Boundary Layer Ingesting Propulsion System Research Articles

A data-driven flow loss prediction model for the blade hub region of a boundary layer ingestion fan rotor

PurposeIn a boundary layer ingestion (BLI) propulsion system, the fan operates continuously under distorted inflow conditions, leading to an increment of aerodynamic loss and in turn impacting the potential fuel burn reduction of the aircraft. Usually, in the preliminary design stage of a BLI propulsion system, it is essential to assess the impact of fuselage boundary layer fluids on fan aerodynamic performances under various flight conditions. However, the hub region flow loss is one of the major loss sources in a fan and would greatly influence the fan performances. Moreover, the inflow distortion also results in a complex and highly nonlinear mapping relation between loss and local physical parameters. It will diminish the prediction accuracy of the commonly used low-fidelity computational approaches which often incorporate traditional physics-based loss models, reducing the reliability of these approaches in evaluating fan performances. Meanwhile, the high-fidelity full-annulus unsteady Reynolds-averaged Navier–Stokes (URANS) approach, even though it can give rather accurate loss predictions, is extremely time-consuming. This study aims to develop a fast and accurate hub loss prediction method for a BLI fan under distorted inflow conditions.Design/methodology/approachThis paper develops a data-driven hub loss prediction method for a BLI fan under distorted inflows. To improve the prediction accuracy and applicability, physical understandings of hub flow features are integrated into the modeling process. Then, the key physical parameters related to flow loss are screened by conducting a sensitivity analysis of influencing parameters. Next, a quasi-steady assumption of flow is made to generate a training sample database, reducing the computational time by acquiring one single sample from the highly time-consuming full-annulus URANS approach to a cost-efficient single-blade-passage approach. Finally, a radial basis function neural network is used to establish a surrogate model that correlates the input parameters and the output loss.FindingsThe data-driven hub loss model shows higher prediction accuracy than the traditional physics-based loss models. It can accurately capture the circumferentially and radially nonuniform variation trends of the losses and the associated absolute magnitudes in a BLI fan under different blade load, inlet distortion intensity and rotating speed conditions. Compared with the high-fidelity full-annulus URANS results, the averaged relative prediction errors of the data-driven hub loss model are kept less than 10%.Originality/valueThe originality of this paper lies in developing a new method for predicting flow loss in a BLI fan rotor blade hub region. This method offers higher prediction accuracy than the traditional loss models and lower computational time cost than the full-annulus URANS approach, which could realize fast evaluations of fan aerodynamic performances and provide technical support for designing high-performance BLI fans.

Numerical study on the Re effects on the tip flow structures of transonic fan under inlet distortion

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.

Non-axisymmetric design and flow field analysis of boundary layer ingesting fans

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.

A Data-Driven Tip Flow Loss Prediction Method for a Transonic Fan Under Boundary Layer Ingesting Inflow Distortion

Abstract In a boundary layer ingesting (BLI) propulsion system, the fan blades need to operate continuously under large-scale inflow distortion. The distortion will lead to serious aerodynamic losses in the fan, degrading the fan performance and the overall aerodynamic benefits of the aircraft. Therefore, in the preliminary design of a BLI propulsion system, it is necessary to evaluate the influence of the fuselage boundary layer under different flight conditions on the fan aerodynamic performance. However, a gap exists in the current computational methods for BLI fan performance evaluations. The full-annulus unsteady Reynolds-averaged Navier–Stokes (URANS) simulations can provide reliable predictions but are computationally expensive for design iterations. The low-order computational methods are cost-efficient but rely on the loss models for accurate prediction. The conventional empirical or physics-based loss models show notable limitations under complex distortion-induced off-design working conditions in a BLI fan, especially in the rotor tip region, compromising the reliability of the low-order computational methods. To balance the accuracy and cost of loss prediction, the paper proposes a data-driven tip flow loss prediction framework for a BLI fan. It employs a neural network to build a surrogate model to predict the tip flow loss at complex non-uniform aerodynamic conditions. Physical understandings of the flow features in the BLI fan are integrated into the data-driven modeling process, to further reduce the computational cost and improve the method’s applicability. The data-driven prediction method shows good accuracy in predicting the overall values and radial distributions of fan rotor tip flow loss under various BLI inflow distortion conditions. Not only does it have higher accuracy than the conventional physics-based loss models but also needs much less computational time than the full-annulus time-accurate simulations. The present work has demonstrated a significant potential of data-driven approaches in complex aerodynamic loss modeling and will contribute to future BLI fan design.

