This paper describes the experimental study about the aerodymamic forces acting on the 3dimensional body and the flow pattems in incompressible flow. The experiments were carried out setting a spheroidal body inclined up to 30 degrees in a wind tunnel. At low incidence, the surfacepressure distributions are close to that of potential flow, and ring-like separation line is observed at the rear part of the body, At high incidence, a pair of the longitudinal vortices, which steadily stay on the lee side, come into existence and affect the air loads.
Experiments on hypersonic flow around a caret wing and a winged-vehicle are conducted by the shock tunnel with conical nozzle, and Mach number of 8. Three force components and surface pressures are measured for a wide range of attack angle up to 45°, and compared with Newtonian flow theory, giving good comparisons for the present, near flatbottomed models, except pitching moments. The source flow produced by the conical nozzle affects measured pressures and forces, and some corrections are necessary.
A new method of measuring heat transfer rate with high spatial resolution and fast response is developed to be used in wind tunnels with long flow duration. For the development of the new method a multi-layered thin film heat transfer gauge is proposed. The principle of the sensor is based on measuring the temperature gradient across a thin, heat resistant layer of SiO with two, thin film resistance thermometers on its upper and lower surfaces. These thin films and the heat resistant layer are deposited by vacuum evaporation technique. The design consideration, the accuracy of the sensor, frequency response, and the calibration technique are discussed. The sensor is applied to the measurement of the heat transfer rate in turbulent boundary layer on a flat plate at Mach number 4, Tw/T0= 0.64, and Re=1.4×107, and shows high sensitivity and fast response. It also shows excellent durability.
This paper deals with the detection characteristics of the human pilot, who is engaged in a compensatory control, to a sudden change in the controlled element's characteristics. Taking the case where the change manifests itself as a variance change of the monitored signal, it is shown that the detection time, defined to be the time elapsed until the pilot detects the change since it takes place, obtained in an experiment is related to the monitored signal and its derivative. Then, the detection behavior is modeled by an optimal controller, an optimal estimator, and a variance-ratio test mechanism that is performed for the monitored signal and its derivative. The result of a digital simulation shows that the pilot's detection behavior can be well represented by the model proposed here.
The optimum glide trajectories of a reentry vehicle are computed as an optimal control problem with inequality constraints on functions of the state and control variables. Inequality constraints are used to simulate the stall, pitch trim and load factor limits. Numerical examples contain the three dimensional turns to maximize the terminal velocity with the terminal location prescribed. Results obtained demonstrate a rational way to reflect various flight limitations on the nominal flight-path determinations.
A combustion and/or pressure-driven shock tunnel with a conical nozzle is constructed. The tailoring Mach number MST of the combustion tumel is from 6.7 to 7.4. The heat transfer rate of sharp leading edge flat plates with a backward facing step in hypersonic flow of Mach number M∞=7.05 and 11.1 are measured by use of the pressure tunnel, and it is shown that the magnitude and the location of the peak heat transfer rate at reattachment region behind the step seem to be affected by the magnitude of M∞. The measured heat transfer rate of a sharp flat plate with angle of attack in M∞=11.1 hypersonic flow from the conical nozzle of the pressure tunnel is in qualitatively proper agreement with the author's analytical result. The behavior of the heat transfer rate near the sharp leading edge of anat plate in M∞=11.1·high stagnation temperature To=7, 900K flow measured by using the combustion tunnel shows tolerable agreement with the existing measured results at lower To in the strong viscous interactlon region.