Bir turbojet motorunun performansının incelenmesi /
Yazar: HASAN KÜÇÜK
Danışman: DOÇ.DR. VELİ ÇELİK
İstanbul Teknik Üniversitesi / Fen Bilimleri Enstitüsü
Derece: Yüksek Lisans, 1995
Konu: Enerji = Energy
Dizin: Gaz türbini = Gas turbine ; Motorlar = Motors ; Performans analizi = Performance analysis ; Uçak motorları = Aircraft engines

Özet:
ÖZET Bu çalışmada bir turbojet motorunun ideal çevrim analizleri yapılarak elde edilen performans özellikleri ile aynı motorun deneysel performans özellikeri karşılaştırılmaktadır. Yapılan teorik analizlerde türbin giriş sıcaklığı, uçuş Mach sayısı ve kompresör basınç oranı gibi çevrim parametrelerinin, performans özellikleri olan tepki ve özgül yakıt tüketimine etkileri de incelenmiştir. İlk bölümde, gaz türbini motorları ile ilgili kısa genel bilgiler verilmektedir. İkinci bölümde, gaz türbini motorlarının gelişimi ve taşıtlarda kullanılan çeşitleri (ramjet, turbojet, turbofan ve turboprop) tanıtılmakta ve kullanım yerleri anlatılmaktadır. Üçüncü bölümde, gaz türbinlerinin çevrim analizlerinde kullanılan temel kavramlar verildikten sonra, turbojet motorunun ideal ve kayıplarla çevrim analizleri detaylı olarak anlatılmaktadır. Dördüncü bölümde, gerçek bileşen karakteristikleri ile turbojet motorunun dizayn koşulları dışındaki performansını hesaplamada kullanılan metod tanıtılmaktadır. Beşinci bölümde, deney düzenekleri hakkında bilgiler, test edilen turbojet motorunun (J79-17) teknik özellikleri ve deney sonuçları verilmektedir. Teorik çevrim analizlerinde kullanılan bilgisayar programlarının akış diyagramları da bu bölümde verilmektedir. Elde edilen teorik ve deneysel sonuçlar karşılaştırıldıktan sonra, teorik çevrim analizlerinden elde edilen sonuçlar da grafikler halinde sunulmaktadır. Teorik çevrim analizlerinden elde edilen performans değerleri ile deneysel sonuçlar arasındaki farklar kabul edilebilir sınırlar içindedir. Teorik çevrim analizlerinin sonuçlan, askeri uçaklarda kullanılan motorlar açısından değerlendirilecek olursa; türbin giriş sıcaklıkları artırılarak, daha hafif ve verimli motorların üretilebileceği söylenebilir. XII

Abstract:
SUMMARY Of the various means of producing mechanical power the turbine is in many respects the most satisfactory. The absence of reciprocating and rubbing members means that balancing problems are few, that the lubricating oil consumption is exceptionaly low, and that reliability can be high. The inherent advantages of turbine were first realized using water as the working fluid, and hydro-electric is still a significant contributor to the world's energy resources. Around the turn of the twentieth century the steam turbine began its career and, quite apart from its wide use as a marine power plant, it has become the most important prime mover for electricity generation. In spite of its successfull development, the steam turbine does have an inherent disadvantage. It is that the production of high-pressure high-temperature steam involves the installation of bulky and expensive steam generating equipment, whether it be a conventional boiler or nuclear reactor. A gas turbine is an engine designed to convert the energy of a fuel into some form of useful power, such as a mechanical (shaft) power or the high-speed thrust of a jet. A gas turbine consists basically of a gas generator section and a power conversion section. The gas generator section consists of a compressor, combustion chamber, and turbine, the turbine extracting only sufficient power to drive the compressor. This results in a high-temperature, high-pressure gas at the turbine exit. The different types of gas turbines result by adding various inlet and exit components to the gas generator. Serious development of the gas turbine began not long before the Second World War with shaft power in mind, but attantion was soon tranferred to the turbojet engine for aircraft propulsion. The gas turbine began compete successfully in other fields only in mid nineteen fifties, but since then it has made a progressively greater impact in an increasing variety of applications. Today, the gas turbine is widely used. During the period from the granting of the first patent in 1791 to the present, the gas turbine has been developed into a very reliable, versatile engine with a high power-to-weight ratio. It is used exclusively to power all new commercial airplanes. Most business aircraft are powered by one form or another of a gas turbine. The gas turbine has been used to power boats and trains, has been widely accepted for electric power generators and gas pipeline compressor drives, and is being tested for use in buses and trucks. It has been tested extensively as a power plant for an otomobile. In order to produce an expansion through a turbine a pressure ratio must be provided, and the first necessary step in the cycle of a gas turbine plant therefore be xiiicompression of the working fluid. If after compression the working fluid were to be expanded directly in the turbine, and there were no losses in either component, the power developed by the turbine would just equal that absorbed by the compressor. Thus if the two were coupled together the combination would do no more than turn itself round. But the power developed by the turbine can be increased by the addition of energy to raise the temperature of the working fluid prior to expansion. When the working fluid is air a very suitable means of doing this is by combustion of fuel in the air which has been compressed. Expansion of the hot working fluid then produces a greater power output from the turbine, so that it is able to provide a useful