On the Optimization of a Geared Fan Intercooled Core Engine Design

2014 ◽  
Vol 137 (4) ◽  
Author(s):  
Konstantinos G. Kyprianidis ◽  
Andrew M. Rolt
Keyword(s):  
Fuel Consumption ◽  
Design Space ◽  
Pressure Ratio ◽  
System Level ◽  
Design Parameters ◽  
Customer Requirements ◽  
Blade Height ◽  
Fuel Burn ◽  
Engine Design ◽  
Last Stage

Reduction of CO2 emissions is strongly linked with the improvement of engine specific fuel consumption (SFC), as well as the reduction of engine nacelle drag and weight. One alternative design approach to improving SFC is to consider a geared fan combined with an increased overall pressure ratio (OPR) intercooled core performance cycle. Thermal benefits from intercooling have been well documented in the literature. Nevertheless, there is little information available in the public domain with respect to design space exploration of such an engine concept when combined with a geared fan. The present work uses a multidisciplinary conceptual design tool to further analyze the option of an intercooled core geared fan aero engine for long haul applications with a 2020 entry into service technology level assumption. The proposed design methodology is capable, with the utilized tool, of exploring the interaction of design criteria and providing critical design insight at engine–aircraft system level. Previous work by the authors focused on understanding the design space for this particular configuration with minimum SFC, engine weight, and mission fuel in mind. This was achieved by means of a parametric analysis, varying several engine design parameters—but only one at a time. The present work attempts to identify “globally” fuel burn optimal values for a set of engine design parameters by varying them all simultaneously. This permits the nonlinear interactions between the parameters to be accounted for. Special attention has been given to the fuel burn impact of the reduced high pressure compressor (HPC) efficiency levels associated with low last stage blade heights. Three fuel optimal designs are considered, based on different assumptions. The results indicate that it is preferable to trade OPR and pressure ratio split exponent, rather than specific thrust, as means of increasing blade height and hence reducing the associated fuel consumption penalties. It is interesting to note that even when considering the effect of HPC last stage blade height on efficiency there is still an equivalently good design at a reduced OPR. This provides evidence that the overall economic optimum could be for a lower OPR cycle. Customer requirements such as take-off distance and time to height play a very important role in determining a fuel optimal engine design. Tougher customer requirements result in bigger and heavier engines that burn more fuel. Higher OPR intercooled engine cycles clearly become more attractive in aircraft applications that require larger engine sizes.

On the Optimisation of a Geared Fan Intercooled Core Engine Design

2014 ◽  
Author(s):  
Konstantinos G. Kyprianidis ◽  
Andrew M. Rolt
Keyword(s):  
Fuel Consumption ◽  
Design Space ◽  
Pressure Ratio ◽  
Specific Fuel Consumption ◽  
Design Parameters ◽  
Customer Requirements ◽  
Blade Height ◽  
Fuel Burn ◽  
Engine Design ◽  
Last Stage

Reduction of CO2 emissions is strongly linked with the improvement of engine specific fuel consumption, as well as the reduction of engine nacelle drag and weight. One alternative design approach to improving specific fuel consumption is to consider a geared fan combined with an increased overall pressure ratio intercooled core performance cycle. Thermal benefits from intercooling have been well documented in the literature. Nevertheless, there is little information available in the public domain with respect to design space exploration of such an engine concept when combined with a geared fan. The present work uses a multidisciplinary conceptual design tool to further analyse the option of an intercooled core geared fan aero engine for long haul applications with a 2020 entry into service technology level assumption. The proposed design methodology is capable, with the utilised tool, of exploring the interaction of design criteria and providing critical design insight at engine-aircraft system level. Previous work by the authors focused on understanding the design space for this particular configuration with minimum specific fuel consumption, engine weight and mission fuel in mind. This was achieved by means of a parametric analysis, varying several engine design parameters — but only one at a time. The present work attempts to identify “globally” fuel burn optimal values for a set of engine design parameters by varying them all simultaneously. This permits the non-linear interactions between the parameters to be accounted. Special attention has been given to the fuel burn impact of the reduced HPC efficiency levels associated with low last stage blade heights. Three fuel optimal designs are considered, based on different assumptions. The results indicate that it is preferable to trade overall pressure ratio and pressure ratio split exponent, rather than specific thrust, as means of increasing blade height and hence reducing the associated fuel consumption penalties. It is interesting to note that even when considering the effect of HPC last stage blade height on efficiency there is still an equivalently good design at a reduced overall pressure ratio. This provides evidence that the overall economic optimum could be for a lower overall pressure ratio cycle. Customer requirements such as take-off distance and time to height play a very important role in determining a fuel optimal engine design. Tougher customer requirements result in bigger and heavier engines that burn more fuel. Higher overall pressure ratio intercooled engine cycles clearly become more attractive in aircraft applications that require larger engine sizes.

