Band 39, Heft 2

The effect of altering the orientation angle of trenched film hole on film cooling effectiveness (ƞ) is analyzed through numerical investigation. The film holes were oriented at 6 different angles viz., 0°, 30°, 45°, 90°, 135° and 180° with a fixed inclination angle 30° and with a pitch to diameter (P/D) ratio 5.The length and depth of the trench embedding the film hole are fixed at 3D and 0.75D respectively. Three different blowing ratios M=0.5, 1 and 1.5 were studied. Results from the study reveal that changing the orientation angle of film hole with in trench offered a considerable improvement on ƞ from near film hole region to far downstream location on test surface. For a low blowing ratio of 0.5, the film cooling effectiveness distribution changes with respect to film hole orientation angle and higher ƞ zones are distributed along the corresponding film hole orientation angle. The film hole oriented at 90° with respect to stream wise direction delivered highest ƞ at all three different blowing ratios due to better lateral spreading and uniform coverage of injected coolant. Except near film hole region, 30° and 45° hole orientation produced uniform lateral effectiveness compared to 0° hole orientation.

This investigation aims to present the jet mixing characteristics and thrust variations of the subsonic jet employed with plain triangular tab and semi-circular corrugated tab by numerical simulation. A triangular tab with semi-circular corrugations is used in this regard at the exit plane of a convergent nozzle, to study the behavior of the jet and its structure. The near jet flow field is studied for different Mach numbers of 0.6, 0.8 and 1, and the comparisons were done for the jet employed with plain triangular tab. To validate the numerical results, experimental validation is carried out for 0.6 Mach jet. The thrust and the potential core length of any jet depend mainly on the percentage of blockage ratio. Since the relationship between the thrust and blockage ratio is such that, the blockage ratio increases, the thrust and the potential core length decreases and vice-versa. The blockage ratio is kept 8.27 % for both the corrugated and plain triangular tabs. From the results, it is found that the Potential core length of the free jet is cut down to 66 % by the jet employed with plain triangular tab, whereas it is 64.5 % for the corrugated tab enabled jet. It is also concluded that the corrugated tab enhances the thrust by 4.43 % for the same blockage ratio and increases potential core length by 3.33 % when compared with the plain triangular tab. This increase in thrust is there by an added advantage of this investigation.

Aircraft engine load spectrum is the basis of stress analysis of components, structural design criteria, the durability test of components and complete engine and lifetime analysis. The main features are as follows: 1.it is difficult to obtain; 2. samples are insufficient; 3. parameters amount are huge, complex correlation. As a result, the lifetime prediction for aircraft engine and the new engine development is in urgent need of the simulation method of load spectrum. The mathematical model of aircraft engine, in specific, the use-related multiple parameters load spectrum was established by using the Principal Component Analysis (PCA). By using this model, simulation about the normal, non-normal and measured load spectrum of aero-engine can be made. In this paper, the simulated method of use-related multi-parameter load spectrum based on PCA can simulate the original load spectrum at any time by few sample measured and keep the correlation between parameters unchanged. The result indicated that the cumulative frequency curve of the simulated load spectrum and the cumulative frequency curve of the original load spectrum are in good agreement, which ensures that the fatigue damage caused by the simulation spectrum is consistent with the original spectrum.

The objective of the current study is to investigate the effects of geometric parameters on the fluid flow patterns, diffusion factor, load-split, and off-design performance of the tandem configuration under subsonic, rectilinear, cascade flow. A 2-D computational model using finite volume formulation has been developed and validated against available experimental data. The obtained data is presented in a format involving loss parameter and loading and will help the designer to have sufficient number of combinations within the useful range of application. It is observed that general performance behavior of tandem cascade is similar to a single airfoil data. Load shifting from the front airfoil to the rear airfoil is noted at positive axial overlaps and higher camber ratios. An optimum gap ratio of around 2 is observed for the range of configurations investigated. Contrary to the previous literature, the present study indicates higher loading of the front airfoil for the best overall loading-to-loss ratio of the tandem configuration.

This paper deals with the cyclic-periodic structure with piezoelectric shunt circuits or a parallel piezoelectric network. The objective is to obtain the vibration suppression effect of them on the cylic-periodic structure. The background of the research is about vibration reduction of bladed disks in aero-engines, and the system is simulated by a lumped parameter model. The tuned and mistuned dynamic equations of the system are derived. The method of equivalent blisk model (EBM) is used to identify the lumped parameters of a finite element model, which is related to the experimental model. Then the Modified Modal Assurance Criterion (MMAC) is used to evaluate the vibration suppression ability of shunt circuits and the network. The numerical results show that both of these two systems can control vibration well, even the mechanical and electrical mistuning level are up to 10 %, and there is an optimal resistance which can be used to design the system to get a good performance of vibration suppression. An experiment is performed to validate the analytical results. The experimental results demonstrate that the ability of vibration suppression for both two systems are almost the same, but the piezo-network has an advantage in vibration delocalization of the mistuned system.

