Overall control to improve the control characteristics of an aircraft, CA (Control Augmentation), is used to realize the desirable motion of the aircraft in relation to the pilot’s control action. C∗ criterion is an important factor for the pilot’s preferred longitudinal motion. The time history of C∗ corresponding to the step input is specified within the upper and lower envelope, and it is desirable to be near the center of the envelope. In this research, the control laws of control augmentation for small supersonic aircraft were designed with the use of fuzzy logic control to obtain the C∗ response near the center of the envelope. The evaluation of the designed control laws showed good performance in all flight conditions. Here the control laws were varied by only one scaling factor for dynamic pressure. This means that virtually no gain schedules by the Mach number and the angle of attack are necessary. This paper shows that fuzzy logic control is an effective and flexible method when applied to control laws for the control augmentation of aircraft.
An understanding of the aerothermophysical feature of the high-enthalpy flow is important for designing atmospheric hypersonic flight vehicles based on the ground-heating test results. Nonequilibrium temperatures of nitric oxide of the arc-heated airflow were measured by the optical diagnostic method in the Institute of Space and Astronautical Science (ISAS) arc wind tunnel facility as a part of the flow characterization. Nonequilibrium temperatures of NO γ-band (A2Σ–X2Π1⁄2) measured experimentally by 2-line and multiline laser-induced fluorescence spectroscopy techniques exhibit separation between vibrational and rotational temperatures. The temperature relaxation along the flow has been compared with the emission spectroscopic data.
A blended-wing-body airplane is in essence a flying wing configuration. The flying wing is regarded as an alternate configuration to reduce drag and structural weight. This paper presents the design process of a medium size Blended-Wing-Body airplane. Because of the abrupt change of thickness and chord distribution, it requires special design tools. In the present study, the wing surface is generated by the use of the RAPID method. Takanashi’s inverse design method is used to obtain the airfoil shape. The required target pressure distribution is specified by the constrained target pressure specification technique. The constraints include the space requirement and the requirements on the aerodynamic performances. Our study shows that the proposed design tools overcome the design difficulties.
A new design method of a robust guidance law for missiles is presented. It has two features. One is that the guidance law is designed based on the heuristic idea that keeping the line-of-sight angular rate small can make the miss distance small. The other is that a linear robust control method, i.e., the μ-synthesis, is employed. When these are incorporated, uncertainties and disturbances in the homing system can explicitly be taken into account in the design to achieve the control or guidance objective. Specifically, the uncertainties and disturbances considered here include time delays in the missile dynamics, range variation between missile and target, measurement noise of the line-of-sight angular rate, and normal target acceleration. The guidance law obtained by this approach is a 4th order dynamic compensator requiring the line-of-sight angular rate as the only measurement. The miss distance is evaluated through nonlinear simulation. The simulation study shows that the proposed guidance law is generally superior to the proportional navigation guidance law and is also superior or equivalent to the suboptimal guidance law in miss distance.
The base-flow characteristics of aerospike nozzles are discussed based on the results of numerical simulations. The nozzle is truncated at a 20% portion of the isentropic length, and the condition both with and without external flow is considered. The computed results realize that the ambient pressure influences the base pressure when the pressure ratio is low and becomes independent from the ambient pressure when the pressure ratio is high, as pointed out in previous studies. Although the ambient pressure is mainly transferred to the base region by the envelope shock wave, some computed cases show that the ambient pressure is not transferred to the base surface even when the envelope shock wave impinges on the subsonic part of the base region. Two high-pressure regions, one created by the shear layer impingement of the nozzle exhaust flow and the other by the envelope shock wave impingement, appear on the nozzle axis. The flow stagnates either of the higher pressure regions of these two, and the strong reverse flow toward the base surface occurs. The ambient pressure is transferred to the base surface when the flow stagnates at the high-pressure region created by the envelope-shock impingement on the nozzle axis. It occurs when the pressure ratio is low. The base pressure becomes independent from the ambient pressure when the stagnation point changes to the shear-layer impingement on the nozzle axis. It occurs when the pressure ratio is high.
This paper discusses the static stability concepts of airplanes. The definition of static stability has not been systematically discussed. For example, although “Cmα<0,” “Clβ<0,” and “Cnβ>0” are merely the aerodynamic criteria corresponding to each static stability, they are used just like the definition of static stability. Moreover, “flight path stability” is well known as the static stability concept on one translational motion, but the stability concepts in regard to the heaving and sliding motions of an airplane are indefinite. And though “tuck under” and “flight path divergence” are also static stability problems, they are treated not systematically but individually as special problems. So, this paper proposes a definition system consisting of six static stability concepts. Moreover, in the first report of this paper, “static flight speed stability” and “static angle of attack stability” are discussed in detail, and their stability criteria in terms of the aerodynamic characteristics of an airframe are obtained.
This paper discusses the static stability of airplanes. The conventional definitions of static stability are neither clear nor systematic. On the basis of these situations, in the first report of this paper a system of definitions of six static stability concepts was proposed. “Static flight speed stability” and “static angle of attack stability” were discussed in detail, and their stability criteria were obtained in terms of the aerodynamic characteristics of an airframe. In this second report, the definitions and the criteria on static pitch stability are discussed in detail, and besides “dδe⁄dΘ<0,” the derived form of definition “dδe⁄dU>0” is deduced from the fundamental definition “dM⁄dθ<0.” The general criterion is also obtained from the fundamental definition. This criterion gives “Cmα<0” as the aerodynamic requirement in a normal low subsonic flight region and in each condition in which the phenomena of “flight path divergence” and “tuck under” arise.
To systematically study the static stability for airplanes, we proposed the system of definitions of six static stability concepts. The details of “static flight speed stability” and “static angle of attack stability” were discussed in the first report of this paper, and “static pitch stability” was studied in the second. In this third report, the static stability concepts of lateral and directional motions are discussed, and their stability criteria are obtained in terms of the aerodynamic characteristics of an airframe. They are “static sideslip stability,” “static roll stability,” and “static yaw stability,” and their fundamental definitions are respectively “dY⁄dβ<0,” “dR⁄dΦ<0,” and “dN⁄dΨ<0,” where R denotes rolling moment. The general criteria are deduced from these definitions. Moreover, they give “CYβ<0” for static sideslip stability, “Clβ⁄CYβ>0” for static roll stability, and “Cnβ>0” for static yaw stability, respectively, as the criteria in normal and low subsonic flights. As a conclusion, a summary of this entire paper is given.
GHTA (Gas Hollow Tungsten Arc) welding experiments were conducted under a simulated space environment that consists of vacuum and microgravity. A vacuum chamber for the experiments was placed in the aircraft cabin. Aluminum and titanium plates were employed as work pieces. The experiments prove that the arc discharge phenomena, the welding phenomena, and the microstructure of welded metal receive almost no effect from gravity. Moreover, it is also demonstrated that the weld metal of the plate welded under the simulated space environment hangs down less than that of the plate welded on the ground.