DOCSLIB.ORG

Simulating Indirect Thrust Measurement Methods for High

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Simulating Indirect Thrust Measurement Methods for High THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 93-G-335 345 E. 47th St., New York, N.Y. 10017 The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the Society or of its Divisions or Sections, © or printed in its publications. Discussion is printed only if the paper is pub- lished in an ASME Journal, Papers are available from ASME for 15 months after the meeting. Printed in U.S.A. Copyright © 1993 by ASME Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1993/78910/V03BT16A091/2403731/v03bt16a091-93-gt-335.pdf by guest on 26 September 2021 SIMUAIG IIEC US MEASUEME MEOS O IG YASS UOAS . Stvnn nd . Srvntt Department of Mechanical and Aerospace Engineering Carleton University Ottawa, Ontario, Canada ABSTRACT ABBREVIATIONS As yet, there is no known reliable method for directly measuring BPR bypass ratio the thrust of a turbofan in flight. Manufacturers of civil turbofans CDP HP Compressor delivery total pressure use various indirect thrust measurements to indicate performance EPR engine pressure ratio of an engine to the flight deck. Included among these are: Engine EGT LP Turbine exhaust gas total temperature Pressure Ratio (EPR), Integrated Engine Pressure Ratio (IEPR), Fan FPR fan pressure ratio mechanical speed (N,), and various Turbine Gas Temperatures such GG gas generator (core of turbofan) as ITT or EGT. Of key concern is whether these thrust indicators HDTO hot day take-off (ISA +15°C at SLS) give an accurate account of the actual engine thrust. The accuracy HPC,HPT high pressure compressor, turbine of these methods, which are crucial at take-off, may be IEPR integrated engine pressure ratio compromised by various types of common engine deterioration, to ISA international standard atmosphere the point where a thrust indicator may give a false indication of the ITT inter-turbine total temperature health and thrust of the engine. A study was done to determine the LPC,LPT low pressure compressor (Boosters), turbine effect of advanced engine cycles on typical values of these OPR overall pressure ratio parameters. A preliminary investigation of the effects of common PR pressure ratio kinds of turbofan deterioration was conducted to see how these RPR ram pressure ratio faults can affect both actual engine performance and the indirect SLS sea level static thrust indicators. TIT HP turbine inlet total temperature TOC top of climb (Mach 0.8 @ 35000 ft) NOMENCLATURE SUBSCRIPTS AND STATION NUMBERING A area bl HP Turbine cooling bleed (% core massflow) a air C airflow velocity c cold (fan or bypass) c, specific heat (at constant pressure) cc combustion chamber m massflow g gas h hot (core) N1 , N2 mechanical LP and HP spool speeds p, T static pressure, temperature i intake mech P. T. ambient pressure, temperature mechanical N nozzle plane PPb Combustion pressure drop (% CDP) po, To total pressure, temperature PN , TN nozzle plane static pressure, temperature Q non-dimensional massflow parameter corrected pressure (pa/l.01325, po in Bars) specific heat ratio 11_ polytropic efficiency 11 isentropic efficiency O corrected temperature (T0/288.16, To in K) Presented at the International Gas Turbine and Aeroengine Congress and Exposition Cincinnati, Ohio May 242, This paper has been accepted for publication in the Transactions of the ASME Discussion of it will be accepted at ASME Headquarters until September 0, Hence, F = pN A CN +A (PN -p) (2) 0 2 a 8 A oe = PN A • 2 C (T - TN) + A (pn - p°) N 3 5 7Coe = p A PN P° p ° _ 1 CCoe 9 p - 1 p° p° TN C HP LPT +A ° — — - 1 ° Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1993/78910/V03BT16A091/2403731/v03bt16a091-93-gt-335.pdf by guest on 26 September 2021 p°Figure 1 Twin-Spool Turbofan Station Numbering ° (3) INTRODUCTION Substituting the value of pN : P° Although jet engines have been used commercially for 40 years, there is still no direct method of measuring thrust in flight. 2 Indirect methods of indicating thrust are used, based on pressure =1 _L _ measurements or rotational speeds. The use of engine pressure A° Y-1 1l1 ° Y2 (4 ratio (EPR), which is defined as the ratio of total exhaust pressure + [(2 - over total inlet pressure, became dominant in early turbojet engines P. where there was a direct relationship between EPR and thrust. A problem arose with the introduction of high bypass turbofans, This gives, where a substantial portion of the total net thrust is provided by the cold or bypass stream; EPR considered only the core engine F_ P° 11 2 pressure ratio and hence the thrust contribution from the hot (Y+ 1) - 1 (5) stream. One alternative was to use integrated engine pressure ratio A P° P° Y+1 ) (IEPR) which is based on an area weighted average