Exploring GasTurb 12 for Supplementary Use on an Introductory Propulsion Design Project

2017 ◽  
Author(s):  
Aaron R. Byerley ◽  
Kurt P. Rouser ◽  
Devin O. O’Dowd
Keyword(s):  
System Analysis ◽  
Pressure Ratio ◽  
Supplementary Information ◽  
Design System ◽  
Inlet Temperature ◽  
Turbofan Engine ◽  
Design Project ◽  
Engine Design ◽  
Aeronautical Engineering ◽  
Engine Cycle

The purpose of this paper is to explore GasTurb 12, a commercial gas turbine engine performance simulation program, for supplementary use on an introductory propulsion design project in an undergraduate course. This paper will describe several possible opportunities for supplementing AEDsys (Aircraft Engine Design System Analysis) version 4.012, the engine design software tool currently in use. The project is assigned to juniors taking their first propulsion course in the aeronautical engineering major at the USAF Academy. This course, Aeronautical Engineering 361, which focuses on cycle analysis and selection, is required of all aero majors and is used to satisfy the ABET Program Criterion requiring knowledge of propulsion fundamentals. This paper describes the most recent design project that required the students to re-engine the USAF T-38 with the aim of competing for the Advanced Pilot Training Program (T-X) program. The goal of the T-X program is to replace the T-38 aircraft that entered service in 1961 with an aircraft capable of sustained high-G operations that is also more fuel efficient. The design project required the students to select an engine-cycle for a single, non-afterburning, mixed stream, low bypass turbofan engine to replace the two J85 turbojets currently in the T-38. It was anticipated that the high specific thrust requirements might possibly be met through the use of modern component measures of merit to include a much higher turbine inlet temperature. Additionally, it was anticipated that the required 10% reduction in thrust specific fuel consumption might possibly be achieved by using a turbofan engine cycle with a higher overall pressure ratio. This paper will describe the use of GasTurb 12 to perform the same design analysis that was described above using AEDsys as well as additional features such as numerical optimization, temperature-entropy diagrams, and the generation of scaled, two-dimensional engine geometry drawings. The paper will illustrate how GasTurb 12 offers important supplementary information that will deepen student understanding of engine cycle design and analysis.

Computational Aid to AEDsys for Designing a Variable-Area Gas Turbine Nozzle

2018 ◽  
Author(s):  
Devin O. O’Dowd ◽  
Aaron R. Byerley
Keyword(s):  
Gas Turbine ◽  
United States Air Force ◽  
System Analysis ◽  
Graphical Representation ◽  
Aircraft Engine ◽  
The United States ◽  
Design System ◽  
Engine Design ◽  
Critical Feedback ◽  
Gas Turbine Nozzle

This paper presents a practical approach to designing a gas turbine nozzle with the help of the Aircraft Engine Design textbook as well as the software program Nozzle, a subprogram within the Aircraft Engine Design System Analysis Software suite AEDsys. The current textbook and software allow for a variable wetted length of the converging and diverging nozzle sections. Critical feedback from industry experts has inspired an attempt to design a nozzle with fixed wetted material lengths. This paper is written to augment classroom treatment, but will also support others who use the Aircraft Engine Design text and software for a preliminary engine design capstone. This approach is further guided by the actual scaling of the Pratt & Whitney F100 variable geometry converging-diverging nozzle, where wetted lengths are fixed. The chief goal is to equip students at the United States Air Force Academy with a refined approach that is more realistic of a manufactured nozzle design, producing a graphical representation of a nozzle schedule at different speed and altitude flight conditions, both with and without afterburner.

