This research focuses on supersonic combustion using Shock wave application theory, utilizing shock induced combustion with commercially available Ansys ® Fluent. The dominant mode employed with the design of scram jets is non-premixed fuel induction that entails complex geometries such as flame holders, cantilevered injectors etc. A premixed mode paired with shocked induced combustion, on the other hand, can be achieved with simpler geometries than other configurations. In current research work combustion as induced by a normal and oblique shock waves is investigated. Hydrogen gas at stoichiometric ratio and equivalence value was used as fuel. Normal shock induced combustion was investigated for flows accelerated in diverging test section at sea level conditions. Initially a non-reacting flow was studied. Subsequently, a case for reacting flow and supersonic combustion with normal-shock-induced ignition was investigated. For H2 combustion with NOx, a twenty-one (21) species variant was explored. The hydrogen combustion sub-mechanism was taken from Li et al. and comprises 21 elementary chemical steps. NOx sub-mechanism is based on the Glarborg group's research available with Fluent. A reaction zone was observed at Mach Line as well as radical formation zones were observed at two points beyond the Mach line. Hydrogen gas and Oxygen mass fraction was reduced across reaction zone whereas water formation was observed in the chamber. In the second case, the model is chosen from the experiment of Tan et al to model oblique shock induced combustion. A premixed air-hydrogen gas mixture at stoichiometric ratio is incident at Mach 5, hits a dual ramp configuration at varying incident angles. A global reaction mechanism is chosen for hydrogen gas combustion reaction along with FR/ED model for TCI in Ansys ® Fluent. Oblique shocks are created, and close coupling between the reaction zone and the shockwave occurs as a shock induced combustion, where burnt gases at elevated static pressure, density and temperature are observed post shock wave. The flow analysis for reacting flow shows a positive correlation between Mach numbers, flow turning angle and the heat of reaction. Whereas, there is a negative correlation between altitude and heat of reaction. Combustion is also modelled with Eddy Dissipation Concept (EDC) and it predicts a greater pressure ratio jump at shock wave formation. It also predicts more product formation compared to FR/EDM.
The performance of a wind turbine rotor is determined by the site's wind conditions as well as the blades' aerodynamic design. The torque and power generated by the rotor are determined by the blade geometry. There are some differences in the design of small wind turbine blades compared to large blades. Since small wind turbine airfoils have a significantly lower Reynolds number than large ones, large wind turbine airfoils may perform poorly in small - scale applications. Simple flat-plate airfoil blade is simple to fabricate as compared to conventional airfoil and can be used in small scale wind turbines without subduing performance. In this study, Blade Element Momentum Theory (BEM) is applied to design a HAWT blade using flat plate as an airfoil utilizing characteristics of both Trailing edge (increased CL at same angle of attack) and Leading-edge flaps (Increased Stall Angle) for a horizontal axis wind turbine. Wing tip is being included in design ensuring a particularly efficient flow and reducing induced drag. A complete aerodynamic analysis has been performed using Commercial CFD Software for blade designing along with inclusion wing tip and slots. Optimized parameters for wind tip location, cant angle, sweep angle, taper ratio and span. The analysis at different tip speed ratio is carried out to get the complete power curve for determining total power produced by the wind turbine for annual Electrical power generation. A combination of different blades optimized for different tip speed ratios are being analyzed in order to enhance power envelop for three bladed turbine.
This research is an effort to computationally evaluate the flow field properties across a fan rotor stage of an in-service low bypass turbofan engine whose design has been refined and improved. From the comparative analysis of the performance map of the original configuration with the experimental results, it was highlighted that the fan was unable to achieve the design point parameters. The systematic investigation revealed minor measurement flaws in the CAD model of the blade and hub assembly that could be improved by making minute amendments in the geometric features such as the blade incidence angle, lean angle, and hub diameter. After the requisite refinement in the improved design, the computed performance map was generated. The comparison of computed and experimental results manifested a handsome agreement with an error of less than 2%. Variation in different fluid flow parameters such as pressure, velocity and turbulent viscosity was studied. The improved fan was able to achieve the design point parameters after the amendments.
Coning simulates motions encountered during spin and it can be visualized as a cone formed by the aircraft. An aircraft in fully developed spin displays a coning motion, which can be described as aircraft’s longitudinal axis rotating about wind direction axis (velocity vector) such that aircraft’s center of gravity lies on the wind direction axis. Two extremes of coning are yawing motion perpendicular to wind axis and rolling motion parallel to wind axis. Coning motion is a complex aerodynamic phenomenon and very difficult to capture using CFD. An attempt has been made to understand this coning motion for a supersonic aircraft using commercially available CFD software. The results from CFD analysis are compared with available wind tunnel data, because of complex nature of this type of testing the number of facilities having rotary rigs in their wind tunnels is small which further increase the importance of using CFD for such simulations. The CFD is performed on a full-scale aircraft whereas wind tunnel results are of a 1:13 scale model. For wind tunnel testing aircraft model is supported with a rear sting for angles of attack less than 45° and with a dorsal sting for angles of attack more than 45°. Wind tunnel testing was performed for a specific range of rotation rate which is scaled for CFD calculations. In addition to this range CFD calculations were performed for an enhanced range in order to be representative of a realistic aircraft scenario. The simulations were carried out using Shear Stress Transport (SST) k-ω turbulence model at four different angles of attack i.e., 15°,40°,50°,60°. Aerodynamic forces and moments coefficients obtained as a result of CFD simulations show an agreement with wind tunnel data for angles of attack 15° and 40°. While angles of attack 50° and 60° show a difference in lateral forces and moments due to the presence of dorsal sting attachment for angles of attack higher than 45°, lack of model/attachment bias wind tunnel testing and geometric imperfections in wind tunnel model. This analysis can contribute to better understand aircraft spin and spin recovery characteristics utilizing aerodynamic data as inputs for a 6 DOF mathematical model.
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