The dominant parameters of criteria for autoignition in a supersonic combustor with a backward-facing step and perpendicular wall injectors are investigated. The injectant concentration and the residence time within the recirculation region becomes constant value when the momentum ratio exceeds 0.1 at the same spacing of the injectors. The injectant concentration is correlated with the mass flux ratio (ρu )_j /(ρu )_a at different injectant species and total temperature of the main stream, and it is correlated with the ratio of the step height and the momentum thickness H /θ at different step height and boundary layer thickness. The residence time is correlated with the ratio of the step height and the velocity of the main stream H /U_a at any step height, main stream conditions and injectant species. The step height and the location of injectors appropriate to autoignition is predicted by making contour lines of the equivalent ratio and the residence time on L-H plane.

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A low-noise helicopter blade, AT1, was designed with the concept of reducing noise without the drop of rotor performance. In the concept, High-Speed Impulsive (HSI) noise is reduced by applying a thin airfoil in the tip region and a dog-tooth like extension in the leading-edge of the tip region. Blade-Vortex Interaction (BVI) noise is reduced by applying the extension and a strong taper near the tip end. The stall angle of the blade is increased by the effect of the vortex generated from the leading-edge extension. As a result, the drop of rotor performance caused by the thin airfoil and the reduction of rotor rotational speed is recovered. The low-noise characteristics and the performance of AT1 were evaluated by a model rotor test conducted at Deutsch Niederländischer Windkanal (DNW). It is shown that AT1 reduces HSI noise and BVI noise and has good performance in forward flight conditions. However, the improvement of performance in high-lift conditions still remains as a future problem.

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Direct methods utilizing nonlinear programming such as SQP (Sequential Quadratic Programming) are frequently used as a numerical method for optimal control problems. Although they have advantages in terms of the computational robustness and the usefulness for practical problems, it is usually difficult to select appropriate initial solutions. Therefore, GA (Genetic Algorithm) has become popular as a global search method without particular attention in selecting initial solutions. Since original GA solves the unconstrained optimization problem, the penalty function approach is utilized to handle the constraints. However, the estimation of the appropriate penalty parameter is so difficult that good convergence property is seldom obtained. Therefore, this paper proposes a new selection method, which introduces the multiple criteria, i.e., the distance of the genes, the performance index, and the penalty function. Through the application to the simple test problem and to the ascent trajectory optimization problem of a space plane, it is demonstrated that the proposed method can simultaneously and effectively achieve the global search of the performance index as well as the feasibility search, and it can provide an excellent initial guess for the direct method using nonlinear programming.

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Ignition delays of cool flame (τ₁) and of hot flame (τ_t) are experimentally measured for single n-decane, n-dodecane, n-tetradecane and n-hexadecane droplets, which have corresponding volatility to prevailing commercial hydrocarbon fuels, in hot air. Droplet diameter is 0.7 mm. Ambient pressure is 0.3 and 1.0 MPa. Ambient temperature is 550–1000 K. Results show the similarity of the examined fuels in terms of reactivity of the low- and high-temperature reactions. τ₁ and τ_t are longer for the fuels with less volatility. τ₂(=τ_t-τ₁) is similar for the examined fuels. Ignition limits for cool flame and for hot flame are also similar, while they are slightly lower for the fuels with less volatility.

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Recently, two-stage-to-orbit (TSTO) spaceplane is regarded as one of candidates for fully reusable space transportation system. It consists of a hypersonic booster propelled by airbreathing engines and an orbiter by rocket engines. In many concepts of the airbreathing engines, ATREX engine has been developed in ISAS. The aim of this paper is to apply an optimization method to conceptual designs of TSTO spaceplane with the ATREX engines and to obtain necessary vehicle size and its optimal flight trajectory. First, analysis methods are integrated to define optimization problems. Then, the optimal solutions show that it is necessary to reduce each component weight in order to achieve the practicable vehicles. Especially, the boosters need the huge ATREX engines, which requires improving ATREX engine performance. In addition, it is demonstrated that, by using optimal control technique, the booster can fly back to a launch site by little propellant consumption.

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The basic characteristics of the two-dimensional cavitating flow of liquid helium through a horizontal converging-diverging channel near the lambda point are numerically investigated to realize the further development and high performance of new multiphase superfluid cooling systems. First, the governing equations of the cavitating flow of liquid helium based on the unsteady thermal nonequilibrium multifluid model with generalized curvilinear coordinates system are presented, and several flow characteristics are numerically calculated, taking into account the effect of superfluidity. Based on the numerical results, the two-dimensional structure of the cavitating flow of liquid helium though horizontal converging-diverging channel is shown in detail, and it is also found that the generation of superfluid counterflow against normal fluid flow based on the thermo mechanical effect is conspicuous in the large gas phase volume fraction region where the liquid- to gas-phase change actively occurs. Furthermore, it is clarified that the mechanism of the He I to He II phase transition caused by the temperature decrease is due to the deprivation of latent heat for vaporization from the liquid phase.

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This paper reports a flutter investigation of a re-entry space vehicle having an elastic rotational mode caused by its launching rocket. The elastic rotational mode is taken into consideration as an elastic roll mode or an elastic yaw mode. Flutter experiments were conducted in NAL Transonic Wind Tunnel. The DPM (Doublet-Point Method) is used to calculate flutter boundaries. It is shown that the elastic roll mode may lower critical flutter speed, because its existence alters the natural frequency of an anti-symmetric bending mode with which flutter occurs. A coupling between the elastic yaw mode and an anti-symmetric bending mode of a tip-fin wing is also shown to be critical.

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Aerodynamic performance of scramjet engines was measured by using 0.44-m-long models in a M6.7 wind tunnel. Drags and wall pressure distributions were measured to evaluate the total pressure and friction drag in the engine internal flow. The internal drag of the models with various struts was dicussed. The internal and the external drags, the pressure and the friction drags were estimated. Consistency between the force balance measurements and the pressure measurements was examined. The internal drag obtained from the force balance agreed with that based on the wall pressure measurements.

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In the engine wind tunnels, the test cell pressure does not coincide with the nozzle static pressure and the forces (thrust, lift and pitching moment) are influenced by change of the nozzle pressure ratio. In order to separate effects by the external flow from the force measurement, experiments were conducted by using M3.4 and M5.4 subscale wind tunnels. The three component force was measured by a force balance and the effects of the external flow were investigated by controlling of the nozzle pressure ratio. Measurements of the engine wall pressure enabled to itemize the force distribution delivered by the engine internal flow. The results showed that the external drag of engine occupied from 36% to 70% of the total drag. The nozzle pressure ratio varied the drag coefficient by more than 20% in the M3.4 wind tunnel and greatly influenced the lift and the pitching moment. The boundary layer ingestion promoted the engine unstart especially in higher Mach number testing. The inlet wall pressure was lowered and the pressure distribution was smeared by the ingestion of boundary layer. However, it was found that the engine drag was not affected by it under the operation condition far from the engine unstart.

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