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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]

In this study, the stator geometry of a single-stage transonic fan was optimized to improve its aerodynamic performance. Steady and unsteady simulations were performed on the transonic fan stage to maximize the efficiency and stall margin. In a transonic fan stage, the rotor blade generates different shock waves at various operating points. When the operating condition is changed from the design point to the stall point, the passage shock wave moves upstream, leading to a change in the location of flow separation on the suction surface of the rotor blade toward the upstream. This results in a larger rotor wake near the stall and complicates the stator inlet flow after interaction between the rotor and stator blade. In this study, an optimal geometry of the stator blade is proposed to improve the aerodynamic performance and stall margin of the fan stage. The central composite face-centered method as the design of experiment was used to sample the test cases, and the response surface method was applied to find the optimal case. The transient blade row-time transformation model was applied as an unsteady simulation approach. The optimized fan was found to increase the efficiency at the near-stall point and the stall margin by 0.24 % and 0.36 %, respectively. [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]

• Re effect leads to higher frequency and stronger circumferential momentum of compressor tip leakage flow. • Impact of tip leakage flow decreases from leading edge to mid-chord, but increases from mid-chord to trailing edge at low RNI condition. • Flow instability, identified by DMD, shows growth of low-frequency stall modes and weakening of blade-passing modes. • High RNI condition exhibits non-uniform high-pressure region due to shock-TLV (tip leakage vortex) interaction, leading to delayed flow field response to blade passing. • Low RNI condition results in uniform high-pressure region due to enhanced circumferentia momentum of tip leakage flow, leading to secondary leakage formation. With the continuous increase in aircraft flight altitude, low Reynolds number effects in high-altitude environments significantly impact the stability of compressor operation. In this study, dynamic mode decomposition (DMD) of the flow field at various time instances near stall conditions in a transonic compressor was performed to investigate the influence of Reynolds number (Re) effects on critical flow characteristics and instability mechanisms. The results indicate that under different Reynolds number index (RNI) conditions near stall conditions, the frequency of tip leakage flow varies. Under high RNI condition, the influence of tip leakage flow is concentrated mainly from the leading edge to the mid-chord. The stall mode is primarily caused by the interaction of shock with the Tip Leakage Vortex, resulting in the breakdown of the vortex. The development of vortex structures inside the passage leads to a delay in the flow field response to blade passing. Under low RNI condition, the tip clearance leakage flow is stronger, with a relatively reduced influence from the leading edge to the mid-chord and an increased impact from the mid-chord to the trailing edge. The main cause of the stall mode is the circumferential movement of vortex structures, causing secondary leakage over adjacent blade tip clearances. The blockage formed by vortex structures within the passage is weaker, and the flow field response to blade passing remains consistent with the blade passing frequency. However, as stall develops further, the flow field response significantly decreases. [ABSTRACT FROM AUTHOR]

Three-dimensional throughflow models represent a turbomachinery cascade via a force distribution without the need for detailed geometric modelling in the numerical solution, saving consistent computational resources. In this paper, we present the application of a body force method on an axial transonic fan implemented into an in-house tool for axisymmetric throughflow simulations. By a systematic comparison of local and integral quantities with a validated numerical solution, the capabilities and limitations of the model are discussed for different operating regimes. The implementation is first validated at the peak efficiency calibration point, providing a good duplication of blade flow variables and radial profiles. The design total pressure is matched with a 0.6% absolute difference and a slightly higher slope of the characteristic towards the stall. The isentropic efficiency curve is penalised after the choking mass flow rate calibration, presenting an absolute difference close to 2%, although with a consistent off-design trend. In general, the model provides a satisfactory representation of the flow field and the outflow spanwise distributions, with locally larger discrepancies near the endwalls. Finally, the method is applied to simulate the fan and outlet guide vanes installed into an isolated turbofan nacelle. The onset of intake stall at a high angle of attack is compared between the body force and a boundary conditions-based approaches, highlighting the importance of adopting fully coupled solution methods to study fan/airframe interaction problems.

