In a conventional method for attitude determination of a spinning satellite by processing so called cone angle data, iterative calculation is used to minimize the weighted squares of the difference between the observed and calculated cone angle. This method requires a considerable amount of calculation. As another conventional method, KALMAN filtering technique is used. This method requires the knowledge of the satellite attitude dynamics. In this paper, a new algorithm for determining the attitude of a spinning satellite is proposed. The proposed algorithm offers a simple method for processing cone angle data of earth sensor and makes calculation program extremely simple and reduces the required CPU time remarkably. A method for estimating wobble angle and phase using sun and earth sensors is also proposed. An example of the results of its application to CS satellite is presented. This example shows that the earth sensor data of 2/3 day period is sufficient to estimate the spin axis direction and sensor bias.
This paper presents experimental studies on the gust response of a wing at stalled region. Although there have been many published papers on the dynamic stall of oscillating wings, few are found on the dynamic stall in the gusty air. In this paper, the experimental approach on the dynamic stall problem of a wing in the gust is performed by wind tunnel tests. Dynamic pressure distributions upon wing surface are measured through the pressure transducer. The wing is two-dimensional, and has NACA-0012 as its cross sectional form. It is installed in the gusty flow (sinusoidal normal gust) at large angle of attack (12-30°). The induced angle of attack by the gust is about 2 or 3 degree. The detail processes of the dynamic stall are made clear from these pressure distributions. The lift forces are obtained by the numerical integration of pressure distribution. The hysterisis loops of the lift force clearly show the characteristic over-shoot and under-shoot of the dynamic stall.
Optimal control theory is used to determine thrust and glide path angle programs for minimum time or minimum noise landing approach of aircraft modeled by an approximated point-mass dynamics. Necessary conditions for optimal programs are derived and numerical results are presented for a small jet airplane. Significant noise abatement and savings in time are obtained compared with reference thrust and path angle programs. The minimum time solution indicates; 1) thrust is a dominant control, and glide path angle γ has little influence on the performance index; 2) Optimal path mainly consists of a constant velocity descending path, which is a singular solution, and a decelerating path with γmax. Optimal thrust switches from its maximum to its minimum on the latter path. The minimum noise solution indicates; 1) optimal thrust has neither its maximum nor its minimum, and varies continuously so as to achieve a smooth EPN level curve; 2) optimal path mainly consists of a decelerating horizontal path and a decending path with γmax. These results may be regarded as a theoretical base for a two-segment approach for noise abatement and time savings.
In this paper, model reference adaptive control systems (MRACS) theory is applied to the design of an aircraft control augmentation system. The most attractive feature of MRACS over other adaptive control systems is that a parameter identification mechanism is not required at all. The problem considered here is to have the aircraft output cp* follow a model (c* model) output cm*, where the model is defined to meet a new handling qualities criterion, namely, c* criterion. First, an error equation of motion is derived from the model and aircraft equations. Then, the stability of the above error equation system is confirmed by LIAPUNOV'S direct method and the KALMAN-YACUBOVICH Lemma. The block diagram obtained in this manner contains the second order c* model, seven adaptive gains (variable) and three state variable filters which were employed by MONOPOLI to avoid error derivatives. The adaptive gain scheme is somewhat analogous to the well-known P-I type controller used in many classical applications. The adaptive control law employs the proportional plus integral of the product of the error and other signals which are filter outputs, plant output cp*, elevator deflection, etc. A digital computer simulation was performed at three flight conditions. The result proved that the present system has a great adaptability and that the cp* responses are satisfactory. However, normal acceleration response nz is poor at low airspeed, which fact indicates that c* criterion is not necessarily complete.