of the bypass and core nozzle pressure ratios. Another approach was to use fan It follows that rotational speed (N,) to indicate thrust. Advanced cycle turbofans may operate at significantly F = K p° - 1 where K = f (1) (6) ige aues o yass aio ( a oea essue aio (OPR) which may influence the choice of parameter for indicating A ° ° thrust. The effects of in-service deterioration may also be This can be extended to work with EPR instead, and noting that important in selecting the most suitable parameter. = P_' P°t This paper analyses the performance of a typical high i.e. nozzle PR = EPR RPR performance turbofan in the hope of providing objective data for P. P°I P. comparing the usefulness of various thrust setting parameters. Then, THE ORIGINS OF EPR F + 1l I RPR = K(EPR) (7) AP, J It is instructive to consider the basis of EPR as a means of L indicating the thrust of a simple, fixed nozzle turbojet: (where K = 1.2594 for y = 1.333) Thus, for a fixed nozzle turbojet, the ideal thrust can readily P0 Pa be found from the EPR. In the case of a turbofan, however, the To EPR will depend on both the bypass ratio and fan pressure ratio ------------- :i N sice o o ese ae a eec o e oa essue o e LPT ^ TN Turbine exhaust, and there is no longer a direct relationship Exit between total net thrust and EPR. The concept of IEPR attempts — — -^ to include the substantial thrust contribution from the fan or bypass Figure 2 nozzle through an area weighted average defined as follows: For a choked nozzle, IEPR = po2A0 + po7At' (8) pot (A, + Ah) P Y F = m CN + A (PN - p°) , andd = (2 )Fr (1) This requires extensive pressure measurements in the fan duct but p° Y+1 appears to be a more logical way to indicate the thrust of a high Also: bypass turbofan. The thermodynamic derivation of EPR is straightforward, PN 2 but to pilots, EPR may appear to be an abstract concept which = pAC C =2 CP (° N) PN = RN 7 ; does not give an instinctive "feel" for thrust developed. Rotational 2 speeds, such as N1 , can readily be related to thrust; anyone who has ever driven a car equipped with a tachometer can readily recognize COMPUTER SIMULATION METHODS USED red line limits and instinctively keeps within them. a) On-Design Model CHOICE OF DESIGN POINT The off-design performance calculation procedure used required detailed knowledge of several design point parameters of Modern high-bypass turbofans are designed around several the chosen engine cycle, particularly the engine cycle parameters flight conditions or design points, and not just the Sea Level Static such as FPR, BPR, OPR and TIT and fixed nozzle areas. A (SLS) condition as was perhaps the common design point in early straightforward thermodynamic cycle calculation was used based on engines when adequate take-off thrust was the primary concern. LP and HP spool work balances, polytropic compression and Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1993/78910/V03BT16A091/2403731/v03bt16a091-93-gt-335.pdf by guest on 26 September 2021 Current modem designs are usually optimized around three expansion processes, and separate nozzles. While the precise common flight conditions (Philpot, 1992): simulation of a particular engine was not the aim of this study, reference was made to manufacturers' information on similar high Sea Level Static (SLS) Takeoff: Usually at ISA +15°C, 1) bypass designs so that reasonable on-design performance could be to allow a flat rated take-off thrust up to some hot ambient simulated. This, in general, required a number of iterations since temperature condition (Hot Day Take-off, or HDTO). Most many of the design parameters (such as component efficiencies) are modem engines automatically control the take-off fuel flow to essentially educated guesses, especially since TOC engine cycle prevent the TIT from exceeding a specified safety limit. In parameters are rarely quoted. practical terms, the Inter-turbine Temperature (ITT) is used as the limiter due to difficulties with directly measuring the high TITS b) Off- common in modem turbofans. The HDTO condition is important, Design Simulation Model limiting the maximum allowable TIT and hence influencing the ultimate thermodynamic cycle choice. Note that the ITT is The off-design simulation procedure used in this study was the twin-spool turbofan matching procedure based on the use of sometimes called the gas generator Exhaust Gas Temperature (EGT) by many engine manufacturers. In this paper, EGT will be component characteristics or performance maps. This straight- used to denote the LP Turbine exhaust temperature. forward procedure is outlined only briefly below; a more detailed description can be found in Gas Turbine Theory (Cohen, Rogers and Saravanamuttoo, 1987). 2) Cruise: A majority of the fuel burned during a typical civil airliner flight will be during cruise, especially for long range.
