Coupled aeropropulsive design optimisation of a boundary-layer ingestion propulsor

2018 ◽  
Vol 123 (1259) ◽  
pp. 121-137 ◽  
Author(s):  
Justin S. Gray ◽  
Joaquim R. R. A. Martins
Keyword(s):  
Boundary Layer ◽  
Power Transmission ◽  
Pressure Ratio ◽  
Relative Size ◽  
Thermodynamic Cycle ◽  
Transmission Efficiency ◽  
Navier Stokes ◽  
Design Optimisation ◽  
Fuel Burn ◽  
Design Variables

AbstractAirframe–propulsion integration concepts that use boundary-layer ingestion (BLI) have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA’s STARC-ABL aircraft configuration, which offers the potential for fuel burn reduction by using a turboelectric propulsion system with an aft-mounted electrically driven BLI propulsor. So far, attempts to quantify this potential fuel burn reduction have not considered the full coupling between the aerodynamic and propulsive performance. To address the need for a more careful quantification of the aeropropulsive benefit of the STARC-ABL concept, we run a series of design optimisations based on a fully coupled aeropropulsive model. A 1D thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation to model the aft propulsor at a cruise condition and the effects variation in propulsor design on overall performance. A series of design optimisation studies are performed to minimise the required cruise power, assuming different relative sizes of the BLI propulsor. The design variables consist of the fan pressure ratio, static pressure at the fan face, and 311 variables that control the shape of both the nacelle and the fuselage. The power required by the BLI propulsor is compared with a podded configuration. The results show that the BLI configuration offers 6–9% reduction in required power at cruise, depending on assumptions made about the efficiency of power transmission system between the under-wing engines and the aft propulsor. Additionally, the results indicate that the power transmission efficiency directly affects the relative size of the under-wing engines and the aft propulsor. This design optimisation, based on computational fluid dynamics, is shown to be essential to evaluate current BLI concepts and provides a powerful tool for the design of future concepts.

Shape Optimization of Forward-Curved-Blade Centrifugal Fan with Navier-Stokes Analysis

2004 ◽  
Vol 126 (5) ◽  
pp. 735-742 ◽  
Author(s):  
Kwang-Yong Kim ◽  
Seoung-Jin Seo
Keyword(s):  
Response Surface ◽  
Stokes Equations ◽  
Computing Time ◽  
Relative Size ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Centrifugal Fan ◽  
Navier Stokes Equations ◽  
Design Variables ◽  
Curved Blade

In this paper, the response surface method using a three-dimensional Navier-Stokes analysis to optimize the shape of a forward-curved-blade centrifugal fan is described. For the numerical analysis, Reynolds-averaged Navier-Stokes equations with the standard k-ε turbulence model are discretized with finite volume approximations. The SIMPLEC algorithm is used as a velocity–pressure correction procedure. In order to reduce the huge computing time due to a large number of blades in forward-curved-blade centrifugal fan, the flow inside of the fan is regarded as steady flow by introducing the impeller force models. Four design variables, i.e., location of cutoff, radius of cutoff, expansion angle of scroll, and width of impeller, were selected to optimize the shapes of scroll and blades. Data points for response evaluations were selected by D-optimal design, and a linear programming method was used for the optimization on the response surface. As a main result of the optimization, the efficiency was successfully improved. Effects of the relative size of the inactive zone at the exit of impeller and momentum fluxes of the flow in scroll on efficiency were further discussed. It was found that the optimization process provides a reliable design of this kind of fan with reasonable computing time.