Sector Tests of a Low NOx, Lean-Direct-Injection, Multipoint Integrated Module Combustor Concept

2002 ◽  
Author(s):  
Robert Tacina ◽  
Changlie Wey ◽  
Peter Laing ◽  
Adel Mansour
Keyword(s):  
Direct Injection ◽  
Pressure Ratio ◽  
Inlet Pressure ◽  
Inlet Temperature ◽  
Single Element ◽  
Nox Emissions ◽  
Fuel Injectors ◽  
Lean Direct Injection ◽  
Test Conditions ◽  
Engine Cycle

Results of a low-NOx combustor test with a 15° sector are presented. A multipoint, lean-direct injection concept is used. The configuration tested has 36 fuel injectors and fuel-air mixers in place of a dual annular arrangement of two conventional fuel injectors. An integrated-module approach is used for the construction where chemically etched laminates that are diffusion bonded, combine the fuel injectors, air swirlers and fuel manifold into a single element. Test conditions include inlet temperatures up to 866K, and inlet pressures up to 4825 kPa. The fuel used was Jet A. A correlation is developed relating the NOx emissions to the inlet temperature, inlet pressure, and fuel-air ratio. Using a hypothetical 55:1 pressure-ratio engine, cycle NOx emissions are estimated to be less than 40% of the 1996 ICAO standard.

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.

An Investigation on Turbine Tip and Shroud Heat Transfer

2003 ◽  
Vol 125 (3) ◽  
pp. 513-520 ◽  
Author(s):  
Kam S. Chana ◽  
Terry V. Jones
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Turbine Stage ◽  
Pressure Distributions ◽  
Engine Design ◽  
Nozzle Guide

Detailed experimental investigations have been performed to measure the heat transfer and static pressure distributions on the rotor tip and rotor casing of a gas turbine stage with a shroudless rotor blade. The turbine stage was a modern high pressure Rolls-Royce aero-engine design with stage pressure ratio of 3.2 and nozzle guide vane (ngv) Reynolds number of 2.54E6. Measurements have been taken with and without inlet temperature distortion to the stage. The measurements were taken in the QinetiQ Isentropic Light Piston Facility and aerodynamic and heat transfer measurements are presented from the rotor tip and casing region. A simple two-dimensional model is presented to estimate the heat transfer rate to the rotor tip and casing region as a function of Reynolds number along the gap.

Comprehensive Performance Analysis of the Turbofan With a Multi-Annular Rotating Detonation Duct Burner

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
Zifei Ji ◽  
Huiqiang Zhang ◽  
Bing Wang ◽  
Wei He
Keyword(s):  
Performance Analysis ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Shaft Speed ◽  
Analysis Model ◽  
Operating Characteristics ◽  
Turbofan Engine ◽  
Surge Line ◽  
Duct Burner ◽  
Rotating Detonation

Abstract The performance analysis of mixed-exhaust turbofan engine with multi-annular rotating detonation duct burner (RDDB) is conducted for the first time, considering that the flow path of the bypass duct is ideal for a rotating detonation combustor (RDC). The configuration of the multi-annular rotating detonation combustor is constructed aiming at the advantages of a wider operation range and uniform outlet parameters over the single-annular one. Then, a parametric analysis model of the mixed-exhaust turbofan engine with a rotating detonation duct burner is developed. Thereafter, the effects of duct burner parameters on the engine performance and operating characteristics are investigated. The mixed-exhaust turbofan engine with a rotating detonation duct burner shows superior overall performance to that of one with an isobaric afterburner (ICAB) over a wide operation range. The separate-exhaust rotating detonation duct burner can hold characteristics that are higher than those of the mixed-exhaust one at lower values of fan pressure ratio, while the mixed-exhaust one corresponds to lower values of turbine inlet temperature. When the rotating detonation duct burner is “on,” the low-pressure rotor operating line moves toward the surge line on the low corrected shaft speed side but away from the surge line on the high corrected shaft speed side.