Load more

Recommended publications
  • Aircraft Engine Performance Study Using Flight Data Recorder Archives
    Aircraft Engine Performance Study Using Flight Data Recorder Archives Yashovardhan S. Chati∗ and Hamsa Balakrishnan y Massachusetts Institute of Technology, Cambridge, Massachusetts, 02139, USA Aircraft emissions are a significant source of pollution and are closely related to engine fuel burn. The onboard Flight Data Recorder (FDR) is an accurate source of information as it logs operational aircraft data in situ. The main objective of this paper is the visualization and exploration of data from the FDR. The Airbus A330 - 223 is used to study the variation of normalized engine performance parameters with the altitude profile in all the phases of flight. A turbofan performance analysis model is employed to calculate the theoretical thrust and it is shown to be a good qualitative match to the FDR reported thrust. The operational thrust settings and the times in mode are found to differ significantly from the ICAO standard values in the LTO cycle. This difference can lead to errors in the calculation of aircraft emission inventories. This paper is the first step towards the accurate estimation of engine performance and emissions for different aircraft and engine types, given the trajectory of an aircraft. I. Introduction Aircraft emissions depend on engine characteristics, particularly on the fuel flow rate and the thrust. It is therefore, important to accurately assess engine performance and operational fuel burn. Traditionally, the estimation of fuel burn and emissions has been done using the ICAO Aircraft Engine Emissions Databank1. However, this method is approximate and the results have been shown to deviate from the measured values of emissions from aircraft in operation2,3.
    [Show full text]
  • 2.0 Axial-Flow Compressors 2.0-1 Introduction the Compressors in Most Gas Turbine Applications, Especially Units Over 5MW, Use Axial fl Ow Compressors
    2.0 Axial-Flow Compressors 2.0-1 Introduction The compressors in most gas turbine applications, especially units over 5MW, use axial fl ow compressors. An axial fl ow compressor is one in which the fl ow enters the compressor in an axial direction (parallel with the axis of rotation), and exits from the gas turbine, also in an axial direction. The axial-fl ow compressor compresses its working fl uid by fi rst accelerating the fl uid and then diffusing it to obtain a pressure increase. The fl uid is accelerated by a row of rotating airfoils (blades) called the rotor, and then diffused in a row of stationary blades (the stator). The diffusion in the stator converts the velocity increase gained in the rotor to a pressure increase. A compressor consists of several stages: 1) A combination of a rotor followed by a stator make-up a stage in a compressor; 2) An additional row of stationary blades are frequently used at the compressor inlet and are known as Inlet Guide Vanes (IGV) to ensue that air enters the fi rst-stage rotors at the desired fl ow angle, these vanes are also pitch variable thus can be adjusted to the varying fl ow requirements of the engine; and 3) In addition to the stators, another diffuser at the exit of the compressor consisting of another set of vanes further diffuses the fl uid and controls its velocity entering the combustors and is often known as the Exit Guide Vanes (EGV). In an axial fl ow compressor, air passes from one stage to the next, each stage raising the pressure slightly.
