The taketonbo is a very famous flying and traditional flying toy from old times. However, the flight characteristics was not yet studied. The flight of the taketonbo is decided by the initial condition. The initial condition is decided by the state of open air and the movement of hands at the time of throw. To obtain the accurate flight characteristics of the taketonbo, the initial condition must be kept constant. The flight test performed by using an experimental device which generates a spinning motion only. In the experiment, the taketonbo was thrown out to vertical direction with the experimental device, and it was pictured from the vertical and horizontal directions with two high speed video cameras. Flight data of the taketonbo were analyzed and the basic flight characteristics of the taketonbo were clarified.
There is a concern that a thermal stress in a solar array panel under cryogenic environment would damage a solar cell and its interconnector, which might result in a drop of electrical power. Therefore it is important to perform the verification test on the ground to assure the healthiness of solar cells in orbit. In this study, we have developed the health monitoring method for solar array panel using Fiber Bragg Grating (FBG) sensor which is one of optical fiber sensors. In this paper, to reveal the thermal deformation behavior of the substrate composed of very thin Carbon Fiber Reinforced Plastics (CFRP) faceskin and aluminum honeycomb core in cryogenic environment, we measured the strain of CFRP faceskin near the center of honeycomb cell and showed that the depth of dimple increased linearly as temperature lowered. In addition, to confirm the effectiveness of our health monitoring method, we measured the internal strain of solar array panel and showed that we could detect its internal damage by strain change, and we observed the cross section of the damaged area and identified the damage point.
Mechanism of drag reduction acting on a rectangular cylinder with recessed corners at its front edges was investigated in a low speed wind tunnel by means of flow visualization and PIV technique. The aspect ratio of the recessed corner h/t was changed for rectangular cylinders with aspect ratio of ranging from 0.45 to 2.05. As an indicator of drag reduction, the back pressure of the rectangular cylinders was measured. The results show that the back pressure increases to the maximum value for all rectangular cylinders when h/t is about 1. The flow visualization reveals the mechanism of the drag reduction as follows. The edge of the recessed corners makes the separated shear layer turbulent. It results in suppressing the vortex formation in the wake behind a cylinder. This vortex suppression reduces the curvature of the separated shear layer from the other side of the cylinder and increases the back pressure. As the result, drag of the rectangular cylinder is decreased.
A microphone array comprised of 195 microphones was expanded at Noto Airport, and flyover tests on noise source localization around the aircraft were carried out with JAXA flying test bed ‘Hisho’. Analysis condition of time domain beamforming method is adjusted properly, which improves both resolution of noise source and signal-to-noise ratio. In the landing configuration, a nose gear, main gears, and edges of flaps are proved to be main noise sources. Integration of noise source maps obtained by beamforming gives good agreement with sound pressure level measured by a microphone placed at the center of the microphone array.
Effect of nozzle configuration on the takeoff noise of a hypersonic transport is investigated based on small-scale nozzle experiments. In the Japan Aerospace Exploration Agency, the precooled turbojet (PCTJ) engine aimed at the hypersonic flight is under development. The takeoff noise of the full-scale vehicle is assessed before it is put into the operation using an acoustic simulation method. By matching the jet velocity and jet Mach number with the realistic PCTJ engine condition, the acoustic characteristics of the hypersonic nozzle are duplicated with a small-scale nozzle model. The jet noise database around the hypersonic nozzle is established through a number of experiments, and by taking into account 1) the nozzle scale, 2) the number of nozzles, 3) the direction of installation, 4) the measurement distance and angle based on the flight trajectory, and 5) the atmospheric absorption, the effective perceived noise level (EPNL) at the flyover point is calculated. It is shown that the nozzle configuration has a large impact on the takeoff noise, and that such an assessment could contribute to the improvement of the integration design of the vehicle.
Experiments were conducted to investigate effects of vortex generators on flow separation for a cranked-arrow wing at a low speed wind tunnel. A kind of delta-shaped vortex generator was applied to control the boundary layer separation flow near the hinge line of the deflected leading-edge flap. Effects on aerodynamic performance were estimated according to the force measurement data and flow patterns were observed by oilflow visualization technique. It was found that the drag coefficient was reduced and the lift-to-drag ratio was improved as the vortex generator was suitably equipped on the hinge line of the leading-edge flap. It was observed that swirling longitude vortices were generated from the leading-edge of the vortex generator, and thus suppressed the boundary layer flow separation.