A Methodology for Measuring Turbocharger Adiabatic Maps in a Gas-Stand and its Usage for Calibrating Control Oriented and 1D Models at Early ICE Design Stages

2019 ◽  
Author(s):  
José Ramón Serrano ◽  
Francisco José Arnau ◽  
Luis Miguel García-Cuevas ◽  
Alejandro Gómez-Vilanova ◽  
Stephane Guilain ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Internal Combustion Engines ◽  
Gas Flow ◽  
System Analysis ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Boost Pressure ◽  
Gas Temperature ◽  
Hot Gas ◽  
The Impact

Abstract Turbocharged engines are the standard architecture for designing efficient spark ignition and compression ignition reciprocating internal combustion engines (ICE). Turbochargers characterization and modeling are basic tasks for the analysis and prediction of the whole engine system performance and this information is needed in quite early stages of the engine design. Turbocharger characteristics (efficiency, pressure ratio, mass flow rates...) traditionally rely in maps of pseudo non-dimensional variables called reduced variables. These maps must be used by reciprocating ICE designer and modeler not only for benchmarking of the turbocharger, but for a multiplicity of purposes, i.e: assessing engine back-pressure, boost pressure, load transient response, after-treatment inlet temperature, intercooler inlet temperature, low pressure EGR temperature, ... Maps of reduced variables are measured in gas-stands with steady flow but non-standardized fluids conditioning; neither temperatures nor flows. In concrete: turbine inlet gas temperature; lubrication-oil flow and temperature; water-cooling flow and turbo-machinery external heat transfer are non-standardized variables which have a big impact in assessing said multiplicity of purposes. Moreover, adiabatic efficiency, heat losses and friction losses are important data, hidden in the maps of reduced variables, which depend on the testing conditions as much as on the auxiliary fluids temperature and flow rate. In this work it is proposed a methodology to standardize turbochargers testing based in measuring the maps twice: in close to adiabatic and in diathermal conditions. Along the paper it is discussed with special detail the impact of the procedure followed to achieve said quasi-adiabatic conditions in both the energy balance of the turbocharger and the testing complexity. As a conclusion, the paper proposes a methodology which combines quasi-adiabatic tests (cold and hot gas flow) with diathermal tests (hot gas flow) in order to extract from a turbocharger gas-stand all information needed by engine designers interested in controlling or 1D-modelling the ICE. The methodology is completed with a guide for calibrating said control-oriented turbocharger models in order to separate aerodynamic efficiency (adiabatic) from heat transfer losses and from friction losses in the analysis of the turbocharger performance. The outsourced calibration of the turbocharger model allows avoiding uncertainties in the global ICE model calibration, what is very interesting for turbochargers benchmarking at early ICE-turbo matching stages or for global system analysis at early control design stages.

Micro-Turbine Design Point Definition Using Genetic Algorithm With Single and Multi-Objective Optimization

2012 ◽  
Author(s):  
Diogo F. Cavalca ◽  
Cleverson Bringhenti
Keyword(s):  
Genetic Algorithm ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Specific Work ◽  
Multi Objective Optimization ◽  
Total Cost ◽  
Design Point ◽  
Multi Objective ◽  
Engine Design

During a gas turbine development phase an important engineer task is to find the appropriate engine design point that meet the required specifications. This task can be very arduous because all possible operating points in the gas turbine operational envelope need to be analyzed, for the sake of verification of whether or not the established performance might be achieved. In order to support engineers to best define the engine design point that meet required performance a methodology was developed in this work. To accomplish that a computer program was written in Matlab®. In this program was incorporated the thermoeconomic and thermodynamic optimization. The thermodynamic calculation process was done based in enthalpy and entropy function and then validated using a commercial program. The methodology uses genetic algorithm with single and multi-objective optimization. The micro gas turbine cycle chosen to study was the recuperated. The cycle efficiency, total cost and specific work were chosen as objective functions, while the pressure ratio, compressor and turbine polytropic efficiencies, turbine inlet temperature and heat exchange effectiveness were chosen as decision variables. For total cost were considered the fixed costs (equipment, installation, taxes, etc.) and variable costs (fuel, environmental and O&M). For emissions were taken into account the NOx, CO and UHC. An economic analysis was done for a recuperated cycle showing the costs behavior for different optimized design points. The optimization process was made for: single-objective, where each objective was optimized separately; two-objectives, where they were optimized in pairs; three-objectives, where it was optimized in trio. After, the results were compared each other showing the possible design points.