A method of variable speed control system for turboprop engine is presented in this paper. Firstly, the steady operation state of turboprop engine is analyzed, and the operating line is figured out in the steady state characteristic diagram, which is the design basis of Engine Thrust Management System (ETMS). Secondly, the reference model sliding mode multivariable control is used to design the control law to follow the speed instructions given by ETMS. Finally, the optimization of the minimum fuel consumption operating curve is realized, and the control system designed is applied to a numerical model of a turboprop engine. The simulation results show that compared with the constant speed control system, the variable speed control system can reduce the specific fuel consumption by 2.37 % on average and 3.1 % in steady state conditions. Furthermore, the method can enable the pilot to manipulate the turboprop aircraft by using only one throttle lever, which can greatly reduce the pilot operation burden.

In the present work numerical study of full coverage film cooling on an adiabatic flat plate is carried out. Cooling performance of three configurations of cylindrical holes is studied with downstream injection, upstream injection and mixed injection. In mixed injection configuration one column of holes inject in downstream direction and the holes in the adjacent column inject in the upstream direction. Numerical simulations are carried out at different velocity ratios and circumferentially averaged value of adiabatic film cooling effectiveness is estimated. Simulation results indicate that the mixed injection configuration has better and more uniform cooling, throughout the perforated plate, than with downstream injection. The difference is greater with increase in the velocity ratio. Configuration with upstream injection gives better cooling than mixed injection at front few rows of cooling holes but it shows poorer performance with downstream injection in the downstream rows of cooling holes. The obtained results from this study can be an invaluable input for highly loaded combustion chambers.

Advances in thermal science force us to develop more efficient systems. The efficiency of widely-used gas turbine engines, is highly dependent on turbine inlet temperature. However, a high turbine inlet temperature yields material deterioration and long term degradation of turbines. To prevent material deterioration, cooling the hot zones of gas turbine engines, particularly turbine components and blades, is a priority. In this way, long term degradation of the turbine is prevented, while the thermal efficiency of the gas turbine engine is boosted. In the current paper, a flow field within a rib roughed blade internal cooling channel is discussed. Within this scope, a computational fluid dynamics analysis is conducted using a Standard k-ω turbulence model. After this, the same case is experimentally investigated. Experimental results obtained from particle image velocimetry measurements are used to validate the results of the computational fluid dynamics analysis. At the end of the study, the flow field is fully mapped with the recirculation and separation zones being clearly pinpointed.

The shock wave boundary layer interaction (SWBLI) induced separation bubble formation (SB) and its control has been investigated numerically in the mixed compression type of intake in the scramjet engine. The external compression has occurred due to the three successive oblique shocks formed from the three successive ramps of the forebody with the semi-wedge angle of 7.6, 7.0, and 9.4 respectively. The intake is designed in such a way that all three shocks converge and impinge on the leading edge of the cowl lip for the operating Mach number of 5.0. The numerical simulation is carried out by solving steady, compressible 2-D RANS equations using transitional SST k-ω turbulence model to capture the influence of SWBLI in the performance of supersonic intake. The formation of SB and its control by establishing the perforated wall in its proximity are investigated for three different cases based on the perforation with respect to SSB. Findings of the numerical simulation have concluded that the size of the SB decreases to an acceptable level while establishing the perforation in the entire fore-body wall in the isolator region. The feedback loop established between the upstream and downstream of SB could be a possible reason for reducing its size.

Thermodynamic analysis that provides the status and amount of irreversibility is a powerful tool to determine performance parameters of energy conversion systems. However, specifying the location and amount of irreversibility of a system or component does not reveal the full extent of the relationship of the irreversibility between component-component/component-system. Therefore, splitting the exergy destruction into endogenous and exogenous parts provides detailed information on the relationship of the irreversibility between component-component/component-system. The purpose of this paper is to identify the endogenous and exogenous parts of exergy destruction of an experimental turbojet engine (Unmanned Aerial Vehicle Turbojet – UAVT). Test data used for exergy analysis were evaluated at the maximum load of the experimental turbojet engine which consisted of a compressor, a combustion chamber and a turbine. As a result of the endogenous and exogenous parts of the exergy destruction, the majority of the total exergy destruction (15.1416 kW) was calculated as the endogenous part of exergy destruction (13.5473 kW), and the exogenous exergy destruction rate was calculated to be 1.5943 kW in the compressor component of the turbojet engine. In this study, the endogenous and exogenous exergy destruction amounts of other components are also presented in details.

Reynolds-averaged Navier–Stokes equations are used to simulate a practical scramjet inlet geometry using the shock-unsteadiness modified Spalart–Allmaras (SA) turbulence model. The geometry consists of fore-body ramps, expansion corners, and inlet ducts. The focus is to study the impingement of the cowl shock on the opposite wall boundary-layer. The resulting separation bubble can lead to blockage and inlet unstarts. The shock-unsteadiness correction is employed and is found to improve the computational fluid dynamics (CFD) prediction of flow separation in shock/boundary-layer interactions. The shock-unsteadiness parameter is calibrated against available experimental data of canonical flows, and the predicted flow-field is analyzed in detail. A large separation bubble size normalized to the upstream boundary-layer thickness of 4.6 is observed in the interaction region. Across the reattachment region in the interaction region, a peak value of wall pressure is observed. The inlet performance parameters are also calculated. The total pressure losses of 62% are observed across different shock waves, with an additional loss of 15% due to viscous boundary-layer effects.