Project report power plant

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1 Group 2A2O Amsterdam, January, 2010

2 Preface Project report power plant This project report is made by project group 2A2O as part of the foundation course of the education Aviation Studies at the Hogeschool van Amsterdam. The group has worked seven weeks on this fourth project Power plant. During this project we learned much about the operation of the types of engines and especially the turbofan engine. We would like to thank our project teacher Victor Laban for his help during this project. Team 2A2O, 2010 Group 2A2O Amsterdam, January, 2010

3 Contents Project report power plant Summary... 1 Introduction Gas turbine research General gas turbine theory Types of engines Purpose of a gas turbine Operation of a gas turbine Indication Subsystems Gearbox Hydraulic system Lubrication system Electric systems Bleed air Gas turbine theory Purpose Inlet Compressor Combustion Turbine a Components Exhaust Nozzle Engine Properties Performance Environment Demands Client Demands Demands Authorities Next generation engines General gas turbine innovations Material innovations Function research Analysis gas turbine parts Morphological overview Group 2A2O Amsterdam, January, 2010

4 2.1.1 Air intake Fan blades Bypass Compressor Combustion Turbine Exhaust nozzle Possible designs Design One: Noise reduced engine Design Two: Low fuel consumption engine Design three: Low emission engine Engine performance Advantages and disadvantages The three designs Advantages and disadvantages overview Conclusion Engine design HVA CS-25E engine The engine specified Sub systems specified HVA CS-25E layout Certification Type certificate Airworthiness certificate Maintenance research High and low maintenance parts Engine inspection Financial aspects Design costs Maintenance costs Benefits Breakeven point Final conclusion Recommendation Reference list Group 2A2O Amsterdam, January, 2010

5 Summary The purpose of this project is to complete a concept design of a new engine for Fokker Aircraft (FA). The report is based on the current Rolls Royce Tay RB MK turbine engine of the Fokker 100. FA has requested this report to develop a new concept engine for the Fokker 100. This concept design must be more fuel efficient and reduce environmental pollution in relation to the current engine. The new devised engine will be used for the current and next generation Fokker 100 s. Different types of engines can be applied on an aircraft. A Fokker 100 utilizes a turbojet engine. Besides the engine, there are several subsystems that making the operation of the aircraft more manageable such as the gearbox, hydraulic system and electric systems. The gas turbine engine is a heat engine using air as a working fluid to provide thrust. The air enters the engine in the inlet. The inlet provides the compressor of air and decreases the velocity of the incoming airflow to reach the desirable axial entrance speed for the compressor. The air will now be compressed in axial lengths by the compressor. The compressor is a multi-stage unit that minimizes the losses in air. The two main components of the compressor are the rotor blades and stator vanes. The air then goes into the combustion chamber. The purpose of the combustion chamber is to convert chemical energy that is located in the added fuel, in heat. The exerted product of the combustion chamber is a heated gas flow. A combustion chamber is subdivided in a primary and dilution zone which ensures that the air will be cooled to an acceptable temperature for the turbine. The turbine has the task of providing the power to drive the compressor and other accessories mounted on the gearbox of the engine. By propelling an amount of air backwards the reaction power forces the airplane forward. The final stage of the gas turbine where the gas exits is called the exhaust nozzle. The exhaust accelerates the gasses to the required velocity and direction before exiting the gas turbine. For safe flight operations of the engine during different flight phases, the International Civil Aviation Organization (ICAO) and the European Aviation Safety Agency (EASA) have demands which recorded in CS-25, CS-E and CS-34. The client has stated that the new concept design must have a minimal thrust of Newton and a minimal range of 1500 nautical miles. A short research on innovations concerning the gas turbine branch is done to enhance knowledge of today s technologies and materials. To create an engine design each stage is examined and the additional parts are described. Out of these possibilities a morphological overview is made and three concept designs were chosen out of this. Each of the three engines has a special property. First is the noise engine type, which main focus is to reduce noise nuisance. The second type is the fuel efficient type, which has a good fuel efficiency. The final engine has the main purpose of keeping gas emission as low as possible. The three concept designs vary in various components which gives each design its special property. In the process of designing an engine, a full analysis of an engine is needed. With an analysis of an engine there are several calculations to be made. Such as the thrust, range, endurance, thrust specific fuel consumption and the efficiencies of all the stages of the engine. The calculations are the basis for the pro and con overview, which will conclude with the best design. The final design will be fully specified. In this specification, materials and properties of each component and the additional subsystems are explained. This concept engine will be certified and therefore the process of certification is described. Each engine must undergo regular maintenance. Some components need extra maintenance in order to maintain the reliability of the engine. These components will be given per stage. Engine Incorporated B.V. has designed an engine, HVA CS-25E, that is improved in thrust and has a reduced emission. For Fokker 100 of nowadays the engine is not recommended because the break-even point is at nineteen years. For the Fokker 100 of the next generation this design is recommended because it makes operation at 9000 ft on hot day conditions more efficient. Group 2A2O Amsterdam, January,

