sample return missions, in which a touch-and-go sampling probe (TAG probe) equipped with solid rocket motors is used. In the proposed framework, the TAG probe is released from a mother spacecraft, lands on its target body, collects samples, lifts off, and docks with the mother spacecraft. Because the mother spacecraft can stay at a high altitude and only the TAG probe performs risky operations, the possibility of mission failure can be drastically reduced. This study focuses especially on the landing phase and provides the following two major considerations. First, suitable design of solid motors for various celestial bodies is surveyed. Second, a control system for gimbal actuation of the solid motors is developed to enable soft landing. Two operation methods for gimbal control are proposed. In the first one, only the altitude of the probe is controlled through gimbal actuation. In the second one, the attitude motion of the probe is controlled simultaneously. The landing operation can be successfully completed if the controller is designed to converge faster than the motor's burning duration.

This paper introduces Microsatellite-Friendly Multi-Purpose Propulsion (MFMP-PROP) we proposed as a promising propulsion system using low-toxic propellant for microsatellite, clarifies the development policy and positioning of MFMP-PROP among a variety of propulsion system for microsatellites, and reports the current status of its systemization with the corresponding test results. In addition, the high scalability of MFMP-PROP also makes it suitable for a wide range of missions with microsatellites of 2U to 50kg class. Though the performance of MFMP-PROP is not up to the designed value at present because of too large propellant supply in a prototype of MFMP-1U, this will be solved by optimization of the piping or adding flow adjustment elements. We will continue to further systemize MFMP-PROP and conduct environmental tests to achieve an early space demonstration and utilization.

Protuberances on the surface of a space transportation vehicle affect the aerodynamic characteristics, particularly the side force. In this study, wind tunnel tests are performed under transonic conditions (0.7 ≦ M_∞ ≦ 1.3) on a slender body (of slenderness ratio 8.9) with protuberances of varied heights, width, or lengths. The corresponding computational fluid dynamics simulations are also conducted to aid understanding the flow physics and the aerodynamics. The results show that at Mach 0.7, the side force decreases due to the increase in the low-pressure region on the upper surface of the protuberance and also due to the change in the direction of flow around the protuberance. At Mach 1.3, however, the side force barely changes since a shock wave creates positive side force.

In this study, we develop a compressible fluid analysis program that enables fully automatic generation of computational grids using Hierarchical Cartesian grids and immersed boundary methods. However, these methods are associated with low analysis accuracy, oscillations observed in the surface distribution, and violations of conservation laws near an object. This is because the object geometry is represented using a specific algorithm within the flow solver. We demonstrate that the proposed method is sufficiently accurate by comparing its results with those obtained via the conventional analysis method using body fitted grids. We also explain why the surface physical quantity distribution oscillates when the grid resolution is relatively low. In addition, we have clarified why conserved quantities in the vicinity of the object are not conserved and proposed a method to resolve this problem. Herein, we show that mass conservation is strictly followed by the proposed method.

Incoming wakes can induce separated boundary layer instability, and cause performance losses in turbine cascades of a gas turbine. An efficient numerical method that can analyze unsteady incoming wake effects is desirable from practical point of view. In the present study, an extended harmonic balance method (Chebyshev-based harmonic balance, C-HB) is introduced to the analysis of the incoming wake effect on the separated boundary layer. Convection of a one-dimensional wake-like profile and the flow around the flat plate whose front part is replaced by the NACA0012 with the incoming wake at a Reynolds number Re = 92,000 are computed by the C-HB method. The result shows that the C-HB method can resolve the transient phenomenon of wake convection and the associated linear instability of the flow around the blade.

This paper proposes a method to design a multiple-input-multiple-output stabilizing flight control law via controller order reduction. In this procedure, first, Linear Quadratic Regulator (LQR) is designed for a full-order plant model. Then the order of the LQR closed-loop system is reduced by applying balanced truncation to obtain a static output feedback control law, i.e., proportional-control law. The closed-loop order reduction is performed by applying a model order reduction technique called fractional balanced reduction to the controlled system modified by the weighting matrices of the LQR. This design method is applied to linear aircraft models. The simulation results show that the proposed method can design stabilizing flight control laws that provide closed-loop properties close to the LQRs. In addition, the obtained control laws can achieve a control performance better than conventional single-input-single-output roll and yaw dampers and comparable to or better than that designed by a numerical search algorithm and an LMI (Linear Matrix Inequality) approach.