Aerodynamic Shape Optimization of an Aircraft Propulsor Air Intake with Boundary Layer Ingestion

The growth of the airline industry has highlighted the need for more environmentally conscious aviation, leading to the conceptualization of more fuel-efficient aircraft. One concept that has received significant attention and has been associated with improved fuel efficiency is the boundary layer ingesting (BLI) propulsion system, which refers to the ingesting of the aircraft wake by the propulsors. Although BLI has theoretically been proven to reduce fuel burn, this can potentially be offset by the reduced efficiency and stability experienced by the propulsor in the presence of distorted inflow. Therefore, engine intakes must be optimized in order to mitigate the effects of BLI on the propulsion system. In this work, the shape optimization of a BLI intake is investigated using a free-form deformation technique in combination with a multi-objective genetic algorithm, in order to minimize pressure losses and distortion at the engine inlet. The optimization is performed on an S-duct intake at a cruise altitude of approximately 37,000 feet and a free stream Mach number of 0.7. An optimization strategy was developed for the task which was able to produce a Pareto optimal set of designs with improved pressure recovery and distortion. The general trend of the optimal designs shows that to reduce distortion the optimizer accelerates the flow to reduce the size of the low total pressure region and increase the dynamic pressure at the engine inlet. In contrast, the pressure recovery was increased by reducing velocity as well as shifting the maximum velocity region to the outlet, which reduces the viscous dissipation losses within the intake. The final result is a fully autonomous optimization strategy resulting in reduced pressure losses and reduced distortion leading to higher efficiency BLI S-duct intake designs.

Numerical study on rotor tip flow loss in a boundary layer ingesting fan under different operating conditions

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).

Effect of three non-axisymmetric stator schemes on compressor performance with distorted inlet

In the boundary layer ingesting propulsion system, the compressor suffers from a non-uniform flow field. The compressor operating with distorted inflow continuously results in the loss of aerodynamic performance and stability margin. In this paper, three non-axisymmetric configurations are described for the stator of a transonic compressor to match the non-uniform flow field. The flow fields with distorted inflow at near stall condition are obtained and analyzed, the effects of the prototype stator and the three non-axisymmetric stators on aerodynamic performance are compared in detail. Results show that the non-axisymmetric stator schemes can effectively improve the stability margin of the transonic compressor and the maximum stability margin is relatively increased by 22.3% in all the three non-axisymmetric stators. The non-axisymmetric stator design is effective on decreasing the aerodynamic losses and improving the performance of the compressor operating with distorted inflow. Overall, the results show that in the design of the non-axisymmetric stator, the adoption of a curved-twisted blade and the increase of cascade solidity have the potential to reduce loss sources caused by distorted inflow.

Numerical Investigation on the Influences of Boundary Layer Ingestion on Tip Leakage Flow Structures and Losses in a Transonic Axial-Flow Fan

Abstract In a boundary layer ingesting (BLI) propulsion system, the fan is continuously exposed to inflow distortions. The distorted inflows lead to nonuniform loss distributions along the radial and circumferential directions. Since the rotor tip suffers from higher intensive distortion, the local loss increment is a major contributor to the BLI fan performance penalty. To explore the effects of distorted inflows on tip leakage flow evolutions and associated mechanisms for increased loss in a BLI fan, three-dimensional full-annulus unsteady simulations are conducted. Results show that about 54% of total additional losses due to distortion are formed in the tip region and more than 80% of tip region entropy generation is related to the tip leakage flow. The intensities of leakage vortex–shock interactions vary at different annulus locations. When the rotor moves into the distorted region, the vortex–shock interaction is weaker than the undistorted locations due to attenuated leakage flow. At the locations where the rotor is moving out from the distorted region, the vortex–shock interaction is notably enhanced because the front part of the blade tip airfoil suffers a higher load, resulting in a rapid vortex core expansion and eventually vortex breakdown. The increase of flow blockage in the front section of blade tip passages at local circumferential positions leads to a corresponding rise of flow loss. The findings in this study highlight the impacts of tip leakage flow on aerodynamic loss of fan working under BLI inflow distortion and provide improved understandings of loss mechanisms in a BLI fan.

