Turboelectric Distributed Propulsion Modelling Accounting for Fan Boundary Layer Ingestion and Inlet Distortion

2020 ◽  
Author(s):  
Georgios Athanasakos ◽  
Nikolaos Aretakis ◽  
Alexios Alexiou ◽  
Konstantinos Mathioudakis
Keyword(s):  
Boundary Layer ◽  
Parametric Design ◽  
Pressure Ratio ◽  
System Level ◽  
Design Point ◽  
Engine Model ◽  
Inlet Distortion ◽  
Distributed Propulsion ◽  
Inlet Conditions ◽  
The Impact

Abstract A modelling approach of Boundary Layer Ingesting (BLI) propulsion systems is presented. Initially, a distorted compressor model is created utilizing the parallel compressor theory to estimate the impact of inlet distortion on fan performance. Next, a BLI propulsor model is developed considering both distortion effects and reduced inlet momentum drag caused from boundary layer ingestion. Finally, a Turbo-electric Distributed Propulsion (TeDP) model is set up, consisting of the BLI propulsor model, the associated turboshaft engine model and a representation of the relevant electrical system. Each model is validated through comparison with numerical and/or experimental data. A design point calculation is carried out initially to establish propulsor key dimensions for a specified number of propulsors and assuming common inlet conditions. Parametric design point analyses are then carried out to study the influence of propulsors number and location under different inlet conditions, by varying fan design pressure ratio between 1.15 and 1.5. BLI and non BLI configurations are compared at propulsion system level to assess the BLI benefits. The results show that maximum BLI gains of 9.3% in TSFC and 4.7% in propulsive efficiency can be achieved with 16 propulsors and FPR = 1.5, compared to podded propulsors, while further benefits can be achieved by moving the propulsor array backwards in the airframe.

Engine Design Strategy for Boundary Layer Ingesting Propulsion Systems With Multiple Non-Symmetric Engine Inlet Conditions

2013 ◽  
Author(s):  
Jonathan C. Gladin ◽  
Brian K. Kestner ◽  
Jeff S. Schutte ◽  
Dimitri N. Mavris
Keyword(s):  
Boundary Layer ◽  
Design Strategy ◽  
Total Power ◽  
Design Strategies ◽  
Fuel Burn ◽  
Practical Strategy ◽  
Engine Design ◽  
Vehicle Aerodynamics ◽  
Inlet Conditions ◽  
The Impact

Boundary layer ingesting inlets for hybrid wing body aircraft have been investigated at some depth in recent years due to the theoretical potential for fuel burn savings. Such savings derive from the ingestion of a portion of the low momentum wake into the propulsor to reenergize the flow, thus yielding total power savings and reducing required block fuel burn. A potential concern for BLI is that traditional concepts such as “thrust” and “drag” become less clearly defined due to the interaction between the vehicle aerodynamics and the propulsive thrust achieved. One such interaction for the HWB concept is the lateral location of the inlet on the upper surface which determines the effective Reynolds number at the point of ingestion. This is an important factor in determining the amount of power savings achieved by the system, since the boundary layer, displacement, and momentum thicknesses are functions of the local chord length and airfoil shape which are all functions of the lateral location of the engine. This poses a design challenge for engine layouts with more than two engines as at least one or more of the total engines will be operating at a different set of changing inlet conditions throughout the flight envelope. As a result, the engine operating point and propulsive performance will be different between outboard and inboard engines at flight conditions with appreciable boundary layer influence including key flight conditions for engine design: takeoff, top of climb, and cruise. The optimal engine design strategy in terms of performance to address this issue is to design separate engines with similar thrust performance. This strategy has significant challenges such as requiring the manufacturing and certification of two different engines for one vehicle. A more practical strategy is to design a single engine that performs adequately at the different inlet conditions but may not achieve the full benefits of BLI. This paper presents a technique for cycle analysis which can account for the disparity between inlet conditions. This technique was used for two principal purposes: first to determine the effect of the inlet disparity on the performance of the system; second, to analyze the various design strategies that might mitigate the impact of this effect. It is shown that a single engine can be sized when considering both inboard and outboard engines simultaneously. Additionally, it is shown that there is a benefit to ingesting larger mass flows in the inboard engine for the case with large disparity between the engine inlets.

