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In this paper, a multi-objective optimization strategy for efficient design of turbomachinery blades using sparse active subspaces is implemented for a turbofan stage design. The proposed strategy utilized sparse polynomial chaos expansion on a limited dataset to generate a function from which the differential and the covariance matrix can be obtained. Active subspace was used to compute the active variables via singular value decomposition and a hybrid polynomial correlated function expansion was used to construct a surrogate model on the active subspace. Coupled with freeform method and multi-objective genetic algorithm, an automated optimization loop was run at a single operating condition. An improvement in stage efficiency and total pressure ratio of 2.97% and 1.15% was achieved for the optimum design compared with the baseline. Additionally, total pressure loss coefficient decreased by 5.88%, exit flow angle by 34.65% and shock strength by 5.32%. The coupled effect of change in stagger angle, forward sweep, forward lean, and chord length reduced the recirculation at the hub, and blockage at the shroud due the tip leakage flow by decreasing the blade loading. The threshold value hyperparameter was found to be the most influential and must be accurately determined. [ABSTRACT FROM AUTHOR]

To reveal the influence of distortion of S-duct on the flow field and performance of transonic fan, full annulus unsteady numerical simulation was carried out under S-duct/fan integrated condition. This study focuses on the coupling flow of S-duct/fan integrated condition under peak efficiency (PE) condition and near stall (NS) condition. Results show that, compared with the condition that S-duct or fan is investigated alone, the distortion of S-duct is suppressed under S-duct/fan integrated condition. The curved structure of S-duct is the important reason for the flow migration from undistorted region to distortion region. Higher inlet relative Mach number and more work input of the rotor are observed in the counter-swirl region. The flow separation in S-duct is weakened under NS condition compared with that of PE condition. Counter-swirl region of blade tip has larger region of blockage, and rotating stall inception is most likely to occur in this region. [ABSTRACT FROM AUTHOR]

As a key component of boundary layer ingesting (BLI) propulsion system, the BLI fan operates permanently at an inflow distortion condition and the distortion-induced aerodynamic loss in the fan would seriously in turn discount the aerodynamic benefit achievements of BLI propulsion system. Moreover, as the fan rotor rotates along the annulus, the blade tip region will encounter the greatest distortion intensity and operating condition variation, thus contributing to major loss source in the BLI fan rotor. To explore the loss mechanisms in rotor blade tip region of a BLI fan, numerical investigations are conducted to study the influences of BLI inflow distortion on the flow loss in the fan rotor blade tip region at different mass flow working conditions. The results indicate that the flow loss in blade tip region increases notably from near choke to near stall condition, accounting for about 38%-51% of total aerodynamic loss in the BLI fan rotor. The loss in the tip region reaches the maximum at 240° circumferential location where the rotor blade is leaving the distortion region along the annulus because high local blade load and maximum relative Mach number are presented. In the meanwhile, as the rotor blade load increases from near choke to near stall condition, the shock structure in the blade passage varies and the interaction between shock and tip leakage flow is also intensified because of enhanced local leakage flow and forward displacement of shock. At near peak efficiency and near stall conditions, the interaction between the leakage vortex and shock in the tip region is notably intensified at 240° circumferential location and the local leakage vortex expands rapidly after the shock, leading to a vortex breakdown. The resultant flow blockage at 240° circumferential location is much higher than those at other annulus locations and the maximum blockage has increased from 18% to 42% from near choke to near stall condition, contributing fundamentally to the additional loss in the BLI fan. At the same time, by further quantifying local aerodynamic loss, it is also found that the leakage flow loss due to interaction between leakage flow and shock is the major source of loss in the BLI fan rotor, and the vortex breakdown at high blade load condition (from peak efficiency to near stall condition) could even lead to an increment of loss with a slop about three times of that at low blade load condition (near choke condition). [ABSTRACT FROM AUTHOR]

To investigate the effects of different passage structures on the aerodynamic performance of the transonic fans, this paper develops a reliable and practical quasi-3D optimization method for the hub based on the experimental data of Stage 67. In the method, the hub profile of Stage 67 can be optimized without changing the geometrical data of the blades. The optimization results show that stream tube diffusion characteristics depend on the hub profile's curvature in the boundary layer near the hub. In the front part of the hub, the end wall with a concave construction can enhance the expansion of the stream tubes near the root of the rotor blade, which helps control the diffusion flow of viscous fluid effectively to decrease the low-energy fluid's energy degradation and radial secondary flow in the boundary layer. In the latter part of the hub, the end wall with a convex construction facilitates the shrinkage of stream tubes to decrease the secondary flow loss and separated flow loss by controlling the separation of the boundary layer efficiently. This construction of the hub profile is beneficial to promote the aerodynamic performance of a transonic fan. [ABSTRACT FROM AUTHOR]

In order to reveal the different effect mechanisms of blade sweep on the aerodynamic performance when a transonic rotor operates alone or in fan stage environment, two series of forward and backward swept rotors were designed and utilized in the first stage of the dual-stage NASA CR-120859 fan. Results show that, the influence of sweep on the single rotor and the whole stage is different, indicating swept designs for rotor alone may not be suitable for the stage operations. The distinct effect of sweep is account for the difference of the flow field characteristic and stall mechanism of the single rotor and the rotor in the stage environment. The single rotor is tip limited and its stall mechanism is shock/tip leakage vortex (TLV) interaction, whereas the fan stage is hub limited and its stall mechanism is the severe corner separation at stage hub region. For the single rotor, forward sweep increases the stall margin (SM) for all sweep schemes, while backward sweep reduces it in general. For the fan stage, however, backward swept rotor significantly increases stall margin and the stall mechanism is changed to shock/TLV interaction. On the contrary, forward sweep reduces stall margin in general. The flow mechanism is that forward sweep reduces blade loading at tip region near leading edge (LE) and causes the shock to move downstream. Both the variations improve flow field at tip region, while backward sweep exerts an opposite effect. At hub region, backward sweep reduces radial flow tendency by varying radial pressure gradient, causing reduction of corner separation at rotor hub, while forward sweep enhances corner separation. Moreover, with increasing of swept height and swept angle, the chock mass flow, peak efficiency and total pressure ratio of forward sweep are reduced in general, while an opposite effect can be found for backward sweep. [ABSTRACT FROM AUTHOR]

