Volume 38, Issue 1

This present work is aimed at providing detailed understanding of the flow mechanisms in a highly loaded centrifugal compressor with different diffusers. Performance comparison between compressor stages with pipe diffuser and wedge diffuser was conducted by a validated flow solver. Stage with pipe diffuser achieved a better performance above 80 % rotating speed but a worse performance at lower rotating speeds near surge. Four operating points including the design point were analyzed in detail. The inherent diffuser leading edge of pipe diffuser could create a better operating condition for the downstream diffusion, which reduced the possibility of flow separation in discrete passages at design rotating speed. At 60 % rotating speed operating point, there was a large negative incidence angle. The sharp leading edge of pipe diffuser could largely accommodate this negative incidence as comparison of the round leading edge of wedge diffuser. As a result, a better performance was achieved in the pipe diffuser. At 60 % rotating speed near surge, performance of the pipe diffuser dropped below wedge diffuser. Total pressure loss of pipe diffuser exceeded that of the wedge diffuser due to the larger friction loss near wall at throat and ineffective static pressure recovery.

The certification philosophy plays an important role in ensuring the airworthiness qualification of a small gas turbine engine designed as a starter unit. The small and compactness of the engine, high rotational speed of the rotors, requirements of torque and starting capability at sea level to high altitude airfields and consecutive starts within stipulated time impose a great challenge from airworthiness point of view. This paper presents the methodology adopted and various stages of qualification, standards followed and results based on which the starter unit has been qualified for fitment on the designated aircraft.

This paper deals with the creep characteristics of the aircraft turbine disc material of nickel-base superalloy GH4169 under high temperature. From the perspective of continuum damage mechanics, a new creep life prediction model is proposed to predict the creep life of metallic materials under both uniaxial and multiaxial stress states. The creep test data of GH4169 under different loading conditions are used to demonstrate the proposed model. Moreover, from the perspective of numerical simulation, the test data with analysis results obtained by using the finite element analysis based on Graham creep model is carried out for comparison. The results show that numerical analysis results are in good agreement with experimental data. By incorporating the numerical analysis and continuum damage mechanics, it provides an effective way to accurately describe the creep damage process of GH4169.

For rotating machinery, during its running, direct measuring for vibration of rotating components is very difficult. To overcome this defect, we presented a scheme to obtain dynamical responses of a rotor system by synchronizing with a derived virtual system. For a rotor system, a derived system is established, of which some parts are designed to be incorporated with the measurable part of the original rotor system. Then it is proved theoretically that the two systems can reach synchronization, which implies the two systems’ responses are identical. We hope that the derived virtual system can exhibit the dynamical response of the original system, especially in the case of faults occurring. Thus in simulations, first we add a base disturbance to the rotor system. The simulation results reveal that the derived system can reach synchronization with the disturbed rotor quiet well and thus show the disturbed response. Then we assume that there occurs bolt looseness in the rotor support, the simulation shows that the derived system can maintain synchronization with the fault rotor. Thus we can obtain the response of the rotor from the derived system, even the angle response difficult to measure in practice.

Effects of swirling secondary flow on the decay characteristics of non-circular under-expanded supersonic primary jets were experimentally studied. Primary jet is issuing from a convergent-divergent nozzle with different exit geometry such as circle, cruciform, triangle and rectangle. In this study, two different methods were used to impart swirl on the secondary flow as follows, (i) introduction of- angular vanes inside the secondary flow duct and (ii) using tangential inlets at the secondary flow duct. This swirling secondary flow influences the boundary condition of the primary jet and results in reduction of jet core length and shock train length, it also enhances the jet decay rate. Maximum shock train length reduction of 70 % is achieved when swirling secondary flow (tangential injection method) is introduced over the circular primary jet.

Pack aluminization of high pressure turbine nozzle guide vane of an aero gas turbine engine has been carried out following a well defined systematic procedure. The process parameters are first optimized on dummy vanes and optimized process is followed for the actual vanes for evaluation and testing. Visual and binocular examination followed by metallurgical evaluation have been carried out to validate the process and to establish the adequacy and correctness of the coating. The coated vanes are then evaluated through engine level tests for performance and durability. The results of engine level tests and inspection post accelerated mission test cycles ensure that the vanes with aluminide coating can withstand severe engine operating cycles without any damage or failure which would otherwise would have happened without the coating. The condition of vanes post endurance test is also an indication of enhanced life of the vanes with coating.

