Keiichi Okai

Pressure effects on combustion of methanol and methanol/dodecanol single droplets and droplet pairs in microgravity

Abstract This paper presents the results of an experimental investigation on the combustion of single droplets and two-droplet arrays of pure methanol and methanol/dodecanol mixtures in air under microgravity conditions. The initial droplet diameters, d 0 , were nominally 0.9 mm. The independent experimental variables were the ambient pressure (0.1–9.0 MPa), fuel mixture ratio (methanol/dodecanol: 100/0–15/85), and interdroplet separation distance l ( l / d 0 = 2.3–8.0). For pure methanol, the results show that the droplet lifetime decreases with increasing interdroplet separation distances at low pressures. At higher pressures (3.0 MPa and above) the droplet lifetime was independent of separation distance. The flame extinguished at a finite droplet size only for pure methanol at 0.1 MPa, in qualitative agreement with theoretical predictions. The extinction droplet diameter was nearly independent of the droplet spacing. Methanol/dodecanol–mixture droplets exhibited microexplosion for both single droplets and droplet arrays. The paper presents maps of the disruption regime for both single droplets and droplet pairs. The difference between the disruptive behavior of single droplets and droplet pairs is explained by differences in liquid-phase circulation induced by the gas-phase asymmetry of the droplet pair. The paper also presents results of the dependence of the onset of disruption (in terms of both volume and time) on the pressure and initial fuel mixture ratio.

Effects of dilution by aromatic hydrocarbons on staged ignition behavior of n-decane droplets

Spontaneous ignition of isolated two-component fuel droplets has been experimentally studied. The components were n-decane (ND)/1-methylnaphthalene (MN), or ND/1,2,4-trimethylbenzene (TMB). Both are n-alkane/aromatic mixtures, and therefore are candidates for model fuels of multicomponent commercial fuels refined from crude petroleum. ND showed two-stage ignition behavior, while the other two fuels (aromatic hydrocarbons) did not and were less reactive. A suspended droplet was suddenly brought to a high temperature for ignition. Observation by a Michelson interferometer detected cool-flame appearance as well as hot-flame appearance. The ambient gas was air, and the droplet diameter was 0.7 mm. The experimental conditions that were varied were ambient temperature, Ta, (500–1000K), ambient pressure, Pa, (0.1–2.0 MPa), and the initial mole fraction of ND in the liquid phase, Z (0–1). In the case Z≠1, Pa was fixed to 0.3 MPa. Ignition delays for cool-flame and hot-flame appearance (t1 and ttot) and the difference between them (t2) were measured, and ignition types (no ignition, cool-flame only, singlestage ignition, and two-stage ignition) were classified on a “Z-Ta” map. It is common to ND/MN and ND/TMB that two-stage ignition region is narrowed and finally vanishes as Z decreases, and that t1, t2, and ttot all increase monotonically as Z decreases. However, TMB was affected more than MN because of its higher volatility. (Normal boiling point: ND 447.3 K, MN 517.9 K, TMB 442.5 K) Numerical simulation with a fully transient one-dimensional model was employed to help the interpretation.

Design study on a small pre-cooled turbojet engine for flight experiments

Design aspects of a small pre-cooled turbojet engine for flight experiments are investigated in this study. Pre-cooled turbojet engine is a hypersonic engine suitable for Two Stage To Orbit (TSTO) Space Planes and other hypersonic cruise vehicles. The engine has a capability to accelerate vehicles with a flight speed range between Mach 0 and 6. The payload injection capability of the vehicle appeared to be higher than that with rocket engines, in the former analysis. Some component models for the engine have been tested under simulated conditions. Aerodynamic performances of a variable air intake have been investigated by wind tunnel tests. Anti-frosting technology has been established by using a light weight pre-cooler. System performance of pre-cooling cycle under sea level static condition has been obtained. High temperature structure for a ram combustor and a variable exhaust nozzle has been examined by firing tests. A small pre-cooled turbojet engine for flight experiments is designed based on the component models. The engine will be used for the experiments to obtain thermo-dynamic data of the pre-cooled cycle with the designed flight speed range. Results of thermo-dynamic analyses, CFD analyses and component tests are described in this study. Those results are used to determine the final scale and shape of each component.

Combustion of single droplets and droplet pairs in a vibrating field under microgravity

This paper presents results of an experimental investigation on acoustic effects on combustion of single droplets and droplet pairs in microgravity. The ambient gas was air at atmospheric temperature and pressure, with octane as the fuel. A loudspeaker at the bottom of the chamber produced the acoustic field. Experimental results of single droplets showed that at low frequency and small to moderate acoustic intensities the evaporation rate increases, and the burning rate constant is nearly proportional to the product of frequency, f, and square of displacement, Xa2, fXa2. At higher acoustic intensities, the burning rate constant either remains constant or decreases, and in some cases, flame extinction occurs at a finite droplet diameter. The burning rate constant for a droplet pair is consistently lower than that for a single droplet. At lower frequencies, the burning rate constant reaches a maximum at an intermediate acoustic intensity. At higher frequencies, the burning rate constant increases monotonically with increasing acoustic intensity. The flame size decreases as a result of interactions, as does the critical spacing that indicates a merged flame around the droplet pair versus individual flames surrounding the droplets. The results also show that interactions stabilize the flame, in that droplet pairs burn to completion under conditions in which the flame surrounding a single droplet extinguishes at a finite droplet diameter.

