A pseudolite can be a solution to strengthen GNSS or create an independent navigation system. However a pseudolite has very serious near-far problem where the operational area becomes smaller and the usage limited. The aim of this work is to extend the pseudolite operational area. To minimize the near-far problem, we split the pseudolite pulse into several pulses. To maintain tracking of weak pseudolite signals, we increased receiver integration time. To validate these solutions, we simulated distributed pulsing and tested the increased integration time. To increase the integration time of conventional and low-cost receivers, we used a data-less pseudolite with no navigation data. Using these solutions, we extended the mid zone of the operational area and reduced the near and far zone.
For the success of hypersonic vehicles, their shape must be optimized to achieve a high lift-to-drag ratio as well as a low aerodynamic heating rate in the hypersonic regime. In addition, the transonic lift-to-drag ratio must also be optimized to realize quick acceleration to the hypersonic cruise speed. The three-dimensional lift-to-drag ratio can be improved even by the two-dimensional section shape (i.e., airfoil) optimization in the region where the sweep back angle is small. Here, prior to three-dimensional shape optimization, a study is done to optimize airfoils of hypersonic vehicles based on these three parameters. At optimization, the hypersonic lift-to-drag ratio is maximized while the transonic lift-to-drag ratio and the aerodynamic heating rate are constrained. The optimum lift coefficient for hypersonic cruise at the maximum lift-to-drag ratio is investigated. The relation between the leading edge radius, which determines the aerodynamic heating rate, and the hypersonic lift-to-drag ratio is also investigated. Results show that to improve the hypersonic lift-to-drag ratio, the airfoil thickness around the leading edge should be small as long as an appropriate compromise with the transonic lift-to-drag ratio is achieved. Results also show that the optimum lift coefficient for hypersonic cruise is much lower than that for typical supersonic vehicles. Small cruise lift coefficient suggests that the wing loading of a hypersonic vehicle should be small. The leading edge radius should be determined by a compromise between the hypersonic lift-to-drag ratio and leading edge heating. Airfoil optimization can provide an appropriate initial guess of the three-dimensional optimum shape. By using an appropriate initial guess, the computation time of the three-dimensional shape optimization is expected to be reduced.
To clarify tonal noise generation, an experimental study on airfoil tonal noise was undertaken using a conventional wind tunnel, which allows acoustic reflection on test section walls. A two-dimensional wing model with the NACA0015 cross-section was used at 5 degrees angle of attack. Most previous experiments conducted in anechoic environments commonly show that the tonal noise frequency is selected in an overall trend of U1.5 (U is uniform velocity) locally consisting of a step-like structure, and Tollmien-Schlichting disturbances are rapidly amplified in the backflow region near the trailing edge of the pressure surface. The present experiments in an acoustically resonant environment show that the tonal noise emanates in accordance with the aforementioned features. However, the ladder-like structure has a different local slope from that observed in anechoic flow. These characteristics suggest that acoustic resonance does not play a fundamental role in tonal noise generation. Observation by hot-wire and smoke visualization techniques shows that unsteady disturbances rather than Tollmien-Schlichting waves are rapidly magnified by the Kelvin-Helmholtz instability in the backflow region. The frequency selection mechanism at tonal noise generation still remains unsolved.
An experimental study was performed to investigate the flame-holding of a hydrogen jet injected into a supersonic cross-flow interacting with an incident shock wave. Pre-burned rich hydrogen or air was injected into the supersonic airflow at Mach 2.5. The injection pressure was 1.2 MPa or 1.6 MPa. The deflection angle of the shock generator was 10°. Velocity profiles near the wall and of the recirculation zone around the injection slot could successfully be measured using the PTV method newly-devised in this study. The velocity profiles showed that the reattachment point of the recirculation zone downstream of the injection slot moved further downstream when the incident shock wave was introduced downstream of the injection slot. The reattachment point with the incident shock wave at the injection pressure of 1.6 MPa was not very different from that of 1.2 MPa. However, the extinction limit at the injection pressure of 1.6 MPa extended further than that of 1.2 MPa.
One of the most critical technical issues with regard to supersonic commercial transportation is the sonic boom that occurs during supersonic cruising flight, which causes impulsive noise on the ground. The “supersonic biplane theory” has been proposed to reduce the sonic boom. Shock wave interaction and cancellation between the wings of a supersonic biplane can be realized at a specific design Mach number, but does not work at off-design values. Here, the low-speed aerodynamic performance, as off-design performance, of a baseline supersonic biplane was investigated and discussed using experimental and computational fluid dynamics approaches. The thin airfoil stall characteristics of a supersonic biplane were shown to be caused by the stall of both upper and lower wings at an angle of attack of 20°. Although there was leading flow separation of the upper wing at lower angles of attack, the stall of the lower wing was suppressed by interference with the upper wing. The lift of the lower wing was almost dominant to produce the lift of the supersonic biplane in the low-speed range. However, the lower wing caused greater drag than the upper wing at higher angles of attack.
A new method for the attitude control of a flapping-wing aircraft is proposed. In this method, the variations in wing deformation, that is, the feathering angle and the camber, are controlled by pulling the wing at a certain point with a thread connected to a servomotor. The experimental setup for verifying the practicability of this method was developed, and aerodynamic forces and wing deformation were measured. It was concluded that thread control caused effective wing deformation, and the variation in the deformation generated the pitching moment that controls the attitude of a flapping-wing aircraft.