First and Second Law Analysis of Intercooled Turbofan Engine

2015 ◽  
Author(s):  
Xin Zhao ◽  
Oskar Thulin ◽  
Tomas Grönstedt
Keyword(s):  
High Pressure ◽  
Exergy Analysis ◽  
Pressure Ratio ◽  
Second Law ◽  
Turbofan Engine ◽  
Blade Height ◽  
Turbofan Engines ◽  
Fuel Burn ◽  
Aero Engine ◽  
Last Stage

Although the benefits of intercooling for aero engine applications have been realized and discussed in many publications, quantitative details are still relatively limited. In order to strengthen the understanding of aero engine intercooling, detailed performance data on optimized intercooled turbofan engines are provided. Analysis is conducted using an exergy breakdown, i.e. quantifying the losses into a common currency by applying a combined use of the first and second law of thermodynamics. Optimal intercooled geared turbofan engines for a long range mission are established with CFD based two-pass cross flow tubular intercooler correlations. By means of a separate variable nozzle, the amount of intercooler coolant air can be optimized to different flight conditions. Exergy analysis is used to assess how irreversibility is varying over the flight mission, allowing for a more clear explanation and interpretation of the benefits. The optimal intercooled geared turbofan engine provides a 4.5% fuel burn benefit over a non-intercooled geared reference engine. The optimum is constrained by the last stage compressor blade height. To further explore the potential of intercooling the constraint limiting the axial compressor last stage blade height is relaxed by introducing an axial radial high pressure compressor. The axial-radial high pressure ratio configuration allows for an ultra-high overall pressure ratio. With an optimal top-of-climb overall pressure ratio of 140, the configuration provides a 5.3% fuel burn benefit over the geared reference engine. The irreversibilities of the intercooler are broken down into its components to analyze the difference between the ultra-high overall pressure ratio axial-radial configuration and the purely axial configuration. An intercooler conceptual design method is used to predict pressure loss heat transfer and weight for the different overall pressure ratios. Exergy analysis combined with results from the intercooler and engine conceptual design are used to support the conclusion that the optimal pressure ratio split exponent stays relatively independent of the overall engine pressure ratio.

Optimization of Turbofan Propulsion Cycle Using a Genetic Algorithm

2010 ◽  
Author(s):  
Adel Ghenaiet
Keyword(s):  
Genetic Algorithm ◽  
Fuel Consumption ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Optimization Procedure ◽  
Specific Fuel Consumption ◽  
High Mass ◽  
Design Parameters ◽  
Inlet Temperature ◽  
Simplifying Assumptions

This paper presents an evolutionary approach as the optimization framework to design for the optimal performance of a high-bypass unmixed turbofan to match with the power requirements of a commercial aircraft. The parametric analysis had the objective to highlight the effects of the principal design parameters on the propulsive performance in terms of specific fuel consumption and specific thrust. The design optimization procedure based on the genetic algorithm PIKAIA coupled to the developed engine performance analyzer (on-design and off-design) aimed at finding the propulsion cycle parameters minimizing the specific fuel consumption, while meeting the required thrusts in cruise and takeoff and the restrictions of temperatures limits, engine size and weight as well as pollutants emissions. This methodology does not use engine components’ maps and operates on simplifying assumptions which are satisfying the conceptual or early design stages. The predefined requirements and design constraints have resulted in an engine with high mass flow rate, bypass ratio and overall pressure ratio and a moderate turbine inlet temperature. In general, the optimized engine is fairly comparable with available engines of equivalent power range.