Simplified Analysis on Electromagnetic-Driving Fan for Aircraft Propulsion

An experimental and analytical investigation into a newly proposed electromagneticdriven fan concept is described. Environmental friendliness in aviation is a key feature in the development of future air-breathing engines, and a desirable technology to be introduced is an electromagnetic motion fan system, because it will tremendously reduce the total operating emission as well as reduce weight and increase maintainability increase energy efficiency compared to conventional aero engines. The proposed fan concept has a potential to fulfill the requirements for environmental compatibility and also possesses many other applications such as vector thrust operation. The present paper shows the results of the test on a concept verification model and the result revealed the importance of the controlling parameters to produce efficient power transmission. The accompanying analytical result further explains the motion mechanisms and gives suggestions for the novel design of the fan system.

Spontaneous Ignition of Fuel Droplets in Lean Fuel-Air Mixtures

Experiments have been carried out on the spontaneous ignition of single fuel droplets in lean fuel-air mixtures. The residence time, which is defined as the elapsed time from the introduction of the mixture into the furnace to the insertion of the droplet, is 15 s. The change in the chemical species concentration in fuel-air mixtures before the insertion of the droplet is predicted by the calculation of chemical reaction. In the case of lower ambient temperatures, the methane-air mixture does not react so much and most of methane and oxygen in the ambience survives for the residence time of 15 sec. In this case, the ignition delay time in mixtures is almost the same as that in air, which indicates that the existence of fuel in the ambience has little effect on the spontaneous ignition behavior. On the other hand, the ignition delay time in mixtures becomes larger than that in air for higher ambient temperature. This is mainly due to the decrease in the oxygen concentration in the ambience caused by chemical reaction of fuel-air mixtures during the residence time. However, the quantitative comparison suggests that the increase in the ignition delay time in fuel-air mixtures is not explained by only reduced oxygen concentration.

Performance Analysis of a Fuel Cell Hybrid Aviation Propulsion System

So-called Blended Wing Body (BWB)-type aircraft can accommodate electric fan driven propulsion systems as propulsion devices, providing distributed propulsion. Our earlier report presented a motor concept that is suitable for driving propulsion fans for that aircraft configuration. This report describes an analytical model of a propulsion system power source with size based on the reference vehicle design presented previously by the authors. Take-off and in-flight conditions are presented. The combined fuel-cell and gas-turbine system shows different features in each case.

Investigation of FC/GT Hybrid Core in Electrical Propulsion for Fan Aircraft

Distributed propulsion systems have been investigated by many researchers because the propulsion system can enhance the benefits of unconventional airframe concepts such as blended wing body (BWB) aircraft. The core-separated fan propulsion system mostly considered for use in the distributed propulsion is unique compared to conventional jet engines because the propulsion fans are separated and the driving power is delivered from the core engines placed apart from the propulsion fans. Several power sources might be used, but the most promising one to realize ultralow fuel emission aircraft is electrical. This paper presents the background and the potential for research activities for the core-separated fan propulsion systems. One referenced turboelectric propulsion system was evaluated for validation using a new analytical tool. The authors are emphasizing the configuration of electric fan propulsion system powered by a SOFC-GT core. Benefits and challenges of that core configuration are explained.

MULTI-ROW DISK ARRANGEMENT CONCEPT FOR SPIKE OF AXISYMMETRIC AIR INLET

In this paper are presented the advanced concept of air inlet applying for the air breathing propulsion systems of hypersonic flight vehicle like a space plane. It is designated to be the Multi-Row Disk (MRD) inlet. It was devised to give the better inlet performance (mass capture ratio (MCR) and total pressure recovery (TPR)) over the wide flight conditions up to around Mach 6 by improving its off-design performance. The MRD inlet is a kind of axisymmetric air inlet composed of the center conical envelope spike and the cowl. As shown in Fig. 1, the unique structure is employed on the conical envelope spike that is composed of a tip cone and several round disks arranged so as to shape the conical envelope. The space in between the disks can be changed mechanically and thus the overall spike length of the MRD inlet is adjustable to meet the shock on lip condition at the given flight speed independent of the throat area control. The performance of MCR and TPR are governed respectively by the shock on lip condition to reduce a spillage flow and the throat area control and thus the MRD inlet is easier to improve them independent from each other under the off design conditions. It was made clear from the result of wind tunnel tests that the MRD inlet achieves the same ondesign performance as the conventional inlet with the solid surface conical spike and improves TPR by 10% comparing with the conventional ones in case of flow matching with the turbo fan. It was revealed from the result of wind tunnel tests and numerical simulations that the cavities formed in between the disks gave little effect on a boundary layer growth.

Performance Analysis of Fan-Driving Electric Motors in a Fuel-Cell Hybrid Powered Propulsion System for Aviation

So-called Blended Wing Body (BWB)-type aircraft can accommodate electric fan driven propulsion systems as propulsion devices, providing distributed propulsion. Our earlier report presented a motor concept that is suitable for driving propulsion fans for that aircraft configuration, and provided preliminary sizing of the fan and motor requirement to a given reference vehicle. This paper describes a numerical simulation model of a driving-fan motor with size based on the reference vehicle design the authors presented previously. Numerical simulation is applied first to existing experimental data of the present electric motor concept. Preliminary design and analysis of the motor are provided with dimensions of the 0.71-m-diameter class fan module to be applied to the distributed propulsion system installed in the reference vehicle. Numerical simulation shows that the originally proposed shape of the rotating coil is favorable for the designated purpose, suggesting the basic configuration of the motor driving system.