Optimization of a transonic axial-flow compressor under inlet total pressure distortion to enhance aerodynamic performance

In a boundary layer ingestion (BLI) propulsion system, the compressor rotor operates under the condition of a severe inflow distortion because the boundary layers on the surface of aircraft have been ingested. In this situation, the low-momentum fluid would inevitably raise the blade load and aggravate the local aerodynamic losses, thus severely deteriorating the compressor overall aerodynamic performance. The paper has applied a sweep optimization design based on surrogate model and genetic algorithm to a transonic axial-flow compressor at the condition of total pressure distorted inflow. The optimization objective is to improve the compressor efficiency. The sweep design has been implemented through controlling the shape of the blade stacking line and the time-saving steady numerical simulations are mainly utilized in the optimization process. Then, the full-annulus unsteady CFD calculations have been conducted for validating the optimized design and further analyzing the flow mechanisms of performance improvements. The computational results show that the inflow distortion has led to non-uniform distributions of the tip leakage loss and the hub flow separation loss along the circumferential direction. It is found that the rotor blade has suffered from the highest tip leakage losses and hub separation losses during the blade leaving the distorted region. After rotor sweep optimization, both tip leakage and hub separation losses have been notably reduced along the full circumferential direction. Moreover, for different pitchwise locations, the optimized rotor sweep design has achieved more aerodynamic performance improvements in the process of the rotor blade leaving the distorted region than those in the processes of the rotor blade moving out of, moving towards and entering the distorted region. Overall, the optimized design performs better than the baseline design in terms of adiabatic efficiency in the entire operating range. In the meanwhile, the total pressure ratio has been maintained. Specifically, an approximately 0.5% increment of efficiency has been achieved by the optimized design at the optimization point. For this paper, the major purpose is to offer useful guidelines for fan/compressor design strategy to achieve high aerodynamic performances under distorted inflow working condition.

Non-uniform stator loss reduction design strategy in a transonic axial-flow compressor stage under inflow distortion

In a boundary layer ingesting (BLI) propulsion system, the fans/compressors continuously operate at a highly distorted inflow condition due to ingesting the boundary layers on the aircraft fuselage. The BLI inlet distortion induces non-uniform whirl and axial velocities at the rotor outlet, leading to a circumferential and radial non-uniform incidence distribution for the downstream stator. One of the main challenges for the compressor performance is a bad match between the rotor outflow and the stator inlet geometry, which would cause flow separations in blade passages. This paper has employed a non-uniform stator loss reduction design strategy to make a redistribution of stator incidence at the compressor off-design working point so as to reduce the potential separation loss and enhance the stage aerodynamic performance. Firstly, three-dimensional full-annulus unsteady numerical simulations are carried out to reveal the specific flow features of the non-uniform incidence distribution at the stator inlet. Then, according to the circumferential and radial non-uniform incidence distribution, a spatial non-uniform stator loss reduction design strategy based on adjusting the local inlet metal angles of the blade airfoil sections is utilized to make it a good match between the rotor outflow and the stator inlet geometry. Finally, the inner flow fields are analyzed to test the effectiveness of the non-uniform loss reduction design strategy. The computational results indicate that the high-incidence regions at the stator inlet are notably reduced and a more uniform incidence distribution has been achieved along both the circumferential and radial directions. Moreover, the non-uniform loss reduction design has successfully suppressed the flow separations at both the hub and tip regions. As a result, the stator has achieved an 18.2% reduction of the aerodynamic loss at the re-designed point and the stage efficiency has been improved by 0.83%. In the meanwhile, the re-designed stator has decreased the aerodynamic losses over the whole compressor operating range, and the re-designed stage has also achieved a higher efficiency than the baseline one while maintaining the total pressure ratio.