6 Introduction Fokker Aircraft wants to have new engines for the next generation Fokker 100 and for the Fokker 100 of nowadays. Projectgroup 2A2O has got the assignment to design an engine that is more fuel efficient, less noisy and have less emission. Because Fokker Aircraft operates in South-America, the performance on higher elevations have to be taken in account. The process of designing the engine will be written in a report which contains a minimum of 25 pages and a maximum of 40 pages. Besides this report, an appendix document and a calculation document are included. The report will be written concerning Wentzel (2008) and Siers (2004). Before designing, some basic knowledge about the engine is necessary. First research is done about engines in general. In this paragraph the different types of engines and the workings in basic will be explained. After that, the subsystems of the engine are investigated. The engine has a gearbox with several accessories and the engine uses bleed air to power pneumatic systems. Also, some of the subsystems use electricity to be powered. With this knowledge of the subsystems, the main components and their function of the gas turbine can be researched. These functions are: inlet, compression, combustion, turbine and exhaust. With these functions propulsion is created. In this report the thrust reverses will not be taken in account. This is because these reverses are part of the aircraft and not part of the engine. After discussed these functions, calculations can be made concerning these functions. Formulas will be given about temperatures and efficiencies. Another subject to discuss is the laws and requirements. The design of the engine has to meet with these laws. Also research has to be done concerning the requirement of the different parties, such as; Fokker Aircraft, maintenance personnel, airlines and environment. After that it is needed to know what the innovations are in engine design. After this research a functionally research will be done. (1) In the second part of this report, a morphological overview is made. First every stage is investigated with their several options. These options will be put in a morphological overview. After that, three lines will be drawn, with one line being the most fuel efficient option, one line will be the option with the lowest noise and the last line will be the line with the less emission. These are the three different design possibilities. Research will be done in the advantages and disadvantages of every design option. This will be done with the help of some calculations of every option. After that, a design options can be chosen. (2) When chosen a design, in the third part of the report the design will be developed. In this part, amongst other things, the subsystems that will be installed and the amount of stages of the compressor and turbine will be discussed. This design has to be certified. The engine has to have a type certificate and the Fokker 100 has to be recertified its airworthiness. With this certification a maintenance program is needed. In this maintenance program research will be done concerning the parts that need high maintenance and the parts that need low maintenance. After that, calculations will be made about the costs. It is needed to know what the design costs are and what the maintenance costs will be. One of the benefits is that less fuel is needed for the new designed engine. At last a conclusion can be drawn and a recommendation can be given. (3) The main resources that are used during this project are: Fokker 100 Aircraft Operation Manual and the Gas Turbine Engineering Handbook second edition. The full bibliography can be found on page 42 In the appendix document the project assignment can be found (appendix I). Also a process report is written about the improvements of the group and every group member individual (appendix II). The group has made some group agreements that had to be taken in account during this project (appendix III). Finally, for every abbreviation in this report the abbreviation list can be used (appendix IV). Group 2A2O Amsterdam, January,