First and Second Law Analysis of Intercooled Turbofan Engine

2015 ◽  
Vol 138 (2) ◽  
Author(s):  
Xin Zhao ◽  
Oskar Thulin ◽  
Tomas Grönstedt
Keyword(s):  
High Pressure ◽  
Conceptual Design ◽  
Exergy Analysis ◽  
Second Law ◽  
Turbofan Engine ◽  
Blade Height ◽  
Turbofan Engines ◽  
Fuel Burn ◽  
Aero Engine ◽  
Last Stage

Although the benefits of intercooling for aero-engine applications have been realized and discussed in many publications, quantitative details are still relatively limited. In order to strengthen the understanding of aero-engine intercooling, detailed performance data on optimized intercooled (IC) turbofan engines are provided. Analysis is conducted using an exergy breakdown, i.e., quantifying the losses into a common currency by applying a combined use of the first and second law of thermodynamics. Optimal IC geared turbofan engines for a long range mission are established with computational fluid dynamics (CFD) based two-pass cross flow tubular intercooler correlations. By means of a separate variable nozzle, the amount of intercooler coolant air can be optimized to different flight conditions. Exergy analysis is used to assess how irreversibility is varying over the flight mission, allowing for a more clear explanation and interpretation of the benefits. The optimal IC geared turbofan engine provides a 4.5% fuel burn benefit over a non-IC geared reference engine. The optimum is constrained by the last stage compressor blade height. To further explore the potential of intercooling the constraint limiting the axial compressor last stage blade height is relaxed by introducing an axial radial high pressure compressor (HPC). The axial–radial high pressure ratio (PR) configuration allows for an ultrahigh overall PR (OPR). With an optimal top-of-climb (TOC) OPR of 140, the configuration provides a 5.3% fuel burn benefit over the geared reference engine. The irreversibilities of the intercooler are broken down into its components to analyze the difference between the ultrahigh OPR axial–radial configuration and the purely axial configuration. An intercooler conceptual design method is used to predict pressure loss heat transfer and weight for the different OPRs. Exergy analysis combined with results from the intercooler and engine conceptual design are used to support the conclusion that the optimal PR split exponent stays relatively independent of the overall engine PR.

Multidisciplinary Analysis of a Geared Fan Intercooled Core Aero-Engine

2013 ◽  
Vol 136 (1) ◽  
Author(s):  
Konstantinos G. Kyprianidis ◽  
Andrew M. Rolt ◽  
Tomas Grönstedt
Keyword(s):  
Exchange Rates ◽  
Fuel Consumption ◽  
Pressure Ratio ◽  
Velocity Ratio ◽  
Specific Fuel Consumption ◽  
Design Parameters ◽  
Aircraft Model ◽  
Aero Engine ◽  
Jet Velocity ◽  
Analytical Expressions

The reduction of CO2 emissions is strongly linked with the improvement of engine specific fuel consumption, along with the reduction of engine nacelle drag and weight. One alternative design approach to improving specific fuel consumption is to consider a geared fan combined with an increased overall pressure ratio intercooled core performance cycle. The thermal benefits from intercooling have been well documented in the literature. Nevertheless, there is very little information available in the public domain with respect to design space exploration of such an engine concept when combined with a geared fan. The present work uses a multidisciplinary conceptual design tool to analyze the option of an intercooled core geared fan aero engine for long haul applications with a 2020 entry into service technology level assumption. With minimum mission fuel in mind, the results indicate as optimal values a pressure ratio split exponent of 0.38 and an intercooler mass flow ratio of 1.18 at hot-day top of climb conditions. At ISA midcruise conditions a specific thrust of 86 m/s, a jet velocity ratio of 0.83, an intercooler effectiveness of 56%, and an overall pressure ratio value of 76 are likely to be a good choice. A 70,000 lbf intercooled turbofan engine is large enough to make efficient use of an all-axial compression system, particularly within a geared fan configuration, but intercooling is perhaps more likely to be applied to even larger engines. The proposed optimal jet velocity ratio is actually higher than the value one would expect by using standard analytical expressions, primarily because this design variable affects core efficiency at mid-cruise due to a combination of several different subtle changes to the core cycle and core component efficiencies at this condition. The analytical expressions do not consider changes in core efficiency and the beneficial effect of intercooling on transfer efficiency, nor do they account for losses in the bypass duct and jet pipe, while a relatively detailed engine performance model, such as the one utilized in this study, does. Mission fuel results from a surrogate model are in good agreement with the results obtained from a rubberized-wing aircraft model for some of the design parameters. This indicates that it is possible to replace an aircraft model with specific fuel consumption and weight penalty exchange rates. Nevertheless, drag count exchange rates have to be utilized to properly assess changes in mission fuel for those design parameters that affect nacelle diameter.