Advanced low-noise aircraft configurations and their assessment: past, present, and future

Aircraft noise remains the key inhibitor of the growth of air transportation, but the focus of the noise mitigation strategies has changed. As the propulsor fan pressure ratio is decreased for improved fuel burn and lessened environmental impact, the propulsion system noise can be reduced near or even below the noise level of the airframe. Jet noise has become less of a concern, and during approach and landing, the acoustic signature is predominantly set by the airframe. Novel aircraft concepts and architectures, enabled by distributed, more integrated, and boundary-layer ingesting propulsion systems, pose new aeroacoustic problems which require innovative approaches and call for teaming and collaboration as the technological challenges cut across disciplines. One past example of such a collaborative research effort was the Silent Aircraft Initiative (SAI), aimed at the conceptual design of an aircraft imperceptible to the human ear outside the airport perimeter. The initiative brought together researchers from academia, industry, and government agencies. This chapter gives a brief summary of the Silent Aircraft Initiative, the SAX-40 aircraft design, and the noise reduction technologies which were pursued. A decade past SAI, novel aircraft architectures such as the D8 double bubble aircraft, the outcome of a joint effort between MIT, Aurora Flight Sciences and Pratt and Whitney, are being pursued in the quest of reducing the climate impact of aviation. With regulations continuing to reduce the allowable aviation noise emission levels, both new challenges and new opportunities are emerging. Electric, hybrid, and turbo-electric aircraft concepts are currently being investigated as potential game changers. Independent of the level of electrification, noise will remain a major issue as air transportation is growing and mobility might become a key driver. The chapter discusses a selection of enabling technologies and their implications on acoustics and noise and gives a perspective on future trends and new directions in aeroacoustics required to address the challenges.

Installed Performance Assessment of an Array of Distributed Propulsors Ingesting Boundary Layer Flow

Conventional propulsion systems are typically represented as uninstalled systems to suit the simple separation between airframe and engine in a podded configuration. However, boundary layer ingesting systems are inherently integrated, and require a different perspective for performance analysis. Simulations of boundary layer ingesting propulsions systems must represent the change in inlet flow characteristics, which result from different local flow conditions. In addition, a suitable accounting system is required to split the airframe forces from the propulsion system forces. The research assesses the performance of a conceptual vehicle, which applies a boundary layer ingesting propulsion system—NASA's N3-X blended wing body aircraft—as a case study. The performance of the aircraft's distributed propulsor array is assessed using a performance method, which accounts for installation terms resulting from the boundary layer ingesting nature of the system. A “thrust split” option is considered, which splits the source of thrust between the aircraft's main turbojet engines and the distributed propulsor array. An optimum thrust split (TS) for a specific fuel consumption at design point (DP) is found to occur for a TS value of 94.1%. In comparison, the optimum TS with respect to fuel consumption for the design 7500 nmi mission is found to be 93.6%, leading to a 1.5% fuel saving for the configuration considered.

Method for simulating the performance of a boundary layer ingesting propulsion system at design and off-design

Boundary layer ingestion has emerged as a potential propulsion concept on novel aircraft configurations for the future. As these concepts progress, preliminary design tools are required that enable the simulation of these aircraft and the rapid analysis of multiple configurations. Simulation tools for boundary layer ingesting propulsion systems tend to focus on proving performance benefits at design point. However, the simulation of aircraft configurations that utilise boundary layer ingestion requires a method to simulate the propulsion system at a range of flight conditions other than design point. A tool is therefore required to enable simulations at off-design. This research presents a work flow to simulate a boundary layer ingesting propulsion system at design and off-design. The process is intended as a tool for design space exploration and the rapid analysis of concepts at the conceptualisation phase. Boundary layer calculations have been combined with conventional 1-D gas turbine performance methods to predict performance of a propulsion system at design point. This method is then extended to enable simulations at off-design conditions for a range of flight conditions or propulsion system power settings. The formulation provides a thrust-drag representation that supports conventional aircraft simulation tools. A case study of an aircraft configuration which utilises an array of boundary layer ingesting propulsors is used to demonstrate the process. The performance of individual propulsors in the array is compared at off-design. Simulations found that, although each propulsor was sized for the same propulsive force at design point, off-design performance diverged depending on operating conditions. In addition, the performance of the propulsor array as a whole was predicted as a function of altitude and Mach number. The case study is used to draw general conclusions on the performance characteristics of a boundary layer ingesting propulsor.