A Methodology for Quantifying Distortion Impacts Using a Modified Parallel Compressor Theory

2018 ◽  
Author(s):  
Manish Pokhrel ◽  
Jonathan Gladin ◽  
Elena Garcia ◽  
Dimitri N. Mavris
Keyword(s):  
Boundary Condition ◽  
Design Space ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Interface Plane ◽  
Design And Optimization ◽  
Fan Performance ◽  
Fuel Burn ◽  
Distributed Propulsion ◽  
The Impact

Efforts to achieve NASA’s N+2 and N+3 fuel burn goals have led to various future aircraft concepts. A commonality in all these concepts is the presence of a high degree of interaction among the various disciplines involved. A tightly integrated propulsion/airframe results in distortion in the flow field around the engine annulus. Although beneficial in terms of propulsive efficiency (due to boundary layer ingestion), the impact of distortion on fan performance and operability remains in question for these concepts. As such, rapid evaluation of the impacts of distortion during the conceptual design phase is necessary to assess various concepts. This is especially important given the expansion of the design space afforded by turbo-electric and hybrid-electric distributed propulsion concepts, in which the gas turbine generator and propulsive devices can be decoupled in space. A simple and rapid methodology to assess operability of compressors is the theory of Parallel Compressors (PC). PC theory views the compressor as two compressors in parallel, one with a uniform high Pt and the other with a uniform low Pt, both operating at the same speed and exiting to a common static pressure. The assumption of two compressors exiting at the common static pressure is not entirely true, especially when the distortion is high. In this paper, the development of a modified parallel compressor model with parametric boundary condition that can capture the impact of non-uniform inflow on fan performance is introduced and validated. Unlike classical PC model, the modified approach introduces a boundary condition dependent on the intensity of distortion (DPCP) at the Aerodynamic Interface Plane (AIP). Additionally, the concept of PC is also extended to Multi-Per Revolution (MPR) distortion. A modeling environment which follows this methodology is created in PROOSIS, an object oriented 0-D cycle code. The model was created using the “compressor” components acting in parallel and a procedure for implementing both design mode and off-design mode solutions was created using the PROOSIS toolset. The example problem was implemented to demonstrate two capabilities — i) the ability of quantifying impacts on thrust and performance of a ducted fan propulsion system, and ii) the ability of predicting loss in stability pressure ratio. The results clearly show the ability of the tool to quantify distortion related losses. The work described in this paper can be integrated to a Multi-Disciplinary Design and Optimization (MDAO) framework along with other disciplines and can be used to evaluate the viability of design space offered by novel aircraft configurations.

scholarly journals Efficient Design Investigation of a Turbofan in Distorted Inlet Conditions

2019 ◽  
Author(s):  
Benjamin Godard ◽  
Edouard De Jaeghere ◽  
Nicolas Gourdain
Keyword(s):  
Body Force ◽  
Latin Hypercube Sampling ◽  
The Body ◽  
Design Domain ◽  
Efficient Design ◽  
Inlet Distortion ◽  
Trailing Edges ◽  
Inlet Conditions ◽  
Fan Blade ◽  
The Impact

Abstract Designing a turbofan to operate in distorted inlet conditions is an issue of growing interest. In such conditions however, fan design can be computationally challenging. Indeed, subject to neither axi-symmetrical nor periodic inlet conditions, computations must be carried out throughout the whole circumferential domain i.e. 360°. Besides, the classical CFD approach implies the use of URANS computations so as to capture the distortion transfer across the fan stage. Eventually, computations are too onerous to be used in design loops. In this context, this paper presents a methodology to effectively assess a fan blade design domain in distorted conditions. This methodology is based on a body-force source term approach formulated in order to accurately recreate deviations, loads and losses across the fan stage. It notably enables to gain two orders in terms of restitution time and thus the possibility to use optimization tools. The design domain of this study is based on variations of profile chord, blade leading and trailing edges angles applied at two different relative heights. A Latin Hypercube Sampling (LHS) associated with a meta-model based on Radial Basis Functions (RBF) enables to assess the impact of geometric variations on performance and operability. Although this study emphasizes that some modeling improvements are still necessary, it also demonstrates the potential of the body-force methodology to conduct fan design when subject to inlet distortion.

Aerodynamic Comparison Between Circumferential and Wing-Embedded Inlet Distortion for an Ultra-High Bypass Ratio Fan Stage

2019 ◽  
Author(s):  
Daniel Giesecke ◽  
Jens Friedrichs
Keyword(s):  
Boundary Layer ◽  
Pressure Ratio ◽  
Aircraft Design ◽  
Low Noise ◽  
Research Centre ◽  
Aircraft Wing ◽  
Local Flow ◽  
Inlet Distortion ◽  
Distortion Coefficient ◽  
Bypass Ratio