Heat-Exchanged Propulsion Gas Turbines: A Candidate for Future Lower SFC and Reduced-Emission Military and Civil Aeroengines

2009 ◽  
Author(s):  
Colin F. McDonald ◽  
Colin Rodgers
Keyword(s):  
Gas Turbines ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Thermodynamic Cycle ◽  
Simple Cycle ◽  
Specific Power ◽  
Operational Experience ◽  
Fuel Cost ◽  
Turbofan Engines ◽  
Fuel Burn

After seven decades of service the evolution of simple cycle propulsion gas turbines continues with emphasis now being placed on reduced fuel burn, lower emissions, and less noise. With compressor and turbine efficiencies near plateauing, and turbine inlet temperatures paced by materials and blade cooling technologies, improvements in SFC, specific power and weight for conventional engines (including small turboprop, and turboshaft engines and larger turbofans) will likely be incremental compared with the past. With retention of the simple cycle both evolutionary and revolutionary approaches are being taken by the aeroengine industry to improve performance, particularly reduced fuel burn in an era of high fuel cost. In this paper a further step is suggested, that is in concert with meeting performance, economic, and environmental goals of future aeroengines, namely the use of a more complex thermodynamic cycle involving recuperation for turboprop and turboshaft engines, and intercooling together with recuperation for higher pressure ratio turbofan engines. The idea of heat exchanged propulsion gas turbines is not new, but the many concepts identified from studies done periodically over the last 65 years, including the few engines that were static tested and one test flown, didn’t find acceptance in an era of low fuel cost and concerns about recuperator integrity and reliability. With today’s very high fuel cost there is current interest in this type of engine because of its potential for low SFC and reduced emissions. In this paper potential applications are summarized and the features of various heat exchanged aeroengine design concepts together with projected performance are presented. Included is a discussion on the various issues that must be resolved before they enter service. A postulated deployment scenario is suggested with engines initially developed to meet military aviation needs, such as recuperated turboprop and turbofan engines for extended range UAV’s, followed by a recuperated turboshaft engine for a helicopter. Operational experience and demonstrated reliability from these would pave the way for high efficiency ICR turbofan engines for military and civil aircraft service sometime after the year 2020.

Aircraft conceptual design for optimal environmental performance

2012 ◽  
Vol 116 (1175) ◽  
pp. 1-22 ◽  
Author(s):  
R. P. Henderson ◽  
J. R. R. A. Martins ◽  
R. E. Perez
Keyword(s):  
Operating Cost ◽  
Aircraft Design ◽  
Aircraft Emissions ◽  
Design Optimisation ◽  
Commercial Aviation ◽  
Fuel Burn ◽  
Design Variables ◽  
Multidisciplinary Design Optimisation ◽  
Aircraft Conceptual Design ◽  
Geometry Engine

Abstract Consideration of the environmental impact of aircraft has become critical in commercial aviation. The continued growth of air traffic has caused increasing demands to reduce aircraft emissions, imposing new constraints on the design and development of future airplane concepts. In this paper, an aircraft design optimisation framework is used to design aircraft that minimise specific environmental metrics. Multidisciplinary design optimisation is used to optimise aircraft by simultaneously considering airframe, engine and mission. The environmental metrics considered in this investigation are CO2 emissions — which are proportional to fuel burn — and landing-takeoff NOx emissions. The results are compared to those of an aircraft with minimum direct operating cost. The design variables considered in the optimisation problems include aircraft geometry, engine parameters, and cruise settings. An augmented Lagrangian particle swarm optimiser and a genetic algorithm are used to solve the single objective and multi-objective optimisation problems, respectively.

scholarly journals The Influence of Fan Root Flow on the Aerodynamic of a Low-Pressure Compressor Transition Duct

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
A. Duncan Walker ◽  
Ian Mariah ◽  
Dimitra Tsakmakidou ◽  
Hiren Vadhvana ◽  
Chris Hall
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Navier Stokes ◽  
High Pressure Ratio ◽  
Fuel Burn ◽  
Bypass Ratio

