The pyrolysis of Methylcyclohexane (MCH) is investigated for the purpose of a regenerative cooling system of hypersonic propulsion by using Endothermic Fuel (EF). The experimental apparatus has heating tube made from INCONEL alloy where MCH would pass through and be decomposed. To confirm the effects of endothermic reaction on the heat absorption, heat fluxes are measured at 9 cross sections in this heating tube. For each section, temperatures are measured at the inner and the outer radii of heating tube and heat flux can be evaluated by these temperature differences. The experimental results show that the endothermic reaction of MCH can confine the temperature increment of MCH in the heating tube and increase the heat fluxes when its temperature was greater than 900K. The numerical simulation can also indicate that MCH pyrolysis have begun around the same temperature. The chemical heat capacity by MCH pyrolysis can increase its total heat capacity to 1.4 times. These facts indicate the usefulness of the endothermic fuels and the possibilities of the regenerative cooling by them.
The effect of aspect ratio (AR) on the aerodynamic characteristics of rectangular wings is investigated. Reynolds numbers considered are 7.6×104 and 5.2×104. Experimental results are compared with analytical ones obtained from Lamar's method. It is shown that the aspect ratio of wings affects the stall characteristics when AR≤2.0. The aspect ratio also has an effect on the lift slope if AR≤1.5. Furthermore, maximum lift coefficients increase when AR≤1.0. Comparisons show that the variations of aerodynamic characteristics are caused by tip vortices. It is also considered that the position of the center of pressure is varied with the angle of attack because of the effects of the laminar separation bubble and tip vortices on wing surfaces.
This study evaluates the system performances of a 10-W-class miniature ion thruster designed for 50kg small spacecraft. The miniature ion thruster used here, using microwave discharge, was specially designed for low microwave power operation, as low as 1.0W. Thruster performance of this thruster (ion beam current, required microwave power, and required gas flow rate) was measured by the experiments. This experiment included a neutralizer and power and gas needed for its operation. Specifications of sub-components needed for a miniature ion thruster system was estimated based on commercially available or space qualified products. As a result, performance of the miniature ion thruster system was evaluated using those thruster performance and sub-component specifications. One of the results is that the miniature ion thruster system can generate 297μN thrust with 1100s specific impulse and ΔV of 300m/s for 50kg spacecraft by 15.6W total power consumption and 2.7–3.5kg total wet weight of the system.
The microwave discharge ion engine μ10's thrust force was improved by additional propellant inlets to a discharge chamber. However, internal plasma diagnostics was not carried out while ion beam was extracted. In order to understand the effects of the new propellant inlets, we measured excitation temperatures and axial number density distributions of metastable Xe I 5p5(2P03/2)6s[3/2]02 inside of μ10 by a line pair method and laser absorption spectroscopy respectively. Firstly, the measurement of excitation temperatures was operated in two positions of the probe tip: 0cm and 5cm from a screen grid. This measurement confirmed that the temperatures marked between 0.42 and 0.68eV. Secondly, the number density distribution measurements were realized by a novel laser absorption spectroscopy utilizing optical fibers. As a result, 1017m-3 order of metastable neutral particles were measured by coupling with the excitation temperatures. Consequently, this paper will reveal that the propellant injection from a waveguide inlet increased the electron number density in the waveguide, which disturbed a propagation of microwave to the discharge chamber. It will also reveal that the propellant injection from the discharge chamber was effective to suppress the plasma production in the waveguide, which resulted in the increase of the thrust.
The flow field in a three dimensional biplane model is investigated by using a focusing schlieren method which can visualize a density gradient of arbitrary cross sections along a light path. Experiments were conducted in the 60mm×60mm indraft supersonic wind tunnel at the Institute of Fluid Science, Tohoku University at M∞ = 1.7. Two test models were used in this experiment: the two-dimensional and three-dimensional models. First of all, the performance of the focusing schlieren system was evaluated by using the two dimensional model. The flow visualization in the three dimensional model is conducted by using the conventional schlieren method and more detail flow visualization by using the focusing schlieren method. Additionally, surface pressure distribution is measured by using the pressure-sensitive paint (PSP). From combined measurements, the flow field in the three dimensional model was discussed. It was found that shock wave structures in the test model vary in the span direction. Additionally, the cause of high pressure regions on the wing surface was understood by comparing with the PSP result.