Abstract Future aircraft design concepts often show a somewhat wing embedded ultra-high bypass ratio engine. The aircraft concept of the Coordinated Research Centre 880 (CRC880) is a single-aisle configuration with engines partly integrated over the aircraft wing. The aircraft is designed to take off and land on regional airfields with low noise and fuel emissions to address the guidelines set by the ACARE. As a result of the engine installation, the inlet induces a non-axisymmetric boundary layer ingestion into the fan stage. In experimental setups, inlet distortion has often been seen as a 60-degree circumferential inlet stagnation pressure distortion. However, the fan stage inlet flow of the prescribed engine installation of the CRC880 differs to a great extent from a 60-degree sector. In this paper, an aerodynamic comparison between a realistic inflow situation and a similar 60-degree inlet distortion for the same ultra-high bypass ratio fan stage is given. The realistic inflow situation is a result of the flow moving over the aircraft wing suction side and entering the nacelle. As non-axisymmetric inlet geometry remains the same for both cases, therefore, only the total pressure boundary condition at nacelle inlet was changed between both cases. Hence, full annulus simulations are required. Both inlet distortion cases are equivalent by matching average 60-degree distortion coefficient. This study points out that the method, by using averaged 60-degree segment values, excludes specific inflow characteristics. For the same averaged 60-degree distortion coefficient, the local distortion of the embedded case is up to four times larger at rotor tip compared to the segmental approach. For constant mass flow, fan pressure ratio and isentropic efficiency drop by more than five and eight percent respectively. At peak efficiency operating condition, the decrease is even more significant with more than nine percent in stage efficiency. For future embedded aircraft engine configurations, the fan designer has to take into account the non-axisymmetric local flow changes. Hence, the 60-degree segment method is not sufficient when investigating experimental boundary layer ingesting fans and therefore, further method developments are necessary.

Design of a Mixed-Flow Transonic Compressor for Active High-Lift System Using a 3D Inverse Design Methodology

2020 ◽  
Author(s):  
Peng Wang ◽  
Maria Vera-Morales ◽  
Patrick La ◽  
Mehrdad Zangeneh ◽  
Niklas Maroldt ◽  
...  
Keyword(s):  
Design Methodology ◽  
Collaborative Work ◽  
Pressure Ratio ◽  
Inverse Design ◽  
Mixed Flow ◽  
High Lift ◽  
Design Point ◽  
Transonic Compressor ◽  
Reliable Design ◽  
The Impact

Abstract This paper presents the redesign of an electrically driven mixed flow transonic compressor by using a 3D inverse design methodology. The compressor will be used for an active high-lift system application that aims to delay the onset of stall and thus contributing to the reduction of both the aircraft noise footprint and the impact of aviation emission on local air quality. As part of a collaborative work between the Institute of Turbomachinery and Fluid Dynamics of the Leibniz University Hannover and Advanced Design Technology Ltd., an existing optimized compressor stage for this application is redesigned using a 3D inverse method. The new compressor design presents an increase in pressure ratio and total-to-total isentropic efficiency of more than 5.5% and 1% respectively at design point. The higher PR at design point allows the compressor to be run at lower rotational speeds, which decreases the load on the electric motor and the power electronic systems, and hence contributing further to the overall weight reduction of the entire system. The advantage of using an inverse design methodology is shown in this paper as a method that allows a very simple parameterization, reducing significantly the design time and hence allowing the exploration of wider design spaces, with the potential of reaching more innovative and efficient designs. The fast and reliable design and analysis of components represents an important advantage for the enhancement of aircraft electrification, where long design times are often a barrier for the exploration of system configurations.

Validation of a Modified Parallel Compressor Model for Prediction of the Effects of Inlet Swirl on Compressor Performance and Operability

2015 ◽  
Author(s):  
Reginald S. Floyd ◽  
Milton Davis
Keyword(s):  
Pressure Ratio ◽  
Pressure Rise ◽  
Engine Performance ◽  
The United States ◽  
Turbine Engine ◽  
Inlet Distortion ◽  
Inlet Swirl ◽  
Compressor Performance ◽  
Inlet Conditions ◽  
Map Data