Abstract To reduce fuel-burn and emissions, there is a drive toward higher bypass ratio and smaller high-pressure ratio core engines. This makes the design of the ducts connecting compressor spools more challenging as the higher radius change increases aerodynamic loading. This is exacerbated at inlet to the engine core by fan root flow which is characterized by a hub-low-pressure profile and large secondary flow structures. Additionally, shorter, lighter nacelles mean that the intake may not provide a uniform inlet flow when the aircraft is at an angle of attack or subject to cross winds. Such inlet distortion can further degrade the flow entering the engine. A combination of experiments and computational fluid dynamics (CFD) has been used to examine the effects on the aerodynamics of an engine section splitter (ESS) and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A test facility incorporating a 1½ stage axial compressor was used to compare system performance for a flat rotor exit profile to one with a hub deficient flow. Validated Reynolds averaged Navier–Stokes (RANS) CFD was then used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. These changes were seen to have a surprisingly small effect on the flow at the duct exit. However, increased secondary flows were observed which degraded the performance of the ESS and significantly increased loss. Nevertheless, the enhanced mixing delayed separation in the duct suggesting that overall the design was reasonably robust albeit with increased system loss.

Experimental Investigation of Secondary Flows and Length Reduction for a Low-Pressure Compressor Transition Duct

2021 ◽  
Author(s):  
Dimitra Tsakmakidou ◽  
Ian Mariah ◽  
A Duncan Walker ◽  
Chris Hall ◽  
Harry Simpson
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Cumulative Effect ◽  
Poor Quality ◽  
Secondary Flows ◽  
Flow Structures ◽  
Momentum Exchange ◽  
Fuel Burn ◽  
Length Reduction

Abstract The need to reduce fuel-burn and emissions, is pushing turbofan engines towards geared architectures with higher bypass ratios and small ultra-high-pressure ratio cores. However, this increases the radial offset between compressor spools leading to a more challenging design for compressor transition ducts. For the duct connecting the fan to the engine core this is further complicated by poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. This paper presents an experimental evaluation of a new rotor designed to produce these larger flow structures and examines their effect on the performance of an engine sector stators (ESS) and compressor transition duct. Aerodynamic data were collected via five-hole probes, for time-averaged pressures and velocities and phase-locked hot-wire anemometry to capture the rotating secondary flows. The data showed that larger structures promoted mixing through the ESS increasing momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub boundary layer suggesting the duct had moved further from separation. Consequently, an aggressive duct with 12.5% length reduction was designed and tested and measurements confirmed the duct remained fully attached. Total pressure loss was slightly increased over the ESS, but this was offset by reduced loss in the duct due to improved flow quality. Overall, this length reduction represents a significant cumulative effect in reduced fuel-burn and emissions over the life of an engine.

scholarly journals Application of Advanced Computational Codes in the Design of an Experiment for a Supersonic Throughflow Fan Rotor

1988 ◽  
Vol 110 (2) ◽  
pp. 270-279
Author(s):  
J. R. Wood ◽  
J. F. Schmidt ◽  
R. J. Steinke ◽  
R. V. Chima ◽  
W. G. Kunik
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Navier Stokes ◽  
Nasa Lewis Research ◽  
Two Dimensional ◽  
Final Design ◽  
Dimensional Boundary

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The data base for components of this type is practically nonexistent; therefore, in order to furnish the required information for assessment of this type fan, a program has been initiated at the NASA Lewis Research Center to design, build, and test a fan rotor that operates with supersonic axial velocities from inlet to exit. This paper describes the design of the experiment using advanced computational codes to calculate the unique components required. The fan rotor has constant hub and tip radii and was designed for a pressure ratio of 2.7 with a tip speed of 457 m/s. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier–Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. A translating nozzle was designed to produce a uniform flow parallel to the fan up to the design axial Mach number of 2.0. The nozzle was designed with the three-dimensional Euler/interactive boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code, which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

Aerodynamic Analysis of a Boundary-Layer-Ingesting Distortion-Tolerant Fan

2013 ◽  
Author(s):  
Razvan V. Florea ◽  
Dmytro Voytovych ◽  
Gregory Tillman ◽  
Mark Stucky ◽  
Aamir Shabbir ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
High Performance ◽  
Risk Mitigation ◽  
Pressure Ratio ◽  
Wind Tunnel Test ◽  
Fuel Burn ◽  
Global And Local ◽  
The Impact ◽  
Stage Efficiency