Corner separation limits the increase of the aerodynamic loading in the compressor. Previous numerical studies indicate that the Blended Blade and End Wall (BBEW) technology is useful in delaying, or reducing, or even eliminating the corner separation. This paper presents combined experimental and numerical investigations on a BBEW cascade and its prototype. The experimental results show that the design of Blended Blade and End Wall (BBEW) can improve the performance of the cascade when the incidence angle was positive or at the design point, and the total pressure loss coefficient was reduced by 7 %-8 %. The performance improvement mainly located from 10 %-25 % span heights. Compared with the experiment data, the SST turbulence model shows the best results of the flow field. Based on the numerical results, the details of the flow field and the effect of the Blended Blade and End Wall (BBEW) design on the corner separation are discussed and analyzed.

The physical characteristics of self-sustain shock oscillation and the key factors that affect the shock oscillation in turbine-based combined-cycle(TBCC) inlet were investigated through unsteady numerical simulation method in this paper. The terminal shock oscillation in the throat was caused by the separation bubble appeared in the lower wall of the turbojet flowpath. The angle of the splitter and the offset of the diffuser were the main factors which affect the characteristics of the shock oscillation. When the angle of the splitter increased from 0 deg to 12 deg, the frequency of terminal shock increased from 100Hz to 133Hz; while when it was greater than 18 deg the terminal shock oscillation phenomenon disappeared. Different offset of the turbojet diffuser affect the motion of terminal shock in the inlet. When the y-coordinate value of the control point (y m ) in the turbojet diffuser was smaller than 1.5, the shock oscillated in the inlet; when y m was greater than 2.0, the shock oscillation disappeared. The bleed in the diffuser was able to control the oscillation of the terminal shock.

The design and model simulation of a can combustor has been made for future syngas combustion application in a micro gas turbine. An improved design of the combustor is studied in this work, where a new fuel injection strategy and film cooling are employed. The simulation of the combustor is conducted by a computational model, which consists of three-dimensional, compressible k-ε model for turbulent flows and PPDF (Presumed Probability Density Function) model for combustion process invoking a laminar flamelet assumption generated by detailed chemical kinetics from GRI 3.0. Thermal and prompt NO x mechanisms are adopted to predict the NO formation. The modeling results indicated that the high temperature flames are stabilized in the center of the primary zone by radially injecting the fuel inward. The exit temperatures of the modified can combustor drop and exhibit a more uniform distribution by coupling film cooling, resulting in a low pattern factor. The combustion characteristics were then investigated and the optimization procedures of the fuel compositions and fuel flow rates were developed for future application of methane/syngas fuels in the micro gas turbine.

Based on the aerodynamic matching configuration between a core driven fan stage (CDFS) and high pressure compressor (HPC), matching state point correlation equations have been established in this paper. The radial nonuniformity of the CDFS outlet flow and the aerodynamic loss caused by complex flow structure between these two components are taken full account in the equations, and a high-accuracy prediction program for matching performance has been developed. Based on the experimental performance data of two components, matching performance prediction at given bypass ratio is made and compared with experimental results. The results show that: 1) The outlet temperature of CDFS varies with its state points change, which makes the relative corrected rotation speed of HPC in matching environment change, and the characteristic line of HPC is not based on the same relative corrected rotation speed. 2) The difference exists between the matching performance prediction results and experimental results. One reason is that CDFS performance varies in the component environment and the matching environment due to the different back pressure regulation modes; another reason is that this prediction method doesn’t consider the effect of the CDFS’s outlet flow nonuniformity on the HPC’s performance, but the maximum relative deviation of the flow rate is only 0.74 %.

The part-load performance of gas turbine is of great importance due to the consideration of running time spent at part power, and Variable Geometry Turbine (VGT) is an important way to meet the requirements. Effect of variable geometry guide-vane with cylindrical endwalls on part-load turbine stage performance is quantified by detailed analysis of aerodynamic characteristics. The turbine stage characteristics of corrected mass flow, corrected output power and efficiency are presented to investigate the VGT performance under the condition of four off-design rotating angles. In addition, the flow field, the loss distribution and the leakage flow are analyzed for different rotating angles. Results show that both the corrected mass flow and the corrected output power increase as the vane throat area is increased, the total leakage area and the shroud area are increased with the increase of rotating angle, whereas the hub leakage area is decreased, and the total pressure loss and the entropy change of variable geometry vane are both increased with the decrease of vane throat area.