scholarly journals An Automated Interactive Design System for Advanced Gas Turbines

1974 ◽  
Author(s):  
K. M. Thomas ◽  
J. J. Piendel
Keyword(s):  
Gas Turbines ◽  
Design System ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Air Cooling ◽  
Gas Turbine Engines ◽  
Turbine Inlet Temperature ◽  
The Past ◽  
Engine Design ◽  
Turbine Design

In the past ten years there has been a dramatic increase in turbine inlet temperature in aircraft gas turbine engines. This increase has been made possible by the application of extensive air cooling to turbine parts. The attendent increase in turbine design complexity without an increase in engine design or development time has been made possible by the development of modern computers and computer programs. A computerized turbine automated design system (TADSYS) was developed at Pratt and Whitney Aircraft, which makes extensive use of computer graphics, to meet the needs of modern turbine design.

Estimating Gas Turbine Engine Weight, Costs, and Development Time During the Preliminary Aircraft Engine Design Process

2013 ◽  
Author(s):  
Aaron R. Byerley ◽  
August J. Rolling ◽  
Kenneth W. Van Treuren
Keyword(s):  
Risk Factors ◽  
Gas Turbine ◽  
Design Process ◽  
Development Time ◽  
Aircraft Engine ◽  
Engineering Majors ◽  
Cost Estimating ◽  
Engine Design ◽  
Aeronautical Engineering ◽  
Engine Cycle

This paper describes the application of a weight and cost-estimating methodology used in an undergraduate aircraft engine design course that is taught in concert with a companion course in airframe design. The two preliminary designs, one for the engine and the other for the airframe, must be integrated as subsystems within a system to satisfy the performance requirements of a given mission as outlined in a single “request for proposals”. In recent years, systems engineering management majors have been added to the design teams to work alongside the aeronautical engineering majors to analyze and report on costs, schedule, and technical risk factors in addition to the operational performance factors that have previously been the sole focus of the course. The teaming of technical management majors and aeronautical engineering majors has been driven by a heightened emphasis on system affordability. The cost-estimating methodology for gas turbine engines uses cycle parameters such as turbine rotor inlet temperature, overall pressure ratio, specific fuel consumption, level of technology, and engine dry weight as inputs. A methodology for estimating dry engine weight was developed which uses engine cycle parameters and fan face diameter as inputs in a volume analog scaling factor which was correlated against historical engine weight data. To tie all of the performance, weight, cost, and development time issues together, the paper presents an “analysis of alternatives” example that considers three different engine cycle alternatives. The design tools presented in this paper will provide a strong foundational understanding of how to systematically weigh and evaluate the important tradeoffs between aircraft turbofan engine performance, cost, schedule, and risk factors. Equipping students with the insight and ability to perform these multidisciplinary trade studies during the preliminary engine design process is this paper’s most important contribution.

An Investigation on Turbine Tip and Shroud Heat Transfer

2002 ◽  
Author(s):  
Kam S. Chana ◽  
Terry V. Jones
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Turbine Stage ◽  
Pressure Distributions ◽  
Engine Design ◽  
Nozzle Guide

Detailed experimental investigations have been performed to measure the heat transfer and static pressure distributions on the rotor tip and rotor casing of a gas turbine stage with a shroud-less rotor blade. The turbine stage was a modern high pressure Rolls-Royce aero-engine design with stage pressure ratio of 3.2 and nozzle guide vane (ngv) Reynolds number of 2.54E6. Measurements have been taken with and without inlet temperature distortion to the stage. The measurements were taken in the QinetiQ Isentropic Light Piston Facility and aerodynamic and heat transfer measurements are presented from the rotor tip and casing region. A simple two-dimensional model is presented to estimate the heat transfer rate to the rotor tip and casing region as a function of Reynolds number along the gap.

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