Engine inlet distortion complications have plagued the turbine engine development community for decades, and engineers have developed countless methods to identify and combat the harmful effects of inlet distortion. One such type of distortion that has gained much attention in recent years is known as inlet swirl, which results in a significant flow angularity at the face of the engine. This flow angularity can affect the pressure rise and flow capacity of the fan or compressor, and subsequently affect compressor and engine performance. Previous modeling and simulation efforts to predict the effect inlet swirl can have on fan and compressor performance have made great strides, yet still leave a lot to be desired. In particular, a one-dimensional parallel compressor model called DYNTECC (Dynamic Turbine Engine Compressor Code) has been used to analyze the effects of inlet swirl on fan and performance operability of the Honeywell F109 turbofan engine. However, when compared to experimental swirl data gathered at the United States Air Force Academy (USAFA), the model predictions were found to be inaccurate. This paper documents work done to compare the initial predictions generated by DYNTECC to the latest set of experimental swirl data, analyze the potential shortcomings of the initial model, and modify the existing model to more accurately reflect test data. Extensive work was completed to create a methodology that can calibrate the model to existing clean inlet fan map data. In addition, an in depth study of fan/compressor stalling criteria was conducted, and the model was modified to use an alternate stalling criteria that more accurately predicted the point of stall for various swirl inlet conditions. The prediction of the fan stall pressure ratio for all inlet swirl conditions tested is within 2% of the ground test stall point at the same referred fan speed and referred mass flow.

scholarly journals LES and RANS Analysis of the End-Wall Flow in a Linear LPT Cascade With Variable Inlet Conditions: Part II — Loss Generation

2018 ◽  
Author(s):  
Michele Marconcini ◽  
Roberto Pacciani ◽  
Andrea Arnone ◽  
Vittorio Michelassi ◽  
Richard Pichler ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Flow Characteristics ◽  
Turbulence Models ◽  
Companion Paper ◽  
Wall Boundary Layer ◽  
Wall Flow ◽  
Inlet Conditions ◽  
End Wall ◽  
The Impact

In low-pressure-turbines (LPT) at design point around 60–70% of losses are generated in the blade boundary layers far from end-walls, while the remaining 30%–40% is controlled by the interaction of the blade profile with the end-wall boundary layer. Increasing attention is devoted to these flow regions in industrial design processes. Experimental techniques have shed light on the mechanism that controls the growth of the secondary vortices, and scale-resolving CFD have provided a detailed insight into the vorticity generation. Along these lines, this paper discusses the end-wall flow characteristics of the T106 profile with parallel end-walls at realistic LPT conditions, as described in the experimental setup of Duden and Fottner (1997) “Influence of Taper, Reynolds Number and Mach Number on the Secondary Flow Field of a Highly Loaded Turbine Cascade”, P. I. Mech. Eng. A-J. Pow., 211 (4), pp.309–320. The simulations target first the same inlet conditions as documented in the experiments, and determines the impact of the incoming boundary layer thickness by running additional cases with modified incoming boundary layers. Calculations are carried out by both RANS, due to its continuing role as the design verification workhorse, and highly-resolved LES. Part II of the paper focuses on the loss generation associated with the secondary end-wall vortices. Entropy generation and the consequent stagnation pressure losses are analyzed following the aerodynamic investigation carried out in the companion paper. The ability of classical turbulence models generally used in RANS to discern the loss contributions of the different vortical structures is discussed in detail and the attainable degree of accuracy is scrutinized with the help of LES and the available test data. The purpose is to identify the flow features that require further modelling efforts in order to improve RANS/URANS approaches and make them able to support the design of the next generation of LPTs.

An Experimental Study of Loss Sources in a Fan Operating With Continuous Inlet Stagnation Pressure Distortion

2012 ◽  
Author(s):  
E. J. Gunn ◽  
S. E. Tooze ◽  
C. A. Hall ◽  
Y. Colin
Keyword(s):  
Boundary Layer ◽  
Quantitative Assessment ◽  
Three Dimensional ◽  
Leading Edge ◽  
Stagnation Pressure ◽  
Total Efficiency ◽  
Experimental Measurements ◽  
Inlet Distortion ◽  
Point Reduction ◽  
The Impact

The viability of Boundary Layer Ingesting (BLI) engines for future aircraft propulsion is dependent on the ability to design robust, efficient engine fan systems for operation with continuously distorted inlet flow. A key step in this process is to develop an understanding of the specific mechanisms by which an inlet distortion affects the performance of a fan stage. In this paper, detailed full-annulus experimental measurements of the flow field within a low-speed fan stage operating with a continuous 60-degree inlet stagnation pressure distortion are presented. These results are used to describe the three-dimensional fluid mechanics governing the interaction between the fan and the distortion and to make a quantitative assessment of the impact on loss generation within the fan. A 5.3 percentage point reduction in stage total-to-total efficiency is observed as a result of the inlet distortion. The reduction in performance is shown to be dominated by increased loss generation in the rotor due to off-design incidence values at its leading edge, an effect which occurs throughout the annulus despite the localised nature of the inlet distortion. Increased loss generation in the stator row is also observed due to flow separations that are shown to be caused by whirl angle distortion at rotor exit. By addressing these losses, it should be possible to achieve improved efficiency in BLI fan systems.

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