This paper investigates an optimization for the axisymmetric hub-endwall profile of a transonic fan designed for a civil high bypass ratio turbofan engine. A time-saving integrated design optimization method is proposed based on a new hub-endwall profile modification method in combination with a novel two-stage polynomial response surface. The profile modification method is proposed to change the profile of the baseline hub curve by moving three control points, so as to redesign the hub-endwall. The two-stage polynomial response surface is proposed to obtain an accurate high-order polynomial while requiring fewer sampling points, so as to reduce the computational cost. The results show that the integrated design optimization method is suitable and valuable. After optimization, the corner separation of the studied fan is weakened, and the stall margin of the core fan is also enlarged. Moreover, the aerodynamic performance of the core fan at the design speed is improved. [ABSTRACT FROM AUTHOR]

The performance of fan stage in a small turbofan engines is significantly affected at high-altitude low Reynolds number. In order to examine the effect of low Reynolds number on the fan stage, 3D numerical simulation method was employed to analyse the performance variations and the underlying flow structure in the fan stage. For the sake of decreasing the influence of low Reynolds number, the different bowed stator airfoils were redesigned and the effect of the modified design was evaluated. [ABSTRACT FROM AUTHOR]

Investigations of the phenomena associated with the Fluid-Structure interaction in transonic turbomachines due to the presence of unstable flow behaviors have doubled in significance. Severe restrictions of the experimental methods have developed researchers approach in this field to Numerical methods. Nevertheless, using simple two-dimensional model to investigate the phenomenon of quality is inevitable because of high computational cost of numerical methods in aerodynamic and aeroelastic simulation of full model of turbomachines. In this paper transonic flow in fixed fan cascade and fan cascade with central blade vibration in Forced harmonic pattern is simulated and variations of turbulence characteristic patterns are studied. In order to prevent divergence of the solution and achieve more accurate results, the step by step algorithm is developed. On the other hand, spring methodology with linear torsional springs is used for movement of dynamic grid around the oscillating blade. Mesh quality is assessed by examining maximum Mach number and y+ variation. Comparing the results with the available experimental data indicated a significant difference in the position of the vortices that are detached and reattached. This difference proves the use of turbulence model is more accurate in terms of the wide separation. In this paper, effect of blade geometry, flow separation and central blade oscillation on flow pattern and turbulence characteristics of transonic flow have been investigated. Obtained results explain the effect of the mentioned parameters on the turbulence kinetic energy and dissipation frequency. [ABSTRACT FROM AUTHOR]

In modern high bypass turbofan engines, the fan is a considerable component to the performance and efficiency of the entire engine. Fan efficiency directly affects specific fuel consumption of the engine. In the paper, the three dimensional (3-D) flow field through the transonic axial flow fan and the intake of the CF6-50 high bypass ratio turbofan engine is investigated at the takeoff hot day conditions. The fan under study has a large diameter, about 2.18 m. The fan has a number of 38 blades. Each blade is tapered, highly twisted, and has a long span. The real dimensions of both the fan and its intake are used in the current simulation. Furthermore, steady, compressible, turbulent, and viscous flow in both the intake and fan is solved. The computational domain is a periodic sector through both the fan and the intake zones. Results obtained from the simulation show shock waves formed in the passage at the trailing edge of the suction side and propagate towards the opposite blade pressure side. [ABSTRACT FROM AUTHOR]

In the present study, the off-design performance of an axial transonic flow fan is predicted. The fan is 2.18 m in diameter, highly twisted, and has a number of 38 tapered blades with a long span and law aspect ratio. This fan is installed in the high bypass ratio CF6-50 turbofan engine. The three dimensional (3-D) flow field is solved in the fan and its intake at different flight conditions. A periodic sector of a central angle of (360/38) is generated to represent the required computational domain. Results show that, at the fan blade inlet, the average relative Mach number reaches unity at 54% and 40% of the blade inlet span for takeoff and cruise conditions, respectively. At the blade tip the relative Mach number reaches 1.38 at takeoff and 1.5 at cruise. Furthermore, the fan map is plotted over a range of speed lines varying from 60% to 110% of the design value. In addition, the surge line and the operating line at cruise conditions are also plotted on the fan map. The efficiency curves for each speed are plotted as well for the same relative speeds. Moreover, computations revealed that at the design speed, approximately 10.5% stall margin remained beyond design pressure ratio. [ABSTRACT FROM AUTHOR]

The noise of aerodynamics nature from modern transonic fan is examined from its sources with the perspective of noise reduction through aero-acoustics design using advanced Computational Fluid Dynamics (CFD) tools. In particular the problems associated with the forward propagating noise in the front is addressed. It is identified that the shock wave spillage from the leading edge near the fan tip is the main source of the tone noise. Two different approaches have been studied to reduce the forward arc tone noise and two state-of-art transonic fans are designed using the strategies developed. The following rig tests show that while the fans exhibit other noise problems, the primary goals of noise reduction have been achieved through both fans and the novel noise reduction concept vindicated. [ABSTRACT FROM AUTHOR]