    [Show full text]
  • Aerospace Engine Data
    AEROSPACE ENGINE DATA Data for some concrete aerospace engines and their craft ................................................................................. 1 Data on rocket-engine types and comparison with large turbofans ................................................................... 1 Data on some large airliner engines ................................................................................................................... 2 Data on other aircraft engines and manufacturers .......................................................................................... 3 In this Appendix common to Aircraft propulsion and Space propulsion, data for thrust, weight, and specific fuel consumption, are presented for some different types of engines (Table 1), with some values of specific impulse and exit speed (Table 2), a plot of Mach number and specific impulse characteristic of different engine types (Fig. 1), and detailed characteristics of some modern turbofan engines, used in large airplanes (Table 3). DATA FOR SOME CONCRETE AEROSPACE ENGINES AND THEIR CRAFT Table 1. Thrust to weight ratio (F/W), for engines and their crafts, at take-off*, specific fuel consumption (TSFC), and initial and final mass of craft (intermediate values appear in [kN] when forces, and in tonnes [t] when masses). Engine Engine TSFC Whole craft Whole craft Whole craft mass, type thrust/weight (g/s)/kN type thrust/weight mini/mfin Trent 900 350/63=5.5 15.5 A380 4×350/5600=0.25 560/330=1.8 cruise 90/63=1.4 cruise 4×90/5000=0.1 CFM56-5A 110/23=4.8 16
    [Show full text]
  • Using an Autothrottle to Compare Techniques for Saving Fuel on A
    Iowa State University Capstones, Theses and Graduate Theses and Dissertations Dissertations 2010 Using an autothrottle ot compare techniques for saving fuel on a regional jet aircraft Rebecca Marie Johnson Iowa State University Follow this and additional works at: https://lib.dr.iastate.edu/etd Part of the Electrical and Computer Engineering Commons Recommended Citation Johnson, Rebecca Marie, "Using an autothrottle ot compare techniques for saving fuel on a regional jet aircraft" (2010). Graduate Theses and Dissertations. 11358. https://lib.dr.iastate.edu/etd/11358 This Thesis is brought to you for free and open access by the Iowa State University Capstones, Theses and Dissertations at Iowa State University Digital Repository. It has been accepted for inclusion in Graduate Theses and Dissertations by an authorized administrator of Iowa State University Digital Repository. For more information, please contact [email protected]. Using an autothrottle to compare techniques for saving fuel on A regional jet aircraft by Rebecca Marie Johnson A thesis submitted to the graduate faculty in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE Major: Electrical Engineering Program of Study Committee: Umesh Vaidya, Major Professor Qingze Zou Baskar Ganapathayasubramanian Iowa State University Ames, Iowa 2010 Copyright c Rebecca Marie Johnson, 2010. All rights reserved. ii DEDICATION I gratefully acknowledge everyone who contributed to the successful completion of this research. Bill Piche, my supervisor at Rockwell Collins, was supportive from day one, as were many of my colleagues. I also appreciate the efforts of my thesis committee, Drs. Umesh Vaidya, Qingze Zou, and Baskar Ganapathayasubramanian. I would also like to thank Dr.
    [Show full text]
  • High Pressure Ratio Intercooled Turboprop Study
    E AMEICA SOCIEY O MECAICA EGIEES 92-GT-405 4 E. 4 S., ew Yok, .Y. 00 h St hll nt b rpnbl fr ttnt r pnn dvnd In ppr r n d n t tn f th St r f t vn r Stn, r prntd In t pbltn. n rnt nl f th ppr pblhd n n ASME rnl. pr r vlbl fr ASME fr fftn nth ftr th tn. rntd n USA Copyright © 1992 by ASME ig essue aio Iecooe uoo Suy C. OGES Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 Sundstrand Power Systems San Diego, CA ASAC NOMENCLATURE High altitude long endurance unmanned aircraft impose KFT Altitude Thousands Feet unique contraints on candidate engine propulsion systems and HP Horsepower types. Piston, rotary and gas turbine engines have been proposed for such special applications. Of prime importance is the HIPIT High Pressure Intercooled Turbine requirement for maximum thermal efficiency (minimum specific Mn Flight Mach Number fuel consumption) with minimum waste heat rejection. Engine weight, although secondary to fuel economy, must be evaluated Mls Inducer Mach Number when comparing various engine candidates. Weight can be Specific Speed (Dimensionless) minimized by either high degrees of turbocharging with the Ns piston and rotary engines, or by the high power density Exponent capabilities of the gas turbine. pps Airflow The design characteristics and features of a conceptual high SFC Specific Fuel Consumption pressure ratio intercooled turboprop are discussed. The intended application would be for long endurance aircraft flying TIT Turbine Inlet Temperature °F at an altitude of 60,000 ft.(18,300 m).