For practical CFD-aided design of riblets on aircrafts, we are aiming to develop a Reynolds-Averaged Navier-Stokes (RANS) turbulence model which can simulate drag reduction effects without resolving fine-scale flows near the riblets. Wilcox's rough wall boundary condition for Menter's SST model is modified to reproduce velocity shift in logarithmic region corresponding to riblet's drag reduction effects. Two basic relations for this model are derived from available experimental data and parametric analysis, and are validated in zero-pressure gradient boundary layers with two different-geometry riblets. Drag reduction rates corresponding to experimental results can be obtained with these relations. The new model, however, underestimates the effects of riblet mounted on NACA0012, which suggests that further improvement of the model is needed to account for influences of pressure gradient.
In an effort to realize efficient arrival operations, one of the challenges directs at realizing energy-saving arrivals is Continuous Descent Operation (CDO), with which the arrival aircraft descends to the airport continuously using near-idling thrust. As a potential solution to achieve both time-spacing performance and fuel reduction while conducting CDO in arrival traffic, this paper suggests applying ``Fixed-flight Path Angle (FPA) descent,'' which the arrival aircraft continuously descents to the assigned runway following the fixed vertical path angles. The feasibility of the FPA descent was evaluated via a series of B777-200 full-flight simulator experiments, eleven trials in total. With the use of airline's flight simulator, the pilot's operability and tracking performance following the assigned FPA vertical path were assessed. The effectiveness of the combination of FPA descent and speed control is clarified via a comparison with conventional arrival operations. These simulation results show the applicability of the proposed FPA descent and suggest its future contributions.
This paper presents an experimental study on the identification of impact forces acting on an isogrid-stiffened panel which consists of a thin skin and ribs arranged in a repetitive equilateral triangular pattern. The locations and the force histories of the impacts applied to the panel are identified using the responses of surface bonded piezoceramic disks by employing an identification method based on experimental transfer matrices. The experimental transfer matrix is adopted to relate the force history of the impact to the time history of the response of the PZT sensor, and it is constructed utilizing the measured data obtained by conducting preliminarily impact tests. When the panel is subjected to unknown impacts, the impact locations and force histories are identified so that the sensor responses calculated with the experimental transfer matrices match the measured responses of the PZT sensors. To verify the validity of the present method, the identification of impact forces acting on an isogrid-stiffened aluminum panel is performed experimentally. Here, the identification is demonstrated for a single point impact and also for impacts acting at two locations. The results reveal that the identified impact locations and the force histories agree well with the measured ones.
Aerodynamic performance of a helicopter rotor hovering above or near the obstacles with penetration boundary condition such as grove is studied experimentally. At first, it is shown that effects of recirculation flow in the experimental facilities are not essential on the ground effect of hovering rotor if the model rotor and ground effect plate are properly positioned. The obstacles with penetration boundary condition are simulated by plural rows of pipes. The experiment proves that the required torque coefficient of the hovering rotor near the obstacles with penetration boundary condition remarkably increases compared with that without obstacles.
We applied a method based on Complex-ray theory (CRT) to the stability analysis of three-dimensional boundary layers at supersonic flight test conditions of the JAXA's experimental airplane called, NEXST-1, and investigated its capability. The dispersion relation was numerically obtained by the eigenvalue equations and linear parabolized stability equations (PSE). By adapting two kinematic wave theoretical models, we found out which wave type was dominant at measured transition locations of the flight test in 2005. As a result, it has been found that the dominant wave type on the outer wing section of the NEXST-1 was wedge-shaped disturbance, on the other hand, it was two dimensional wave-packets on the inner wing section. From the view point of transition prediction using an eN method, there was no difference of transition N factors (Ntr.) between the results by the eigenvalue solutions with and without CRT method. Only a use of PSE with CRT, however, was able to reduce the discrepancy of Ntr. among cross sections, owing to the nonparallel effect of flows.
This paper is concerned with the design of Stability/Control Augmentation System (S/CAS) for small Quad Tilt Wing (QTW) Unmanned Aerial Vehicle (UAV) which has been developed as a test bed for guidance and control system demonstrations. We design the basic S/CAS for the aircraft using our previously proposed design method. By comparing the eigenvalues of linearized motions of the newly developed and the previously developed QTWUAVs, the effectiveness of our method for S/CAS design is illustrated. We also show the flight test results using our S/CAS.