This study, primarily reports the development of a 3D design procedure for axial flow tandem compressor stages and then the method is used to design a highly loaded tandem stage. In order to investigate the effects of such arrangement, another stage with conventional loading with single blade for both rotor and stator rows is designed with similar specification. In order to ease the comparison of results, chord lengths and hub/shroud geometries are selected with the same dimensions. At the next stage a three dimensional numerical model is developed to predict the characteristic performance of both tandem and conventional stages. The model is validated with the experimental results of NASA-67 stage and the level of the accuracy of the model is presented. Employing the model to simulate the performance of both stages at design and offdesign operating points show that, tandem stage can provide higher pressure ratio with acceptable efficiency. In another word, tandem stage is capable having the same pressure ratio at lower rotational speed. The safe operation domain and loss mechanism in tandem stage are also discussed in this report.

In this study, a 3D design procedure for axial flow tandem stage is developed, and then a highly loaded stage is designed with this method. Designed stage is numerically investigated with a CFD code and the stage characteristic map is reported. In order to investigate the effect of highly loaded design, a conventionally loaded compressor stage is designed with single blade for rotor and stator rows. For a better comparison, Chord lengths and hub/shroud geometries are selected same between highly and conventionally loaded compressors. Characteristic map of the conventional compressor is reported too and performance of highly loaded and conventionally loaded compressors are compared. To validate the CFD method, another compressor stage is presented that its geometry and experimental results are available. Performance of the recent compressor stage is numerically investigated and compared with experimental results.

There is a severe tendency to reduce weight and increase power of gas turbine. Such a requirement is fulfilled by higher pressure ratio of compressor stages. Employing tandem blades in multi-stage axial flow compressors is a promising methodology to control separation on suction sides of blades and simultaneously implement higher turning angle to achieve higher pressure ratio. The present study takes into account the high flow deflection capabilities of the tandem blades consisting of NACA-65 airfoil with fixed percent pitch and axial overlap at various flow incidence angles. In this regard, a two-dimensional cascade model of tandem blades is constructed in a numerical environment. The inlet flow angle is varied in a wide range and overall loss coefficient and deviation angles are computed. Moreover, the flow phenomena between the blades and performance of both forward and afterward blades are investigated. At the end, the aerodynamic flow coefficient of tandem blades are also computed with equivalent single blades to evaluate the performance of such blades in both design and off-design domain of operations. The results show that tandem blades are quite capable of providing higher deflection with lower loss in a wide range of operation and the base profile can be successfully used in design of axial flow compressor. In comparison to equivalent single blades, tandem blades have less dissipation because the momentum exerted on suction side of tandem blades confines the size of separation zone near trailing edges of blades.

Non-dimensional simulation models are often used to interpret experimental data in order to better understand the combustion processes. Recent investigation has presented a thermodynamic non-dimensional model which is used for evaluating the emitted greenhouse gases of a turbofan engine at flight altitude. Combustion chamber flow has been modeled as a non-dimensional flow in provided model. Required data for this model are combustion chamber inlet, turbine inlet and outlet together with exit nozzle pressure and temperature which are calculated by gas turbine non-dimensional modeling software. In this regard mass flow and rotational speed compatibility, conservation of mass and energy equations have been solved. (GSP and KineTechs software (chemical kinetic solving software) have been used for solving the mentioned equations as well as modeling the non-dimensional flow in the combustion chamber, respectively.) This paper investigates the JT9D engine. The results have been validated with experimental data published by International Civil Aviation Organization (ICAO). Noteworthy is that the obtained experimental data have been considered for low flight altitude phases and therefore the presented model takes into account the Greenhouse Gases estimation values for the mentioned phases. After result validation and conformity of experimental data, engine cruise altitude conditions have been modeled. Subsequently, New York – London airway has been studied and emitted Greenhouse gases quantity at cruise altitude has been calculated in one year while verified with published official Statistical data.

No abstract