فهرست مطالب elyas lekzian

In this paper, a turbofan engine with a double bypass duct (DBP) was studied at design point and off-design conditions. This is a separate exhaust, two-spool turbofan engine with bleed, turbine cooling, power extraction, and convergent exhaust nozzles. The bypass ratio of the main and secondary ducts is equal. Also, a typical turbofan engine with one bypass duct is considered the benchmark case (simple engine) and its bypass ratio is 2. Simulation results show that the DBP engine produces 5.4% thrust more than the simple engine at  and an altitude of 9296.4 m (design point). DBP engine thrust is more than the simple engine at SL altitude and 9296.4 m at off-design conditions in all flight Mach numbers ranging from 0 to 0.8. Moreover, the double bypass duct engine SFC is lower than the simple engine at the aforementioned off-design conditions. Thrust per mass flow rate (F/m ̇) was also studied. By increasing F/m ̇, specific fuel consumption is decreased for both engine types. Another interesting finding is that at constant F/m ̇, the DBP engine SFC is lower than the simple engine at all flight altitudes.

In this study, the co-gasification of mazut as the primary fuel and black liquor as a supplementary fuel with the aim of hydrogen-rich syngas is investigated. Oxygen and steam have been chosen as gasification agents. The present research was done using the equilibrium method and Aspen Plus software. The presented model has been validated through an experimental gasification article consisting of combined fuel. Then, by analyzing the performance parameters of gasification, the optimum range of gasification temperature, the ratio of fuel compounds, and the ratio of gasification agent to fuel were determined. Finally, the effect of adding steam as a secondary gasifying agent on the performance parameters and composition of syngas was assessed. The results show that the best ratio of fuel composition ranges from 0.1 and 0.2 and the optimum gasification temperature is 1200-1400 centigrade. Moreover, choosing an appropriate range of oxygen to fuel and steam to fuel causes hydrogen-rich syngas; So that, hydrogen includes more than 50% of syngas.

In the present study, the direct simulation Monte Carlo method is utilized to investigate the effect of obstacles number, location, and also flow injection on the mixing in a channel with 16 μm length and 1 μm height. A mixing length is defined which is the length at which two species are mixed completely. Eight cases with different blockage ratios are considered to study the obstacle effect on the mixing. The blockage ratio shows the reduced flow cross-section due to the addition of obstacles. In All cases, CO2 and N2 gases enter the domain and are separated by a splitter plate that extends up to 1/3 of the channel. The blockage ratio increasing decreases mixing length by up to 10%. Whereas the mass flow rate decreased significantly. Flow injection into the channel is also studied. Four cases are considered: the first case is a simple channel without injection, the second case has cross injection, the third case has inverse injection, and flow is injected vertically through an obstacle in the fourth case. Mixing length is increased by 17% and 5% for cases 2 and 3, respectively. In case 4, the mixing length is decreased by 2% due to the obstacle.

In the current paper, downstream flow field of a propeller at low Reynolds numbers and at static conditions (zero flight speed) is investigated experimentally. This propeller can be utilized in UAVs. Propeller diameter is 56 centimeter and it is investigated at 2550 to 5670 rpm experimentally. Experiment results show that propeller rpm increasing, increases induction velocity. Flow swirl ratio and axial flow coefficient decrease along propeller radius at different propeller rpm. Experimental results of absolute velocity of swirl flow at the propeller airfoil trailing edge downstream is fairly similar to the free vortex flow theory at static condition along the blade radius.  At static condition for r/R<0.8, semi-empirical equations are suggested for variation of flow swirl ratio and axial flow coefficient at downstream of propeller. The propeller is also simulated with numerical simulations. Relative standard deviation of numerical and experimental results for propeller thrust and power are 0.4 and 4.1, respectively. The exponential coefficient (n) which predicts numerical axial flow downstream of propeller for r/R<0.8 has a 7.7 relative standard deviation with experimental results at static condition.

In the present paper a DSMC solver is utilized to study the effects of wall heating/heater plates on performance parameters of microthruster systems. The solver uses local Knudsen number based on the gradient of flow properties to distinct the molecular and continuum region. This solver uses theory of characterisitcs for determination of inlet and outlet boundary conditions. Proper cell dimensions, number of particles per cell, and grid study are performed to guarantee the accuracy of simulations. Three typical micropropulsion systems are studied. All three systems have a microchannel and a converging-diverging micronozzle. First type is cold gas micropropulsion system, second type is a microthruster with wall heated channel, third type is microthruster with heater plates inside. The first type is considered as reference case and two other systems are compared with type1. It is obsereved that heating the walls in microthruster type2 accelerates the flow and increase the specific impulse of the system. In micropropulsion device type3, heater plates increase downstream temperature of convergent-divergent nozzle and also elevate the specific impulse. Due to considerable mass flow rate decrease of system type3, its thrust is decreased whereas mass flow rate of system type2 is not decreased as much as type3 and therefore the thrust of microthruster type2 is more than type1 and type3. Hence the second microprolusion system configuration has higher performance paratmeters in comparison with two other systems. It is also observed that increasing of wall temperature in microthruster type2 decrease the thrust and specific impulse sensitivity to temperature increase.