Analyses and Optimization of a Propulsion Cycle for Unmixed High Bypass Turbofan

2008 ◽  
Author(s):  
Adel Ghenaiet
Keyword(s):  
Fuel Consumption ◽  
Pressure Ratio ◽  
Specific Fuel Consumption ◽  
Design Parameters ◽  
Inlet Temperature ◽  
Turbine Inlet Temperature ◽  
High Pressure Ratio ◽  
Turbine Inlet ◽  
Design Variables ◽  
Bypass Ratio

This paper deals with a parametric study and an optimization for the design variables of a high bypass unmixed turbofan equipping commercial aircrafts. The objective of the first part of this study is to highlight the effects of the principal design parameters (bypass ratio, compression ratios, turbine inlet temperature etc..) on the uninstalled performance, in terms of specific thrust and specific fuel consumption. The second part concerns the optimization, aiming at finding the optimum design parameters concurrently minimizing the specific fuel consumption at cruise, and meeting the thrust requirement at takeoff. The cycle analyzer (on-design and off-design) as coupled to the optimization algorithm MMFD by adopting a random multi-starts search strategy is shown to be stable and converging. The predefined requirements and constraints have dictated utilizing an engine with a high-bypass ratio, high-pressure ratio and a moderate turbine inlet temperature. In general, the obtained results compare fairly well with typical data available for an equivalent ‘reference’ engine. This elaborated methodology is shown to be consistent with the conceptual design requirements and accuracy, because, it does not use components’ characteristics, and operates on simplifying assumptions. This present methodology can be readily adapted for other configurations of aero-engines as well, and easily integrated in a multi-disciplinary design approach.

Undergraduate Gas Turbine Engine Design Using Spreadsheets and Commercial Software

2000 ◽  
Author(s):  
Kenneth W. Van Treuren ◽  
Brenda A. Haven
Keyword(s):  
Gas Turbine ◽  
Fuel Consumption ◽  
Pressure Ratio ◽  
Sensitivity Study ◽  
Commercial Software ◽  
Turbine Engine ◽  
Mission Analysis ◽  
Engine Design ◽  
Student Team ◽  
Engine Cycle

A unique, three-part undergraduate gas turbine engine design project was developed to acquaint students, working in teams of two or three, with the process of engine cycle selection. The design application is a low-flying, Close Air Support (CAS) aircraft using a separate exhaust turbofan engine. Both spreadsheets and commercial software are used. The commercial software is included with the course textbook, “Elements of Gas Turbine Propulsion” by Dr Jack D. Mattingly. Using commercial software, reinforced by classroom lectures, allows the students to focus on the design decisions. The first part of the project is Mission Analysis which introduces the student teams to the design problem. A spreadsheet template is given to each student team that includes aircraft and mission profile specifications. The students must complete the spreadsheet and develop the relationships for lift, drag, thrust required, and fuel burn to calculate a useable fuel remaining at the end to the mission. The spreadsheet allows the students to obtain an average specific fuel consumption that results in 1500 lbm of fuel remaining at the end of the mission. This target value is used in the second part of the design process, on-design Parametric Cycle Analysis (PCA), as a basis for engine cycle selection. Parametric Cycle Analysis is accomplished using the program PARA.EXE. PARA.EXE generates a carpet plot of possible engine design choices by varying the compressor pressure ratio, bypass ratio, and fan pressure ratio. From these carpet plots the students must identify three possible engine cycles that meet the target value for specific fuel consumption found during the mission analysis. Tradeoffs between thrust and fuel consumption are discussed and the students are required to justify their choices for the engine cycle. The last part of the project is the off-design Engine Performance Analysis (EPA) using the program PERF.EXE. The chosen engines must fly the mission and meet the required performance and mission constraint. Based on the overall mission performance, the students narrow the field of three possible engine cycles to one. Each student team then does a sensitivity study to determine if there is an additional benefit for slight changes in the design choices. The result of this sensitivity study is the students’ final engine cycle. With this cycle, an additive drag calculation is made using the program DADD.EXE to account for losses (off-design) and these losses are then factored back into the performance spreadsheet to check the engine’s capabilities for completing the mission. The iterative nature of the design process is emphasized throughout but only one pass through the process is accomplished. Units are given in English Engineering, as that is what is required for the project. Both SI and English Engineering units are taught in the course.

