Academic literature on the topic 'Allison T56 (Engine)'
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Journal articles on the topic "Allison T56 (Engine)"
1
Kim, J., M. G. Dunn, A. J. Baran, D. P. Wade, and E. L. Tremba. "Deposition of Volcanic Materials in the Hot Sections of Two Gas Turbine Engines." Journal of Engineering for Gas Turbines and Power 115, no. 3 (July 1, 1993): 641–51. http://dx.doi.org/10.1115/1.2906754.
Full textAbstract:
This paper reports the results of a series of tests designed to determine the melting and subsequent deposition behavior of volcanic ash cloud materials in modern gas turbine engine combustors and high-pressure turbine vanes. The specific materials tested were Mt. St. Helens ash and a soil blend containing volcanic ash (black scoria) from Twin Mountain, NM. Hot section test systems were built using actual engine combustors, fuel nozzles, ignitors, and high-pressure turbine vanes from an Allison T56 engine can-type combustor and a more modern Pratt and Whitney F-100 engine annular-type combustor. A rather large turbine inlet temperature range can be achieved using these two combustors. The deposition behavior of volcanic materials as well as some of the parameters that govern whether or not these volcanic ash materials melt and are subsequently deposited are discussed.
2
MacLeod, J. D., and J. C. G. Laflamme. "Compressor Coating Effects on Gas Turbine Engine Performance." Journal of Engineering for Gas Turbines and Power 113, no. 4 (October 1, 1991): 530–34. http://dx.doi.org/10.1115/1.2906273.
Full textAbstract:
In an attempt to increase the time between maintenance actions and to improve performance retention of turboprop engines installed in transport and maritime patrol aircraft, the Canadian Department of National Defence is evaluating an erosion and corrosion-resistant blade coating, for use on compressors. As coatings could appreciably alter engine performance by virtue of their application thickness and surface quality, the National Research Council of Canada was asked to quantify any performance changes that could occur. A project was initiated, utilizing a new Allison T56 turboprop engine, to assess not only the performance changes resulting from the coating, but also those from dismantling and reassembling the compressor, since the compressor must be completely disassembled to apply the coating. This paper describes the project objectives, the experimental installation, and the measured effects of the coating application on compressor performance. Performance variations due to compressor rebuilds on both engine and compressor characteristics are discussed. As the performance changes were small, a rigorous measurement uncertainty analysis is included. The coating application process and the affected overhaul procedures are examined. The results of the pre- and postcoating compressor testing are presented, with a discussion of the impact on engine performance.
3
MacLeod, J. D., V. Taylor, and J. C. G. Laflamme. "Implanted Component Faults and Their Effects on Gas Turbine Engine Performance." Journal of Engineering for Gas Turbines and Power 114, no. 2 (April 1, 1992): 174–79. http://dx.doi.org/10.1115/1.2906567.
Full textAbstract:
Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effect is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library, which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: (a) first-stage turbine nozzle erosion damage; (b) first-stage turbine rotor blade untwist; (c) compressor seal wear; (d) first and second-stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.
Dissertations / Theses on the topic "Allison T56 (Engine)"
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Skidmore, F. W., and n/a. "The influence of gas turbine combustor fluid mechanics on smoke emissions." Swinburne University of Technology, 1988. http://adt.lib.swin.edu.au./public/adt-VSWT20070420.131227.
Full textAbstract:
This thesis describes an experimental program covering the development of certain
simple combustion chamber modifications to alleviate smoke emissions from the
Allison T56 turboprop engines operated by the Royal Australian Air Force.
The work includes a literature survey, smoke emission tests on two variants
of the T56 engine, flow visualisation studies of the combustion system in a
water tunnel and combustion rig tests of a standard combustor and four possible
modifications. The rig tests showed that reductions in smoke emissions of
80% were possible by simple modifications that reduced the primary zone
equivalence ratio and improved mixing in that zone.
Conference papers on the topic "Allison T56 (Engine)"
1
Laughlin, T. P., and Joseph Toth. "T56 Derivative Engine in the Improved E-2C." In ASME 1985 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-gt-176.
Full textAbstract:
Airborne Early Warning for the Navy fleets has been provided for the past 20 years by the Grumman/Allison E-2 Airframe/T56 engine combination. Although avionic capability has been continually updated to meet the increased threat, the airframe and powerplant have seen only minor changes. Projected mission requirements and future avionic system enhancements require payload increases being limited by the power capability of the present T56 powerplant. Of paramount importance in the E-2 carrier deck operation is the single engine rate of climb capability of the aircraft. This paper discusses the logical evolution of a replacement engine for the E-2C — a derivative T56 engine contracted and designated by the Navy as the T56-A-427 — to meet the projected single engine takeoff and other mission requirements. The T56-A-427 provides 24% power and 13% fuel consumption improvements with identical installation interfaces, and substantially improves E-2C performance characteristics across the flight envelope. Furthermore, the paper shows that meeting these stringent performance requirements with a derivative engine results in a low risk development program and an engine with improved maintainability and reliability, which can capitalize on the in-place logistics support base of the T56 — the world’s longest production run gas turbine engine.