Computational Fluid Dynamics-Based Analysis of Boundary Layer Ingesting Propulsion

A computational fluid dynamics-based analysis has been performed to assess propulsive efficiency benefits of a boundary layer ingesting propulsion system. This analysis is based on a full-scale configuration at relevant flight conditions defined in previous vehicle-level system studies and component designs and includes the full propulsor geometry. A method to calculate the net thrust, equivalent power, and propulsive efficiency of boundary layer ingesting propulsion systems is presented and applied to the computational fluid dynamics solutions. It was found that the boundary layer ingesting system results in a propulsive efficiency benefit of 4.2–4.5% over that of a conventional propulsion system for the case of a fan designed primarily for aerodynamic performance. These results compare favorably to those predicted by engine cycle-based system studies and boundary layer ingesting theory and do not account for anticipated additional nacelle drag reduction benefits. It is recognized that future design work will be required to impart aeromechanics capability to the fan, and thus that the subject results provide an estimate of the upper benefit bound for a full-scale boundary layer ingesting propulsor in a relevant environment.

A preliminary method to estimate impacts of inlet flow distortion on boundary layer ingesting propulsion system design point performance

Boundary layer ingestion by fans in propulsion system improves the propulsive efficiency. However, inlet flow distortion will dramatically eliminate these benefits. This paper puts forward a method to deal with inlet flow distortions and examines their impacts on turbofan performance at engine design point. The method models both radial and circumferential distortion and their impacts separately. Firstly, a distorted fan map is calculated by parallel compressor method. Then, the new map is utilised to find the fan exit flow conditions by parallel stream method. Finally, we assume that all the flows mixed well before entering the nozzle without any pressure losses. At all examined fan pressure ratios, boundary layer ingesting improved fuel consumption. However, the benefits reduced by the new method are lower than previous predictions without considering intake distortion. If the fan pressure loss and efficiency drop due to inlet distortion are too high, boundary layer ingestion should not be used with a traditional fan design. Large boundary layer ingestion for future propulsion system should consider new fan blade design.

Computational analysis of the effects of a boundary layer ingesting propulsion system in transonic flow

Continuous requirements for more efficient aircrafts lead to the design and analysis of novel propulsion configurations, with an example being the boundary layer ingestion. The complexity and integration challenges in such aircraft synergistic propulsion system characterize the research in this field, driven by the potential benefits. The aim of this article is to investigate the effects of boundary layer ingestion on the aerodynamics of a transonic wing, together with the quality of the flow ingested by the propulsion system. A two-dimensional computational model of a transonic airfoil with boundary layer ingesting propulsion system is developed in order to assess boundary layer ingestion for a commercial air transport at cruise conditions and highlight the complex integration issues arising from such configuration. A parametric analysis of the effects of flight conditions, nacelle geometry and engine operating point, on lift, pressure recovery, distortion, total pressure and velocity distribution at the intake, comes to enhance understanding of the performance of this configuration. The pressure distribution around the airfoil and the boundary layer growth are both substantially affected by the engine operating condition, which is represented by the mass flow ratio, with a direct impact on pressure recovery and lift. Mach number and angle of attack influences on lift and drag ingested are also investigated. Intake size and position on the airfoil appear to have significant effects on lift and losses ingested. In general, the results of this study include several aspects related to wing aerodynamics and ingested flow quality, which may facilitate design and integration of the boundary layer ingestion propulsion system for future commercial aircraft.