    [Show full text]
  • Helicopter Turboshafts
    Helicopter Turboshafts Luke Stuyvenberg University of Colorado at Boulder Department of Aerospace Engineering The application of gas turbine engines in helicopters is discussed. The work- ings of turboshafts and the history of their use in helicopters is briefly described. Ideal cycle analyses of the Boeing 502-14 and of the General Electric T64 turboshaft engine are performed. I. Introduction to Turboshafts Turboshafts are an adaptation of gas turbine technology in which the principle output is shaft power from the expansion of hot gas through the turbine, rather than thrust from the exhaust of these gases. They have found a wide variety of applications ranging from air compression to auxiliary power generation to racing boat propulsion and more. This paper, however, will focus primarily on the application of turboshaft technology to providing main power for helicopters, to achieve extended vertical flight. II. Relationship to Turbojets As a variation of the gas turbine, turboshafts are very similar to turbojets. The operating principle is identical: atmospheric gases are ingested at the inlet, compressed, mixed with fuel and combusted, then expanded through a turbine which powers the compressor. There are two key diferences which separate turboshafts from turbojets, however. Figure 1. Basic Turboshaft Operation Note the absence of a mechanical connection between the HPT and LPT. An ideal turboshaft extracts with the HPT only the power necessary to turn the compressor, and with the LPT all remaining power from the expansion process. 1 of 10 American Institute of Aeronautics and Astronautics A. Emphasis on Shaft Power Unlike turbojets, the primary purpose of which is to produce thrust from the expanded gases, turboshafts are intended to extract shaft horsepower (shp).
    [Show full text]
  • Improving Engine Efficiency Through Core Developments
    IMPROVING ENGINE EFFICIENCY THROUGH CORE DEVELOPMENTS Brief summary: The NASA Environmentally Responsible Aviation (ERA) Project and Fundamental Aeronautics Projects are supporting compressor and turbine research with the goal of reducing aircraft engine fuel burn and greenhouse gas emissions. The primary goals of this work are to increase aircraft propulsion system fuel efficiency for a given mission by increasing the overall pressure ratio (OPR) of the engine while maintaining or improving aerodynamic efficiency of these components. An additional area of work involves reducing the amount of cooling air required to cool the turbine blades while increasing the turbine inlet temperature. This is complicated by the fact that the cooling air is becoming hotter due to the increases in OPR. Various methods are being investigated to achieve these goals, ranging from improved compressor three-dimensional blade designs to improved turbine cooling hole shapes and methods. Finally, a complementary effort in improving the accuracy, range, and speed of computational fluid mechanics (CFD) methods is proceeding to better capture the physical mechanisms underlying all these problems, for the purpose of improving understanding and future designs. National Aeronautics and Space Administration Improving Engine Efficiency Through Core Developments Dr. James Heidmann Project Engineer for Propulsion Technology (acting) Environmentally Responsible Aviation Integrated Systems Research Program AIAA Aero Sciences Meeting January 6, 2011 www.nasa.gov NASA’s Subsonic
    [Show full text]
  • The Predicted Performance of a Two-Spool Turbofan Engine in Rainstorms
    THE PREDICTED PERFORMANCE OF A TWO-SPOOL TURBOFAN ENGINE IN RAINSTORMS By Tarik Baki Thesis presented for the degree of Masters of Science MSc to the faculty of Engineering Department of Mechanical Engineering Glasgow University April 1993. © Tarik Baki. ProQuest Number: 11007734 All rights reserved INFORMATION TO ALL USERS The quality of this reproduction is dependent upon the quality of the copy submitted. In the unlikely event that the author did not send a com plete manuscript and there are missing pages, these will be noted. Also, if material had to be removed, a note will indicate the deletion. uest ProQuest 11007734 Published by ProQuest LLC(2018). Copyright of the Dissertation is held by the Author. All rights reserved. This work is protected against unauthorized copying under Title 17, United States C ode Microform Edition © ProQuest LLC. ProQuest LLC. 789 East Eisenhower Parkway P.O. Box 1346 Ann Arbor, Ml 48106- 1346 ^1581 GLASGOW 1 UNIVERSITY LIBRARY f This thesis is dedicated to the memory of my kind and loving grand-mother Mama Hadja and my dear grand-father Hadj Tahar. You will always be present in our hearts. CONTENTS ACKNOWLEDGEMENTS NOMENCLATURE STATION NUMBERING SUMMARY PAGE CHAPTER I : GENERAL INTRODUCTION AND BACKGROUND 1.1. Introduction 4 1.1.1. Contribution made in the presentinvestigation 6 1.2. Historical development of gas turbines 7 1.3. Modem developments 9 1.4. Turbojet engine development 10 1.5. Transient behaviour of aircraft gas turbine 11 CHAPTER II : PERFORMANCE PREDICTION PROGRAM FOR TWO-SPOOL TURBOFAN ENGINE 2.1. Modelling of gas turbines 13 2.1.1.