2
Guy, S. R. D., W. D. E. Allan, Marc LaViolette, and P. R. Underhill. "Optical Patternation of Gas Turbine Fuel Sprays During Simulated Engine Operating Conditions." In ASME Turbo Expo 2009: Power for Land, Sea, and Air. ASMEDC, 2009. http://dx.doi.org/10.1115/gt2009-59011.
Full textAbstract:
Fuel atomizer condition can have a significant impact on gas turbine hot section component life. In order to investigate the depth of this influence, an experimental test apparatus was constructed, which allowed for optical access to the primary zone of a Rolls-Royce/Allison T56–A–15 turboprop combustion chamber. Test conditions were matched to simulate altitude cruise conditions of a C–130H Hercules military transport aircraft. T56 fuel nozzles of various conditions were tested in free air and then in the test rig using optical patternation techniques. Results indicated that spray characteristics observed in quiescent ambient air persisted under the representative engine operating conditions both burning and non-burning. The optical patternation tests also revealed the influence of combustion liner airflow patterns on the spray within the region of the primary zone that was observed. Conclusions were drawn such as the persistence of spray features observed in open air testing when nozzles were tested at engine representative conditions and recommendations were made for future experimentation.
3
Kim, J., M. G. Dunn, A. J. Baran, D. P. Wade, and E. L. Tremba. "Deposition of Volcanic Materials in the Hot Sections of Two Gas Turbine Engines." In ASME 1992 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1992. http://dx.doi.org/10.1115/92-gt-219.
Full textAbstract:
This paper reports the results of a series of tests designed to determine the melting and subsequent deposition behavior of volcanic ash cloud materials in modern gas turbine engine combustors and high pressure turbine vanes. The specific materials tested were Mt. St. Helens ash and a soil blend containing volcanic ash (black scoria) from Twin Mountain, New Mexico. Hot section test systems were built using actual engine combustors, fuel nozzles, ignitors, and high pressure turbine vanes from an Allison T56 engine can-type combustor and a more modern Pratt and Whitney F-100 engine annular-type combustor. A rather large turbine inlet temperature range can be achieved using these two combustors. The deposition behavior of volcanic materials as well as some of the parameters that govern whether or not these volcanic ash materials melt and subsequently deposit are discussed.
4
MacLeod, J. D., V. Taylor, and J. C. G. Laflamme. "Implanted Component Faults and Their Effects on Gas Turbine Engine Performance." In ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1991. http://dx.doi.org/10.1115/91-gt-041.
Full textAbstract:
Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effort is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: a) 1st stage turbine nozzle erosion damage, b) 1st stage turbine rotor blade untwist, c) compressor seal wear, d) 1st and 2nd stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.
5
Liburdi, J., D. R. Nagy, and V. R. Parameswaran. "Erosion Resistant Titanium Nitride Coating for Turbine Compressor Applications." In ASME 1992 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1992. http://dx.doi.org/10.1115/92-gt-417.
Full textAbstract:
While operating in dusty environments, the blades and vanes in turbine compressors are prone to degradation by solid particle erosion which causes surface roughening as well as loss of airfoil contour and changes in airfoil geometry. This results in decreased compressor performance, higher specific fuel consumption, and significantly increased operational costs. Erosion damage is more prominent in flight engines that cannot be protected by inlet filters. This paper describes the development and application of a thin ceramic titanium nitride coating to improve the erosion resistance of compressor airfoils. The coatings were produced by a Reactive Ion Coating (RIC) process and optimized to produce a very adherent erosion resistant coating structure. The coating process was successfully scaled up and applied to a complete Allison T56 compressor for engine test. The control laboratory tests showed that the thin coating had no significant influence on either the resonance frequency or the fatigue resistance of the blades and the instrumented engine tests confirmed that the performance was typical of overhauled engines. Therefore, titanium nitride coatings are suitable for service and can be retrofitted on existing engines to improve the life of the compressors.
6
MacLeod, J. D., and J. C. G. Laflamme. "Influence of a Thermal Barrier Coating on the Performance of a Turboprop Engine." In ASME 1992 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1992. http://dx.doi.org/10.1115/92-gt-038.
Full textAbstract:
Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada has evaluated the influence of applying a thermal barrier coating on the performance of a gas turbine engine. The effort is aimed at quantifying the performance effects of a particular ceramic coating on the first stage turbine vanes. The long term objective of the program is to both assess the relative change in engine performance and compare against the claimed benefits of higher possible turbine inlet temperatures, longer time in service and increased time between overhauls. The engine used for this evaluation was the Allison T56 turboprop with the first stage turbine nozzles coated with the Chromalloy RT-33 ceramic coating. The issues addressed in testing this particular type of hot section coating were; 1) effect of coating thickness on nozzle effective flow area; 2) surface roughness influence on turbine efficiency; This paper describes the project objectives, the experimental installation, and the results of the performance evaluations. Discussed are performance variations due to coating thickness and surface roughness on engine performance characteristics. As the performance changes were small, a rigorous measurement uncertainty analysis is included. The coating application process, and the affected overhaul procedures are examined. The results of the pre- and post-coating turbine testing are presented, with a discussion of the impact on engine performance.