7 1 Gas turbine research Fokker aircraft wants a new gas turbine engine which fits on the classic Fokker 100 aircraft and next generation Fokker 100 aircraft. For engineering an engine and for better understanding of the aircraft, an engine analysis must be performed. This research will contain general gas turbine theory (1.1). The next step is to analyse the subsystems which are involved with the engine (1.2). After this a thorough examination of the Rolls Royce Tay RB MK turbine engine of the Fokker 100 will be done to obtain knowledge about the stages in the engine (1.3). Thereupon the theory, equations and calculation methods used on the calculation of engine performance shall be discussed (1.4). The engine which needs to be designed has to meet several regulations and demands when the engine needs to be certified (1.5). For designing an engine the latest improvements and inventions are studied (1.6). The processes from air intake to engine control of the engine will be described in the functionary research (1.7). The main sources used in this chapter are Rolls Royce (1996) and Hüncke (2003). 1.1 General gas turbine theory A gas turbine is a power plant, which produces thrust, electricity, hydraulic pressure and pneumatic pressure. Gas turbines come in many forms, such as the turbojet, turboprop, propfan and turbofan. With such a variety of engine configurations, a new engine cannot be designed without research of these engines (1.1.1). The actual purpose of the engine should also be clear (1.1.2). For the design of the engine knowledge of the main operation of a gas turbine engine is needed (1.1.3). To operate the engine inside the cockpit, operating panels, handles and indication screens are available (1.1.4) Types of engines There are five types of engines which can be used in the aviation industry: the turbojet engine (1.1.1a), the turboprop engine (1.1.1b), the prop-fan engine (1.1.1c), the unducted fan engine (1.1.1d) and the turbofan engine (1.1.1e). These engines each have their different specifications and the use of a specific engine is based on the location where the aircraft has to fly. The turbo shaft engine is also a gas turbine engine but it is used for helicopters so this engine will not be discussed a Turbojet engine The first and simplest type of gas turbine engine is the turbojet (appendix V). Turbojet engines are fuel inefficient if flown at high subsonic and transonic speeds and are noisy. The turbojet engines are still used in military aircraft, due to their high exhaust speed, low frontal area and relative simplicity. These engines were also used because they were able to achieve very high altitudes and speeds, much higher than propeller engines. This is because of their better compression ratio and because of their high exhaust speed. However, because of their fuel inefficiency and noise these engines are rarely applied to modern civil aircraft b Turboprop engine A turboprop engine is a gas-turbine engine that delivers almost all of its power to a shaft to drive a propeller (appendix VI). The propeller drives a relatively large mass of air backwards fairly slowly, while the gas turbine propels a small mass of air backwards relatively quickly. For small aircraft with airspeeds between 216 and 377kts the turboprop engine provides the highest propulsive efficiency. However above the speed of sound the propeller efficiencies drop off quite rapidly, due to the disturbance of the airflow at the tips of the blades. So the turboprop engines are used on slow and small aircraft because of their fuel efficiency at lower airspeeds c Propfan A propfan engine can be described as a modified turbofan engine. With this engine the fan is placed outside of the engine nacelle and the fan is on the same axis as the compressor blades. The design of the engine has the intension of offering the speed and the performance of a turbofan but with the fuel economy of a turboprop engine. The propfan engine delivers a reduction of 30 percent in fuel use compared with turbofan engines. The engine has not yet proved the airworthiness, aerodynamic characteristics, and noise signature regulations. The current gas price and reduced emissions of this Group 2A2O Amsterdam, January,