output in addition to the power necessary to drive the compressor. This represent the gas turbine or internal-combustion turbine in its simplest form. The three main components are a compressor, combustion chamber and turbine, connected together. In earliest days of the gas turbine, two possible systems of combustion were proposed: one at constant pressure, the other at constant volume. Theoretically, the thermal efficiency of the constant volume cycle is higher than that of the constant pressure cycle, but the mechanical difficulties are very much greater. With heat addition at constant volume, valves are necessary to isolate the combustion chamber from the compressor and turbine. Combustion is therefore intermittent, which impairs the smooth running of the machine. It is difficult to design a turbine to operate efficiently under such conditions. In the constant pressure gas turbine, combustion is a continious process in which valves are unnecessary and it was soon accepted that the constant pressure cycle had the greater possibilities for future development. It is important to realise that in the gas turbine the processes of compression, combustion and expansion do not occur in a single component as they do in a reciprocating engine. They occur in components which are separate in the sense that they can be designed, tested and developed individually, and these components can be linked together to form a gas turbine unit in a variety of ways. The possible number of components is not limited to the three already mentioned. Other compressors and turbines can be added, with intercoolers between the compressors, and reheat combustion chambers between the turbines. A heat-exchanger which uses some of the energy in the turbine exhaust gas to preheat the air entering the combustion chamber may also be introduced. These refinements may be used to increase the power output and efficiency of the plant at the expense of added complexity, weight and cost. Two main factors affecting the performance of gas turbines: component efficiencies and turbine working temperature. The higher they can be made, the better all-round performance of the plant. The overall efficieny of the gas turbine cycle depends also upon the pressure ratio of the compressor. The difficulty of obtaining a sufficently high pressure ratio with an adequate compressor efficiency was not resolved until the science of aerodynamics could be applied to the problem. The development of the gas turbine has gone hand in hand with the development of this science, and that of metallurgy, with the result that is now possible to find advanced engines using pressure ratios up to 35:1, component efficiencies of 85-90 per cent, and turbine inlet temperature up to 1650 K. XIVShort explanations on gas turbine engines are given in the first part of the study. In the second part, the improvement of the gas turbine engines (ramjet, turbojet, turbofan and tuboprop) used in aircrafts are introduced and their applications are given. In the third part, after introducing basic concepts which are used in analyses, ideal cycle analyses of turbojet engine and cycle analysis with losses are given in details. In the fourth part, the method used in the off-design performance analysis of turbojet engine by using real component characteristics is introduced. In the fifth part, the knowledge on test cells, technical data and experimental results of the J79-I7 engine are given. Flow diagrams of computer programs used in theoretical cycle analyses are also given in this part. After comparing experimental and theoretical results, results of theoretical analyses are shown in the form of graphs. In the last part, obtained results are summerised and the comments about these results are presented. In this study, the differences between engine performances which are supplied from theoretical cycle analysis and experimental studies are in the tolerable range. If theoretical analyses results are dealed with respect to the military aircraft engines, it is seen that lighter and more efficient engines can be designed by incraesing turbine inlet temperature. xxcompression of the working fluid. If after compression the working fluid were to be expanded directly in the turbine, and there were no losses in either component, the power developed by the turbine would just equal that absorbed by the compressor. Thus if the two were coupled together the combination would do no more than turn itself round. But the power developed by the turbine can be increased by the addition of energy to raise the temperature of the working fluid prior to expansion. When the working fluid is air a very suitable means of doing this is by combustion of fuel in the air which has been compressed. Expansion of the hot working fluid then produces a greater power output from the turbine, so that it is able to provide a useful output in addition to the power necessary to drive the compressor. This represent the gas turbine or internal-combustion turbine in its simplest form. The three main components are a compressor, combustion chamber and turbine, connected together. In earliest days of the gas turbine, two possible systems of combustion were proposed: one at constant pressure, the other at constant volume. Theoretically, the thermal efficiency of the constant volume cycle is higher than that of the constant pressure cycle, but the mechanical difficulties are very much greater. With heat addition at constant volume, valves are necessary to isolate the combustion chamber from the compressor and turbine. Combustion is therefore intermittent, which