A New H∞ control theory which uses a plant description of the normalized left coprime factorization (NLCF) is applied to a pitch attitude control of the short period mode of airplanes. In this theory, the selection of pre and/or post compensators called shaping functions is of practical importance to determine the closed-loop characteristics. This paper presents that the closed-loop poles determined by the theory are given as the eigenvalues of two matrices, one for the augmented plant poles and the other for the controller poles. It is shown as well that the closed-loop poles of the augmented plant are the same as those of the linear quadratic regulator (LQR) theory. These properties enable us to design a compensator that takes not only frequency domain properties (e. g. robustness) but also time domain ones (e. g. output step response) into consideration. Some trade-off is necessary between robustness and fast, less overshooted output responses.
Low-speed aerodynamic characteristics and effects of the blockage ratio of an axisymmetric capsule configuration are experimentally investigated. The flow conditions in a closed-type test section with high blockage ratio are obtained by setting a circular duct coaxially in the open-type test section of the wind tunnel. The aerodynamic loads of similar scale model of three kinds with body diameter of 100, 180 and 240mm are measured in both open-type and closed-type test sections. Also the pressure distributions along the model surface and along the duct wall are measured. From the investigations of experimental results, the following major conclusions are obtained on aerodynamic characteristics of such capsule-type configurations: (1) It is expected that the drag force of capsule-type configuration is strongly affected by blockage ratio effects. (2) The usual method of the wind tunnel wall effect correction is applicable to prediction of lift and pitching moment coefficients, but not sufficient to drag coefficient. (3) The velocity profiles in the closed type test section are much different from those in the open-type test section. In case of the high blockage ratio, the flow which has been accelerated because of the decrease of the section area, is bent by the curvature of the model tail and separates at the trailing edge. Such phenomena make the flow field complicated and result in the increase of drag force.
There are a number of methods for predicting aerodynamic performance of WIG (Wing-in-Ground) effect vehicles. This paper presents a thorough-going critical survey over them. Whole methods are divided into three categories, empirical method, “classical” theories, and “Split Orifice Flow Model” The basic performance criterion may be lift/drag ratio, while the simpler and easier one is the ratio of effective to geometric aspect ratio Aeff/A. There are significant variety for selection of source parameters causing the ground effect, such as B/h, c/h or the midway [(B/h)(c/h)]1/2 Most methods agree fairly well each other, within practically important domain, where 2<e<5 and 10<B/h<50. A Russian empirical chart suggests existence of upper limit for L/D of WIG, which is about 32.
Wingtip vortex breakdown process has been investigated experimentally by hotwire measurement and smoke visualization. On the shear layer which consists of a vortex sheet starting from the leading edge and the boundary layer developing on pressure side surface, unsteady regular oscillation starts to appear at the inflection point of the streamwise velocity profile. As a result, the vortex seems to become unstable due to this inflection point instability in streamwise direction. Therefore, such unstable flow condition seems to play an important role on the wingtip vortex breakdown process.
Radiative heat transfer from the strong shock layer generated around a hypersonic flight model is experimentally investigated by using a ballistic range (two-stage light-gas gun). A plastic projectile of 1.2cm diameter is accelerated in this facility up to 5.3km/s (M=15), and the radiation spectra and global power emission from the projectile are measured. As a result, radiation spectra from some carbon-containing molecules such as CN and CO, were observed along with the air spectra, and the radiative heat transfer rate was estimated. Since the projectile launched in a ballistic range is two orders of magnitude smaller than the actual space vehicles, the model-dimension effect on shock layer structure is examined by analytical and numerical considerations, and a scaling law has been developed. Radiative heat transfer rate around a projectile was calculated, and the validity of conventional numerical model is discussed by comparing the computed result with the measured one.