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]

Abstract Flutter is a long standing issue for fan blades of civil aero-engines and becomes of further concern for modern light-weight designs with increasing fan diameters to reach ultra-high bypass ratios. Accurate flutter prediction is therefore of prime consideration in the design process in order to avoid catastrophic blade failure in operation or expensive redesign iterations if spotted in ground or flight tests. The traditional energy method, which is based on the assumption of negligible aeroelastic coupling, has been used to great extend to predict flutter of turbomachinery components including aero-engine fan blades, and is today by far the most widely applied technique. The underlying assumption of fluid–structure decoupling, however, has to be questioned for large fan blades that are characterized by low mass ratios and low stiffness. Implications of the violation of the system decoupling assumption on the prediction capabilities of the energy method are important to understand for the fan designer in order to allow an informed decision on the flutter prediction tool to use. In this work a comprehensive comparative study is presented in which the energy method is contrasted to the predictions of a strongly coupled fluid–structure interaction method for varying values of mass ratio and blade stiffness of a transonic three-dimensional fan rotor. The strength of aeroelastic coupling is evaluated in terms of the aeroelastic frequency shift and its impact on the prediction accuracy of the energy method is investigated. The results show the capability of the energy method to accurately predict flutter for a wide range of mass ratio and stiffness configurations, but its prediction accuracy is reduced for combined low mass ratio and low stiffness blades. Mechanisms governing the aeroelastic frequency shift are explained to allow a better understanding of the effect and a method for its prediction based on results of a decoupled analysis is shown. [ABSTRACT FROM AUTHOR]

Abstract The redesign and optimization of a low-aspect ratio transonic fan is implemented in this study. An advanced 3D aerodynamic optimization design system is adopted, while flow diagnostic methods are employed to discuss the transonic flow in the blade passages. On the basis of maintaining high aerodynamic efficiency, the study seeks to improve the pressure ratio and through-flow capability of the redesigned fan stage. Furthermore, the dynamic principle for the redistribution of passage flow due to geometry change is revealed. In comparison with the prototype, the total pressure ratio of the redesigned fan is increased by 7.54% at the design point, while its mass flow rate and adiabatic efficiency are raised by 6.30% and 1.25%, respectively. Additionally, a wider high-efficiency operation range is also achieved by the optimization. Under stage flow condition, the control of shock wave at the rotor tip and the removal of low momentum fluid in the stator corner are the two keys in improving the aerodynamic performance of the redesigned fan. Moreover, tangential lean of the stator blade has also succeeded in delaying corner flow separations. Further research for these design techniques would give potential to expand the design system for transonic fan/compressor with low-aspect ratios. [ABSTRACT FROM AUTHOR]

Compared to conventional Pylon-mounted engines, a 5-15% reduction in aircraft fuel consumption is possible with the use of Boundary Layer Ingestion (BLI) propulsion system. However, this type of propulsion system uses a fan that has to operate under continuous inlet distortion and over a wide range of altitude within the flight envelope, which forces researchers to consider the Re effects on the BLI distorted fan stall. For typical spike-type stall, tip leakage flow (TLF) contributes crucially to its generation. Thus, this paper aims to investigate the Re effects on the tip unsteady flow structures under the circumferential distorted inflow near the stall point. In this study, a full-annulus 3D steady/unsteady CFD calculation has been employed to calculate the performance of the NASA 67 transonic fan stage with a 60-deg circumferential distortion at different Re. The results show that the Re effects change the location of peak amplitude of blade tip unsteady flow, without altering the frequency properties. For the further study of the Re effects on the tip unsteady flow, time series analyses of detailed flow structure features and tip leakage vortex (TLV) evolution are performed. At low Re, the angle between the TLF and the axis increases from 77 to 81 deg by comparing the shock structure at 0.99 span, and the secondary flow is enhanced with a marked tendency of blade leading edge spillage. The results indicate that TLV has greater size and strength for low Re condition, and the TLV is more pronounced in the distortion sector than it in the clean sector. Further analysis reveals that the change of shock structure affects the periodic variation in rotor differential pressure on the both surfaces, enhancing the circumferential momentum of TLF near the leading edge, which is the major contributor for changing the vortex structure and causing the stall advance in low Re condition. [ABSTRACT FROM AUTHOR]