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 Ultrashort Nacelles for Low Fan Pressure Ratio Propulsors

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Andreas Peters ◽  
Zoltán S. Spakovszky ◽  
Wesley K. Lord ◽  
Becky Rose
Keyword(s):  
Mach Number ◽  
Pressure Ratio ◽  
Design Tool ◽  
Operating Conditions ◽  
Design Parameters ◽  
Flow Distortion ◽  
Flow Area ◽  
Propulsive Efficiency ◽  
Flow Acceleration ◽  
Fuel Burn

As the propulsor fan pressure ratio (FPR) is decreased for improved fuel burn, reduced emissions and noise, the fan diameter grows and innovative nacelle concepts with short inlets are required to reduce their weight and drag. This paper addresses the uncharted inlet and nacelle design space for low-FPR propulsors where fan and nacelle are more closely coupled than in current turbofan engines. The paper presents an integrated fan–nacelle design framework, combining a spline-based inlet design tool with a fast and reliable body-force-based approach for the fan rotor and stator blade rows to capture the inlet–fan and fan–exhaust interactions and flow distortion at the fan face. The new capability enables parametric studies of characteristic inlet and nacelle design parameters with a short turn-around time. The interaction of the rotor with a region of high streamwise Mach number at the fan face is identified as the key mechanism limiting the design of short inlets. The local increase in Mach number is due to flow acceleration along the inlet internal surface coupled with a reduction in effective flow area. For a candidate short-inlet design with length over diameter ratio L/D = 0.19, the streamwise Mach number at the fan face near the shroud increases by up to 0.16 at cruise and by up to 0.36 at off-design conditions relative to a long-inlet propulsor with L/D = 0.5. As a consequence, the rotor locally operates close to choke resulting in fan efficiency penalties of up to 1.6% at cruise and 3.9% at off-design. For inlets with L/D < 0.25, the benefit from reduced nacelle drag is offset by the reduction in fan efficiency, resulting in propulsive efficiency penalties. Based on a parametric inlet study, the recommended inlet L/D is suggested to be between 0.25 and 0.4. The performance of a candidate short inlet with L/D = 0.25 was assessed using full-annulus unsteady Reynolds-averaged Navier–Stokes (RANS) simulations at critical design and off-design operating conditions. The candidate design maintains the propulsive efficiency of the baseline case and fuel burn benefits are conjectured due to reductions in nacelle weight and drag compared to an aircraft powered by the baseline propulsor.

Multi-Objective Optimization of a Regenerative Rotorcraft Powerplant: Quantification of Fuel Economy and Environmental Impact

2015 ◽  
Author(s):  
Fakhre Ali ◽  
Konstantinos Tzanidakis ◽  
Ioannis Goulos ◽  
Vassilios Pachidis ◽  
Roberto d’Ippolito
Keyword(s):  
Heat Exchanger ◽  
Environmental Impact ◽  
Fuel Economy ◽  
Pareto Front ◽  
Pressure Ratio ◽  
Nox Emissions ◽  
Fuel Burn ◽  
Engine Design ◽  
Compressor Pressure Ratio ◽  
Pressure Compressor

A computationally efficient and cost effective simulation framework has been proposed to perform a multidisciplinary design and optimization of a conceptual regenerative rotorcraft powerplant configuration at mission level. A generic rotorcraft model, representative of a modern twin-engine light civil rotorcraft has been investigated, operating under a representative passenger air taxi mission. The design space corresponding to the conceptual regenerative engine thermodynamic cycle parameters as well as engine and mission design outputs in terms of low pressure compressor pressure ratio, high pressure compressor pressure ratio, turbine entry temperature, mass flow, heat exchanger effectiveness, engine design point specific fuel consumption, engine weight, mission fuel burn and mission CO2 and NOx emissions has been thoroughly investigated through the application of a latin hypercube sampling, design of experiment approach. The interdependencies between the various engine design inputs/outputs are quantified by establishing the corresponding linear correlations between the aforementioned engine inputs/outputs as well as for the corresponding mission output parameters. A multi-objective Particle Swarm Optimizer is employed to derive Pareto front models quantifying the optimum interrelationship between the mission fuel burn and NOx emissions inventory. The acquired engine cycle design parameters corresponding to the span of the Pareto front suggest that the heat exchanger design effectiveness is the key design parameter representing the interdependency between engine fuel economy and environmental impact. The acquired optimum engine models, obtained from the Pareto front, are subsequently deployed for the design of conceptual rotorcraft engine configurations, targeting improved mission fuel economy, enhanced payload-range capability and overall environmental impact.

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