    [Show full text]
  • Download This Issue (PDF)
    QTR_03 12 A QUARTERLY PUBLICATION BROUGHT TO YOU BY THE BOEING EDGE In-Service Supplier Support 787 Propulsion System Reducing Runway Landing Overruns New Process for Component Removal Reduction Securing Airline Information on the Ground and in the Air Cover photo: 737-800 Engine Fan Blades AERO Contents 03 E nhancing Suppliers’ In-Service Support to Airlines Airline assessments of supplier performance enhance airplane support. 05 787 Propulsion System The 787 Dreamliner is powered by new-generation engines that offer improvements in fuel consumption, 05 noise, and emissions. 15 Reducing Runway Landing Overruns A combination of procedural improvements, flight crew knowledge, and flight deck enhancements help mitigate runway overrun excursions during landing. 21 15 N ew Process for Component Removal Reduction Automating the component removal reduction process can save considerable time and help operators reduce main- tenance costs. 25 Securing Airline Information on the Ground and in the Air 21 Protecting information and technology has become an operational requirement for airlines. 25 01 WWW.BOEING.COM/BOEINGEDGE/AEROMAGAZINE I ssue 47 _Quarter 03 | 2012 AERO Publisher Design Cover photography Editorial Board Shannon Myers Methodologie Jeff Corwin Don Andersen, Gary Bartz, Richard Breuhaus, David Carbaugh, Justin Hale, Darrell Hokuf, Al John, Doug Lane, Jill Langer, Russell Lee, Duke McMillin, Editorial director Writer Printer Keith Otsuka, David Presuhn, Wade Price, Jerome Schmelzer, Jill Langer Jeff Fraga ColorGraphics Corky Townsend
    [Show full text]
  • Design of a Turbofan Engine Cycle with Afterburner for a Conceptual UAV
    Design of a Turbofan Engine Cycle with Afterburner for a Conceptual UAV BJÖRN MONTGOMERIE Mixed fl ow turbofan schematic FOI is an assignment-based authority under the Ministry of Defence. The core activities are research, method and technology development, as well as studies for the use of defence and security. The organization employs around 1350 people of whom around 950 are researchers. This makes FOI the largest research institute in Sweden. FOI provides its customers with leading expertise in a large number of fi elds such as security-policy studies and analyses in defence and security, assessment of dif- ferent types of threats, systems for control and management of crises, protection against and management of hazardous substances, IT-security an the potential of new sensors. FOI Defence Research Agency Phone: +46 8 555 030 00 www.foi.se Systems Technology Fax: +46 8 555 031 00 FOI-R-- 1835 --SE Scientifi c report Systems Technology SE-164 90 Stockholm ISSN 1650-1942 December 2005 Björn Montgomerie Design of a Turbofan Engine Cycle with Afterburner for a Conceptual UAV FOI-R--1835--SE Scientific report Systems Technology ISSN 1650-1942 December 2005 Issuing organization Report number, ISRN Report type FOI – Swedish Defence Research Agency FOI-R--1835--SE Scientific report Systems Technology Research area code SE-164 90 Stockholm 7. Mobility and space technology, incl materials Month year Project no. December 2005 E830058 Sub area code 71 Unmanned Vehicles Sub area code 2 Author/s (editor/s) Project manager Björn Montgomerie Fredrik Haglind Approved by Monica Dahlén Sponsoring agency Swedish Defense Materiel Administration (FMV) Scientifically and technically responsible Fredrik Haglind Report title Design of a Turbofan Engine Cycle with Afterburner for a Conceptual UAV Abstract A study of two turbofan engine types has been carried out.
