A nylon cable cutting mechanism triggered by a nichrome burn wire is generally used for cube satellite applications for the purpose of holding and releasing onboard appendages owing to its simplicity and low cost. However, the disadvantages of this mechanism are a lower constraint force and also unavoidable system complexity when it is applied to cube satellites with multiple deployable structures. In this study, as the result of an internal development program for cube satellites, we propose a segmented nut-type holding and release mechanism that affords advantages such as high load capability and negligible induced shock. A demonstration model of the mechanism has been fabricated, assembled and functionally tested under qualification temperatures.
The flow around an ONERA-M6 wing, including the effect of wind tunnel wall interference, is computed using CFD analysis with a porous wall model. The computational domain sets porous walls at the top and bottom of the wing section similar to actual wind tunnel experiments. The computational result captures almost the same shock wave shape as the wind tunnel. This could not be computed exactly in previous works that did not include wall effects. The interference of the porous walls reduces the Mach number and attack angle of the flow, and these effects alter the swept angle of front shock wave and the location of rear shock wave. The aerodynamics coefficients are also affected by the wall interference. The lift coefficient becomes smaller due to the reduction in attack angle. The lower Mach number decreases the drag coefficient, while the reduction in attack angle causes additional drag by a mechanism similar to induced drag.
Throttling technology is a key technology for future space transportation system. One of the most important components of the rocket engine for thrust control is the propellant injection system. The pintle injector is one of the most promising candidates due to its simple structure and combustion stability. However, the number of studies on the pintle injector is limited and detailed phenomena of the pintle injector has not been clarified. Therefore, a combustion test using an ethanol/liquid oxygen rocket engine is carried out with a planar pintle injector in a rectangular combustor. Fundamental combustion characteristics such as characteristic exhaust velocity efficiency are acquired at the combustion pressure of 0.45 MPa and O/F=1.25 to 1.80. C* efficiency increases as the equivalence ratio increases up to unity. Flame images during combustion at 0.45 MPa and O/F=1.75 were acquired. Luminous flame was not observed in throughout all of the region except for the vicinity of the faceplate. However, strong luminous flame was observed at the time of ignition. The combustion chamber used in this research can simulate one of the recirculation zones observed in actual combustor with pintle-type injectors, which contributes to the combustion stability. However, the recirculation zone formed at the top of the injector is not simulated appropriately due to the existence of the lower wall of the combustor.
A design procedure for aircraft auto-landing guidance using time delay control (TDC) is proposed, and the designed guidance system is validated via simulations of general landing procedures, including glide slope, flare and touchdown. The simulations adopt a nonlinear dynamics model that includes the landing gear. Additionally, by employing TDC, a stability and control augmentation system, a longitudinal auto-landing guidance law, and a lateral guidance law considering crosswind are proposed. The designed auto-landing guidance laws are evaluated through simulations with aileron fault and crosswind disturbance. The proposed TDC-based auto-landing guidance law shows good performance and is robust when aileron fault and crosswind are included.
In this study, an efficient methodology capable of systematically constructing an aircraft design database is developed and its application is discussed. The database focuses on fighter and attack aircraft because their design is a particular challenge compared with the design of other types of aircraft. For small conventional aircraft, historical information accumulated over many years on countless previous designs is freely available to designers. Such data for fighter aircraft design are, however, difficult to obtain or are insufficient. The database developed can contribute to the design of fighter and attack aircraft on a conceptual level. In addition, decision-making models are developed by means of data mining as well as approximation models based on the developed database to facilitate initial aircraft configuration decisions. The efficiency and simplicity of the proposed method are described in a case study, which also describes a sample process of conceptual design optimization using initial design information extracted from the developed database.
In this paper, optimization of blade twist distribution to maximize the performance of a rotor during hovering flight is described. The Japan Aerospace Exploration Agency has been developing a hybrid method coupled with computational fluid dynamics (CFD) and a prescribed wake model to reduce the computational cost. In this study, as a first step for the application of this method to optimize the blade configuration, the influence of the blade twist distribution variation on the figure of merit is investigated. The Hover Tip Vortex Structure Test (HOTIS) experiment is used as a calculation model, and the tip Mach number is 0.641. It is found that the result of using the hybrid method for the baseline blade configuration corresponds well with the full CFD of the rotor performance and figure of merit. For the influence of the blade twist on the figure of merit, high figures of merit are obtained when the difference of the blade twist at r/R=0.875 from the baseline Δθ1 is around 1–3 deg and the difference of the blade twist at r/R = 1.0 from the baseline Δθ2 is of negative value. Particularly, in the case of Δθ1 = 2.0 and Δθ2 = −3.7 deg, a maximum value of figure of merit of about 1.5% higher than the baseline design is achieved.
The asteroid explorer Hayabusa-2, which is scheduled to be launched in 2014, is going to perform a global mapping mission after it arrives at the target asteroid. Although most of the global mapping sequence will be the same as that of its predecessor Hayabusa, several automation technologies are planned to be tested to reduce the workload of the operators. In particular, the structure from motion (SFM) and simultaneous localization and mapping (SLAM) techniques are expected to significantly contribute to the automation of asteroid shape estimation and visual spacecraft navigation. These frameworks require automatic landmark tracking on the asteroid surface, but no previous work has discussed the method that should be used to track images of the asteroid taken in space, where the absence of scattering light causes dramatic changes in appearance. In this study, we evaluated the performances of SIFT, SURF, BRISK, ORB, Harris-Affine, Hessian-Affine and MSER for images of the asteroid. We found that SIFT is acceptable for use, while SURF, BRISK and ORB can be used with careful parameter tuning. The affine-invariant detectors might contribute to more accurate tracking, but using them is more challenging owing to an extra normalizing process.