impairs the smooth running of the machine. It is difficult to design a turbine to operate efficiently under such conditions. In the constant pressure gas turbine, combustion is a continious process in which valves are unnecessary and it was soon accepted that the constant pressure cycle had the greater possibilities for future development. It is important to realise that in the gas turbine the processes of compression, combustion and expansion do not occur in a single component as they do in a reciprocating engine. They occur in components which are separate in the sense that they can be designed, tested and developed individually, and these components can be linked together to form a gas turbine unit in a variety of ways. The possible number of components is not limited to the three already mentioned. Other compressors and turbines can be added, with intercoolers between the compressors, and reheat combustion chambers between the turbines. A heat-exchanger which uses some of the energy in the turbine exhaust gas to preheat the air entering the combustion chamber may also be introduced. These refinements may be used to increase the power output and efficiency of the plant at the expense of added complexity, weight and cost. Two main factors affecting the performance of gas turbines: component efficiencies and turbine working temperature. The higher they can be made, the better all-round performance of the plant. The overall efficieny of the gas turbine cycle depends also upon the pressure ratio of the compressor. The difficulty of obtaining a sufficently high pressure ratio with an adequate compressor efficiency was not resolved until the science of aerodynamics could be applied to the problem. The development of the gas turbine has gone hand in hand with the development of this science, and that of metallurgy, with the result that is now possible to find advanced engines using pressure ratios up to 35:1, component efficiencies of 85-90 per cent, and turbine inlet temperature up to 1650 K. XIVAn aircraft gas turbine engine represents a complex system comprising a number of rotating and stationary components, each with its own performance characteristics. The performance of the complete engine depends the performance characteristics of the individual components or sub systems and matching. Moreover a wide range of flight conditions. Hence, testing a complete engine to explore its full performance envelope can be both very costly and extremely laborious. Therefore, simulating the operation of the aero gas turbine engine on a computer has obvious advantages. Cosequently, a number of papers have been published on this subject in the open literature. In the gas turbine field, computer programs have been written for design and cycle analysis, fuel control stability analysis, component matching at off-design, reliability evaluations, stress analysis, aircraft engine mission optimization, component design - the list is endless especially when each of these general subjects heads up a family of more specialized programs. Performance computer programs serve different purposes during the engine development as the information available varies due to the stage of development. For example, in the earlier phases, preliminary cycle selection is based on estimated component performance, while in the latter stages component performance is calculated based on detailed component characteristics. This difference provide a basis which is commonly used to divide engine performance programs into distinct classifications: 1. 'Design Point" programs, used in cycle and configuration selection in which the inputs are the selected values of compressor efficiency, turbine efficiency, compressor pressure ratio, etc. In most cases these this programs calculate the turbine pressure ratios required to satisfy the horsepower balance requirements on the various shafts, and calculate the turbine areas, nozzle areas, etc., as part of the output in addition to the power plant thrust, SFC, etc. 2. 'Off-Design" or Match-Point programs which are much closer to real gas turbine systems since the input includes tables representing the estimated, or tested, compressor and turbine characteristics. At each engine operating point, the performance of the components varies as a result of various relationships established by the continuity requirements, horsepower balances, and control functions. In this study, the theoretical anlyses results of the turbojet engine (J79-17) is compared with the experimental results. Tests were conducted on closed test cell unit in Eskişehir Airforce and Maintanence Factory. Ideal cycle analyses of turbojet engine are studied according to flight and blade Mach numbers. Also cycle analysis of turbojet engine with losses are made by using constant component efficiencies. In theoretical analysis, the effect of the cycle parameters like turbine inlet temperature, compressor pressure ratio and flight Mach number are examined on the performance parameters, thrust, specific thrust and specific fuel consumption. In ideal cycle analysis, the change of the compressor pressure ratio which gives maximum thrust is supplied with change of flight Mach number and turbine inlet temperature. XIXShort explanations on gas turbine engines are given in the first part of the study. In the second part, the improvement of the gas turbine engines (ramjet, turbojet, turbofan and tuboprop) used in aircrafts are introduced and their applications are given. In the third part, after introducing basic concepts which are used in analyses, ideal cycle analyses of turbojet