7
Benson, J., W. Miglietti, S. J. Glass, and K. Getliffe. "Problems in Military Propulsion Engines of the South African Air Force." In ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-425.
Full textAbstract:
The Nimonic 105 first stage turbine blade of the Rolls Royce Viper 22/1 engine has a design life of 2400 hrs. Over the years, these blades have prematurely failed (1 200–1 800 hrs), necessitating a service limit of 1 000 hrs. Initial investigations thought the cause of failure to be linked with higher than desired lead content but the final cause of failure was attributed to an incorrect heat treatment, forming continuous grain boundary carbides, resulting in embrittlement of the blade. A rejuvenation heat treatment was then developed. The IN-713 and IN-X-40 first stage vanes of the Allison T56 engine suffer thermal fatigue cracking. The IN-713 vanes suffer narrow, fine but long cracks, whereas the X-40 vanes suffer severe cracking (cracks as wide as 3 mm) and excessive oxidation due to overheating. The X-40 alloy is protected by a Cr2O3 scale compared with Al3O3 on the IN-713 alloy. Above 1 000 °C the Cr2O3 breaks down and excessive oxidation results. A braze repair was developed for the X-40 vanes.
8
MacLeod, J. D., and B. Drbanski. "Turbine Rebuild Effects on Gas Turbine Performance." In ASME 1992 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1992. http://dx.doi.org/10.1115/92-gt-023.
Full textAbstract:
The Engine Laboratory of the National Research Council of Canada (NRCC), with the assistance of Standard Aero Ltd., has established a program for the evaluation of component deterioration on gas turbine engine performance. As part of this project, a study of the effects of turbine rebuild tolerances on overall engine performance was undertaken. This study investigated the range of performance changes that might be expected for simply disassembling and reassembling the turbine module of a gas turbine engine, and how these changes would influence the results of the component fault implantation program. To evaluate the effects of rebuilding the turbine on the performance of a single spool engine, such as Allison T56 turboprop engine, a series of three rebuilds were carried out. This study was performed in a similar way to a previous NRCC study on the effects of compressor rebuilding. While the compressor rebuild study had found performance changes in the order of 1% on various engine parameters, the effects of rebuilding the turbine have proven to be even more significant. Based on the results of the turbine rebuild study, new methods to improve the assurance of the best possible tolerances during the rebuild process are currently being addressed. This paper describes the project objectives, the experimental installation, and the results of the performance evaluations. Discussed are performance variations due to turbine rebuilds on engine performance characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.
9
MacLeod, J. D., and B. Barry. "Combustor Rebuild Effects on Gas Turbine Performance." In ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-160.
Full textAbstract:
The Institute for Aerospace Research of the National Research Council of Canada (NRCC), has established a program for the evaluation of component deterioration on gas turbine engine performance. The effort is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. As part of this project, a study of the effects of combustor rebuild tolerances on overall engine performance was undertaken. This study investigated the range of performance changes that might be expected for simply disassembling and reassembling the combustor module of a gas turbine engine, and how these changes would influence the results of any component modification testing. To evaluate the effects of rebuilding the combustor on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of three rebuilds was carried out. This study was performed in a similar way to two previous NRCC studies on the effects of compressor and turbine rebuilding. While the compressor and turbine rebuild studies found performance changes in the order of I% on various engine parameters, the effects of rebuilding the combustor have proven to be of similar magnitude. Based on the results of the combustor rebuild study, new methods to improve the assurance of the best possible tolerances during the rebuild process are currently being addressed. This paper describes the project objectives, the experimental installation, and the results of the performance evaluations. Discussed are performance variations due to combustor rebuilds on engine performance characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.
10
Kotzer, Clayton, Marc LaViolette, and William Allan. "Effects of Combustion Chamber Geometry Upon Exit Temperature Profiles." In ASME Turbo Expo 2009: Power for Land, Sea, and Air. ASMEDC, 2009. http://dx.doi.org/10.1115/gt2009-60156.
Full textAbstract:
The purpose of this research was to investigate the effects of combustion chamber geometry on exit temperature fields using an ambient pressure test rig. The apparatus contained a 120° sector of a combustion section of a Rolls Royce (previously Allison) T56-A-15 gas turbine engine. A thermocouple rake acquired high-resolution temperature measurements in the combustion chamber exit plane. Rig test conditions were set to simulate an engine operating condition of 463 km/h (250 knots) at 7620 m (25000ft) by matching the Mach number, the equivalence ratio and the Sauter mean diameter of the fuel spray. To quantify the geometric deviations of the combustion chamber specimens, which varied in service conditions, a three-dimensional laser scanning system was used. Combustion chamber geometric deviations were extracted through comparison of the scanned data to a reference model using the selected software. The relationship between combustion chamber exit temperature profile and geometric deviation was then compared. The main conclusion of this research was that small deviations from nominal dimensions in the dilution zone of the combustion chamber correlated to an increase in pattern factor. A decrease in the mixing of the products of combustion and dilution air was observed as damage in the dilution zone increased. This reduction in mixing created a more compact, higher temperature core flow. The results obtained from this research were compared to past studies.