The forced ignition instability and combustion instability in a scramjet engine were investigated experimentally and numerically using modal analysis. In this study, proper orthogonal decomposition (POD) was applied to OH^* chemiluminescence image data obtained during previous supersonic combustion experiments. In this scramjet combustor, a micro-rocket torch was installed in the cavity. First, to clarify the combustion instability phenomenon in the cavity and flow paths, a POD analysis was performed using the OH^* chemiluminescence image data. The maximum power spectrum density peak shifted to the high-frequency region in high-order modes. The combustion oscillations between 100 and 500 Hz were attributed to injector flame feedback from the fuel or torch jet gas because the fuel and torch gas were directly injected into the shear layer, where fuel combustion occurred. Next, to clarify the forced ignition instability phenomenon in the cavity, the POD modes near the torch jet orifice were investigated. The low frequencies near the torch jet orifice within the cavity were likely associated with the torch gas injector flame feedback because the frequency spectra of the inner pressure oscillations in the micro-rocket torch matched the frequency spectra peak near the torch jet orifice in the cavity.

This study focuses on the measurement of the flame base in a gaseous H₂/O₂ coaxial jet diffusion flame under high-pressure (1.0–5.0 MPa) rocket combustion conditions. OH(2,0) band excitation was utilized for OH-planar laser-induced fluorescence (OH-PLIF) measurement to obtain the flame cross-section. The injection Reynolds number was varied under each pressure to examine the flame structure variation. Numerical simulation was also conducted to visualize the injector vicinity flow field. Results showed that the OH-PLIF measurements had sufficient signal intensity for evaluation. An analysis of the OH-PLIF images revealed variations in the flame thickness and the position of the flame edge under each condition. A numerical simulation of the cold flow was also performed to aid analysis of the OH-PLIF results.

The vertical drag or download on the structure under the rotor of a helicopter caused by the downwash deteriorates hovering performance. On the other hand, the partial ground effect provided by the structure under the rotor improves hovering performance. The design of the fuselage planform should simultaneously consider these aerodynamic interactions, namely the effect of download and the ground effect. This is more important for winged helicopters, which have a large planform area. In this research, as a first step, the download and the partial ground effect of a disk-like structure are investigated to obtain general knowledge of aerodynamic interaction. Experimental parameters are the disk radius, the distance between the rotor and the disk, and the root cutout of the rotor. It is revealed that when the root cutout is relatively small, if the disk radius is up to about half the size of the rotor radius, hovering performance does not deteriorate. Then, as a second step, it is shown that this general knowledge of the disk can be successfully applied to the problem of determining the flap width of winged helicopters.

One of the most important components for air traffic is the runways, which can act as a bottleneck for the arrival flow and cause delays. At Tokyo International Airport, the arrival flow is managed according to the characteristics of air traffic flow from two directions to two runways. This study focused on the effects of changing landing runways on flight times and taxiing times. A queuing model was derived for each runway from the statistical results of surveillance radar data during the busy hours. The flight times depending on changing the landing runway were estimated by the queuing models. To show the effectiveness of each landing runway change, a performance index averaging the times was minimized under the constraint of the number of landing runway changes. The results revealed that some landing runway changes effectively reduced times as intended, while others increased times or had little effect.

Shock wave formation in aircraft drastically retards the maneuverability of aircraft by increasing drag, destabilizing buffet, and diminishing engine efficiency. Conventional method of shock wave strength reduction by structural modifications are cumbersome and lack scalability and the repurposability is difficult to incorporate. The aim of this study was to investigate the effectiveness of nanomaterial coatings in reducing shock wave strength around airfoil surfaces. A nanocoating of graphene was applied to the surface of a TsAGI S-12 airfoil and the aerodynamic characteristics were evaluated. The flow visualization captured from the wind tunnel experiments at various Mach reveals 18% reduction in shock wave angle. The reduced shock wave angle and pressure coefficient confirmed that flow stability increased with increase in Mach, emphasizing that the technique is highly effective for supersonic speeds. Furthermore, to investigate the role of surface roughness, AFM measurements were performed, revealing that the surface roughness decreased from 10 to 2 nm. Notably, this decrease in surface roughness has a direct and vital influence on the reduction of shock wave strength. Specifically, the smoother surface offered by the nanocoating minimizes the irregular reflections that occur on rough surfaces, thereby dissipating the shock wave more effectively. The observed reduction in shock wave angle, surface drag, and static pressure has transpired as a consequence of the reduced surface roughness upon nanomaterial coating. Nanomaterial surface coating is a simple and highly effective method for shock wave strength reduction with scalability that can be readily employed in the aircraft industry.

A combination of water and aluminum powder has been studied for space propulsion. These propellants are particularly suitable for chemical micropropulsion. However, development is being delayed due to many constraints, mainly: safety requirements and system restrictions. This study measured the combustion pressure and reacted aluminum fraction of pulsed combustion using water vapor and aluminum powder under conditions that satisfy the following constraints: water vapor pressure under 100 kPa, aluminum powder diameter over 10 µm, and a combustion chamber smaller than 1 L. In the experimental combustion chamber, the water vapor flow had a mass flow rate of 23.3 mg/s and a velocity of 1.6 m/s. The aluminum powder was lifted and mixed with the nitrogen gas and injected into the water vapor flow and ignited by a spark. The result was organized using an oxidizer-to-fuel ratio in the combustion chamber. When the ratio was around 0.5, the maximum pressure and reacted aluminum fraction marked significant values for micropropulsion: 191.6–313.4 kPa and 0.18–0.36, respectively. The results revealed that the nitrogen gas promoted combustion by diffusing the powder rather than suppressing it, and the importance of the controlling powder diffusion and the ignition timing.