    [Show full text]
  • STS-1000: a High Performance Turboshaft Engine for Hybrid
    AIAA 2018-2018 Engine Design Competition Sharif University of Technology STS-IDOO: A Candidate T urboshaft Engine for Hybrid Electric Medium Altitude Long Endurance Search and Rescue UAV High PowBr to WBight Low FuBI Consumption Modular 6 Compact SIGNATURE SHEET Prof. Kaveh Ghorbanian M. Reza AminiMagham Alireza Ebrahimi Faculty Advisor Project Advisor Team Leader 952166 Amir Nazemi Abolfazl Zolfaghari Hojjat Etemadianmofrad Vahid Danesh 981123 919547 964808 964807 M. Mahdi Asnaashari Saeide Kazembeigi Mahdi Jamshidiha Amirreza Saffizadeh 952842 978931 688249 937080 Copyright © 2019 by FARAS. Published by the American Institute of Aeronautics and Astronautics, Inc., with Permission Executive Summary This report proposes a turboshaft engine referred to “Sharif TurboShaft 1000 (STS-1000)” as a candidate engine to replace the baseline engine TPE331-10 for the next generation “Hybrid Electric Medium Altitude Long Endurance Search and Rescue UAV” by the year 2025. STS-1000, unlike the baseline engine, is a split single-spool turboshaft engine. The hot gas generator is a single spool with a single stage radial compressor, a reverse annular combustion chamber, and an uncooled single stage axial compressor turbine. The required shaft power is produced by a two stage axial power turbine on a separate spool which passes through the spool of the core engine and is intended to drive a power generator at the cold end of the engine. The air intake is of S-type and the exhaust duct has circular cross section. Compared to TPE331-10, STS-1000 has a higher turbine inlet temperature, a lower stage number for the air compressor, and requires less mass flow rate.
    [Show full text]
  • Pzl-10 Turboshaft Engine – System Design Review
    Journal of KONES Powertrain and Transport, Vol. 26, No. 1 2019 ISSN: 1231-4005 e-ISSN: 2354-0133 DOI: 10.2478/kones-2019-0003 PZL-10 TURBOSHAFT ENGINE – SYSTEM DESIGN REVIEW Michal Czarnecki Rzeszow University of Technology, Department of Aircraft and Aircraft Engines Powstancow Warszawy Av. 8, 35-959 Rzeszow, Poland tel.: +48 17 8651609, fax: +48 17 8543116 e-mail: [email protected] John Olsen School of Mechanical and Manufacturing Engineering The University of New South Wales, Sydney, Ainsworth Building, N.S.W. 2052, Sydney, Australia tel.: +61 2 9385 5217, fax: +61 2 9663 1222 e-mail: [email protected] Ruixian Ma School of Energy Science and Engineering Harbin Institute of Technology Harbin, 150001, Heilongjiang, China e-mail: [email protected] Abstract The PZL – 10-turboshaft gas turbine engine is straight derivative of GTD-10 turboshaft design by OKMB (Omsk Engine Design Bureau). Prototype engine first run take place in 1968. Selected engine is interested platform to modify due gas generator layout 6A+R-2, which is modern. For example axial compressor design from successful Klimov designs TB2-117 (10A-2-2) or TB3-117 (12A-2-2) become obsolete in favour to TB7-117B (5A+R-2-2). In comparison to competitive engines: Klimov TB3-117 (1974 – Mi-14/17/24), General Electric T-700 (1970 – UH60/AH64), Turbomeca Makila (1976 – H225M) the PZL-10 engine design is limited by asymmetric power turbine design layout. This layout is common to early turboshaft design such as Soloview D-25V (Mil-6 power plant). Presented article review base engine configuration (6A+R+2+1).
    [Show full text]