The paper describes the aerodynamic CFD analysis that was conducted to address the integration of an embedded-engine (EE) inlet with the fan stage. A highly airframe-integrated, distortion-tolerant propulsion preliminary design study was carried out to quantify fuel burn benefits associated with boundary layer ingestion (BLI) for “N+2” blended wing body (BWB) concepts. The study indicated that low-loss inlets and high-performance, distortion-tolerant turbomachines are key technologies required to achieve a 3–5% BLI fuel burn benefit relative to a baseline high-performance, pylon-mounted, propulsion system. A hierarchical, multi-objective, computational fluid dynamics-based aerodynamic design optimization that combined global and local shaping was carried out to design a high-performance embedded-engine inlet and an associated fan stage. The scaled-down design will be manufactured and tested in NASA’s 8′×6′ Transonic Wind Tunnel. Unsteady calculations were performed for the coupled inlet and fan rotor and inlet, fan rotor and exit guide vanes. The calculations show that the BLI distortion propagates through the fan largely un-attenuated. The impact of distortion on the unsteady blade loading, fan rotor and fan stage efficiency and pressure ratio is analyzed. The fan stage pressure ratio is provided as a time-averaged and full-wheel circumferential-averaged value. Computational analyses were performed to validate the system study and design-phase predictions in terms of fan stage performance and operability. For example, fan stage efficiency losses are less than 0.5–1.5% when compared to a fan stage in clean flow. In addition, these calculations will be used to provide pretest predictions and guidance for risk mitigation for the wind tunnel test.

Unsteady Methods to Investigate Inlet Distortion in a Transonic Tail Cone Thruster Fan Stage With Boundary Layer Ingestion

2020 ◽  
Author(s):  
Ritangshu Giri ◽  
Mark G. Turner ◽  
Mark L. Celestina
Keyword(s):  
Boundary Layer ◽  
Numerical Solutions ◽  
Aerodynamic Performance ◽  
Navier Stokes ◽  
High Fidelity ◽  
Accurate Analysis ◽  
Inlet Distortion ◽  
Fuel Burn ◽  
Time Marching ◽  
Turbomachinery Design

Abstract Boundary Layer Ingestion (BLI) engines have the potential to offer significantly reduced fuel burn, but the fan stage must be designed to run efficiently with a distorted inflow. It must also be able to withstand unsteady aerodynamic loads resulting from a non-uniform flowfield. In a multidisciplinary turbomachinery design cycle involving such a complicated flowfield, high fidelity numerical solutions are required. Two high fidelity unsteady Reynolds Averaged Navier-Stokes (URANS) methods for accurate analysis of a Tail Cone Thruster (TCT) transonic fan stage subjected to inlet distortion have been implemented. They are frequency domain based non-linear harmonic (NLH) and full-annulus complete time domain based time marching methods. This paper demonstrates that the relevant parameters required to accurately compute aerodynamic performance of a fan stage in distorted conditions can be accurately modelled with a few harmonics using the NLH method in a fraction of time compared to the full annulus time marching method. However, the complete aerodynamics of distortion transfer across different blade rows of a fan stage can only be analyzed using the time marching solution. Several physical mechanisms which govern the fan response to an inlet distortion and how different distortion profiles impact the aerodynamic performance of this fan stage are also explained.

Experimental Investigation of Secondary Flows and Length Reduction for a Low-Pressure Compressor Transition Duct

2021 ◽  
Author(s):  
D. Tsakmakidou ◽  
I. Mariah ◽  
A. D. Walker ◽  
C. Hall ◽  
H. Simpson
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Momentum Exchange ◽  
Fuel Burn ◽  
Operational Life ◽  
Length Reduction

Abstract The need to reduce fuel-burn and CO2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phase-locked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO2 over the operational life of an engine.

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