engine and cycle analysis with losses are given in details. In the fourth part, the method used in the off-design performance analysis of turbojet engine by using real component characteristics is introduced. In the fifth part, the knowledge on test cells, technical data and experimental results of the J79-I7 engine are given. Flow diagrams of computer programs used in theoretical cycle analyses are also given in this part. After comparing experimental and theoretical results, results of theoretical analyses are shown in the form of graphs. In the last part, obtained results are summerised and the comments about these results are presented. In this study, the differences between engine performances which are supplied from theoretical cycle analysis and experimental studies are in the tolerable range. If theoretical analyses results are dealed with respect to the military aircraft engines, it is seen that lighter and more efficient engines can be designed by incraesing turbine inlet temperature. xxAn aircraft gas turbine engine represents a complex system comprising a number of rotating and stationary components, each with its own performance characteristics. The performance of the complete engine depends the performance characteristics of the individual components or sub systems and matching. Moreover a wide range of flight conditions. Hence, testing a complete engine to explore its full performance envelope can be both very costly and extremely laborious. Therefore, simulating the operation of the aero gas turbine engine on a computer has obvious advantages. Cosequently, a number of papers have been published on this subject in the open literature. In the gas turbine field, computer programs have been written for design and cycle analysis, fuel control stability analysis, component matching at off-design, reliability evaluations, stress analysis, aircraft engine mission optimization, component design - the list is endless especially when each of these general subjects heads up a family of more specialized programs. Performance computer programs serve different purposes during the engine development as the information available varies due to the stage of development. For example, in the earlier phases, preliminary cycle selection is based on estimated component performance, while in the latter stages component performance is calculated based on detailed component characteristics. This difference provide a basis which is commonly used to divide engine performance programs into distinct classifications: 1. 'Design Point" programs, used in cycle and configuration selection in which the inputs are the selected values of compressor efficiency, turbine efficiency, compressor pressure ratio, etc. In most cases these this programs calculate the turbine pressure ratios required to satisfy the horsepower balance requirements on the various shafts, and calculate the turbine areas, nozzle areas, etc., as part of the output in addition to the power plant thrust, SFC, etc. 2. 'Off-Design" or Match-Point programs which are much closer to real gas turbine systems since the input includes tables representing the estimated, or tested, compressor and turbine characteristics. At each engine operating point, the performance of the components varies as a result of various relationships established by the continuity requirements, horsepower balances, and control functions. In this study, the theoretical anlyses results of the turbojet engine (J79-17) is compared with the experimental results. Tests were conducted on closed test cell unit in Eskişehir Airforce and Maintanence Factory. Ideal cycle analyses of turbojet engine are studied according to flight and blade Mach numbers. Also cycle analysis of turbojet engine with losses are made by using constant component efficiencies. In theoretical analysis, the effect of the cycle parameters like turbine inlet temperature, compressor pressure ratio and flight Mach number are examined on the performance parameters, thrust, specific thrust and specific fuel consumption. In ideal cycle analysis, the change of the compressor pressure ratio which gives maximum thrust is supplied with change of flight Mach number and turbine inlet temperature. XIXShort explanations on gas turbine engines are given in the first part of the study. In the second part, the improvement of the gas turbine engines (ramjet, turbojet, turbofan and tuboprop) used in aircrafts are introduced and their applications are given. In the third part, after introducing basic concepts which are used in analyses, ideal cycle analyses of turbojet engine and cycle analysis with losses are given in details. In the fourth part, the method used in the off-design performance analysis of turbojet engine by using real component characteristics is introduced. In the fifth part, the knowledge on test cells, technical data and experimental results of the J79-I7 engine are given. Flow diagrams of computer programs used in theoretical cycle analyses are also given in this part. After comparing experimental and theoretical results, results of theoretical analyses are shown in the form of graphs. In the last part, obtained results are summerised and the comments about these results are presented. In this study, the differences between engine performances which are supplied from theoretical cycle analysis and experimental studies are in the tolerable range. If theoretical analyses results are dealed with respect to the military aircraft engines, it is seen that lighter and more efficient engines can be designed by incraesing turbine inlet temperature. xx