The inlet distortion problem exists in a boundary layer ingesting propulsion system, and the distortion position is almost stable, which seriously affects the fan performance and stability margin (SM). To effectively improve the distortion resistance of the fan, three non-axisymmetric stator (NAS) schemes are designed according to the parameter distribution law of the distorted flow field in front of the stator under the design conditions in this study, which are based on the solidity matching method of the optimization of the stator inlet geometric angle, chord length, and pitch. The impact of the NAS on the fan performance and the distorted flow field parameters under distorted conditions are numerically studied. The results indicate that NAS can effectively enhance the SM of the fan. Among the three NAS designs, the third one demonstrates a remarkable increase of 74.91% in SM compared to the original fan. The NAS design can promote the fluid re-distribution in the distortion zone, make the aerodynamic parameters more uniform, and thus improve the flow state in the distortion zone. Among the three NAS schemes that are adopted in this research, the NAS design based on the pitch optimization method has the best effect. This design can effectively improve the stator flow field structure, thus significantly improving the SM and the aerodynamic performance of the fan under distortion conditions. [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]

Buzz-saw noise is a main component of the noise measured in the forward-arc of transonic fans during the take-off and climbing of aircraft. Full resolving of this noise component is computationally expensive because it is very sensitive to random blade variations, such as manufacturing tolerance, deformations or defects. To avoid this, artificially-synthesized irregular buzz-saw noise source is widely used for the prediction and further study. However, the irregularities of the synthesized source are hardly linked to the blade geometry and the so-induced flow structures. In this study, it is found that blade stagger, axial and azimuth variations maintain a quasi-linear relation with the so-induced static pressure perturbations. Based on this quasi-linear relation, the disturbances induced by blade variations are represented by a set of acoustic modal perturbations generated by several pairs of "basic" rotors. By superposing such modal perturbations, the buzz-saw noise source and the static pressure field generated by a fan with multiple arbitrary blade variations can be achieved instantly. This method is programmed to develop a code package named the Mode-based Fast Reconstruction (MFR). It is validated from simple to complex application scenarios based NASA rotor 67. An accuracy level of 0.9 dB is achieved for the multiple pure tones generated by rotors with multiple blade variations that satisfy the quasi-linear assumptions. This enables high-accuracy fast reconstruction of buzz-saw noise source and pressure field in more real transonic fans. Additionally, the impact of blade variations on buzz-saw noise is much stronger than the influence on rotor aerodynamic performances. So, the effects of small blade variations should be considered during the design and optimization of transonic fan rotors, even though the aerodynamic performances may not benefit much from this. [ABSTRACT FROM AUTHOR]

This paper addresses a comparison of numerical stall simulations with experimental data at 60% (subsonic) and 95% (supersonic) of the design speed in a modern transonic fan rig. The unsteady static pressures were obtained with high frequency Kulite transducers mounted on the casing upstream and downstream of the fan. The casing pressure varia-tion was clearly visible in the measurements when a stall cell passed below the trans-ducers. Numerical stall simulations were conducted using an implicit, time-accurate, 3D compressible Reynolds-averaged Navier-Stokes (RANS) solver. The comparisons between the experiment and simulation mainly cover peiformance curves and time-domain pres-sure traces of Kulites during rotating stall. At two different fan speeds, the stall charac-teristics such as the number and rotating speed of the stall cells were well-matched to the experimental values. The mass flow rate and the loading parameter under the fully-developed rotating stall also showed good agreement with the experiment. In both the numerical and experimental results, a large stall cell was eventually formed after stall inception regardless of the fan speed. Based on the validation, the detailed flow has been evaluated to understand rotating stall in a transonic fan. In addition, it was found that the mass flow measurement using casing static pressure might be wrong during transient flow if the Kulites were mounted too close to the fan blade. [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]

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]

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]

The aerodynamics of transonic fans is discussed with emphasis on the use of threedimensional design techniques, such as blade sweep and lean, to improve their performance. In order to study the interaction of these 3D features with the shock pattern a series of five different designs is produced and analysed by CFD. It is found that the 3D features have remarkably little effect on the shock pattern near the tip where the shock must remain perpendicular to the casing. Lower down the blade significant shock sweep, and hence reduced shock loss, can be induced by 3D design but this is usually at the expense of reduced stall margin and increased loss elsewhere along the blade span. Overall, very little change in efficiency is produced by blade sweep or lean. However, forwards lean of the rotor does produce a small increase in mass flow. Radial migration of the boundary fluid on the suction surface behind the shock is shown to play a large part in the aerodynamics near the blade tip.