8 engine make this engine interesting for aircraft builders. However, because of the lack of (released) information of the design of the engine and is not fully developed, this engine type will not be chosen for the new to be designed engine d Unducted fan An unducted fan gas turbine engine is an engine which has counter rotating propellers which are coupled to counter rotating coaxial rotors. These rotors and propellers provide a rotation about the longitudinal axis of the engine. A gear is coupled to each of the rotors to provide rotational motion. A shaft is coupled to the gear for rotation about the transverse axis. Each propeller is provided with propeller blades which each are rotatable around the corresponding blade axis. Control is coupled to the blades for varying pitch. The unducted fan provides a fuel saving of 20 to 30 percent. However, unducted fan engines are not yet certified and so the design cannot be used for short term production so this design will not be used e Turbofan engine The efficiency of turbojet engines for supersonic flight speeds was excellent. However, for high subsonic and transonic ( knots), the velocity of the exhaust gas jet was too high to obtain a good propulsive efficiency. Under these conditions, the bypass engine became a very attractive approach for improving the propulsive efficiency. The Bypass Ratio of an engine is the ratio of the amount of air which is bypassed around the hot core of the engine, to the amount of air which passes through the hot core. An engine with a low bypass ratio has a relative small amount of air which is bypassed around the hot core of the engine (appendix VII). The low bypass ratio results that the core needs to produce less power to drive the fan. An engine with a high bypass ratio has a relative high amount of airflow which passes the engine by the bypass duct (appendix VIII) Purpose of a gas turbine The purpose of a gas turbine engine is to propel a mass of air backwards. The force created by the mass of air and its velocity generates a reaction in the opposite direction driving the aircraft forwards. In other words, the mass of air is given an acceleration which produces a force (Newton s second law). According to Newton s third law: for every force acting on a body, there is an equal and opposite reaction, the force of the air should drive the aircraft forwards Operation of a gas turbine The gas turbine engine inside the aircraft is built out of sections stacked in a line from air intake to exhaust (appendix IX). Air from outside the aircraft passes through the inlet (1) the inlet directs the incoming air evenly across the inlet of the engine. The air then passes through a compressor (2) which takes in an enormous volume of air which must be compressed to a smaller volume than outside the engine. The compressor causes the air to decelerate and the pressure to rise. After the air has passed the compressor it goes to the combustion chamber (3). In the combustion chamber fuel nozzles provide a spray pattern which was ignited by the ignition system. This process continues until the fuel flow is stopped due to fuel cut-off. Between the compressor and exhaust section a turbine (4) is placed that uses some of the energy of the discharging air to drive the compressor. A shaft (5) connects the turbine with the compressor. Finally the air leaves the engine via the exhaust nozzle (6) to provide thrust Indication The aircraft engine can be operated inside the cockpit. The engine can be started using the engine start panel which is placed at the overhead panel (appendix X). Once the engine is started the engine can be operated using the engine thrust levers. Thrust lever angle determines the thrust setting (appendix XI). The thrust rating select push buttons are placed directly behind the throttles (appendix XII). Engine indications on the Fokker 100 are presented at two pages on the Multi Function Display System (MFDS) (appendix XIII). The primary page displays the primary parameters of both engines, these parameters consist out if the Engine Pressure Ratio (EPR), Turbine Gas Temperature (TGT), N1 (Low Pressure Rotor Speed), N2 (High Pressure Rotor Speed), Total Air Temperature (TAT) and Static Air Temperature (SAT). The SAT is the temperature of the air with no compression effects Group 2A2O Amsterdam, January,

9 due to aircraft movements. The TAT is the temperature of the air when it has been brought completely to rest, as in the pitot tube. The ram rise is the temperature increase due to compression and can then be subtracted from TAT to give corrected outside air temperature. The secondary page shows the secondary parameters of both engines. These secondary parameters consist out of: oil pressure indicator, oil temperature indicator, oil quantity, fuel flow indicator, fuel temperature, fuel quantity and the engine vibration indicator. The vibrations in the aircraft engine can also be monitored with the vibration push button which is placed at the engine start panel. This button shows if the vibration of both engines is below or above threshold values (appendix XIV). The MFDS can be controlled using the MFDS Controls (appendix XV). If MFDS is not available or malfunctioning a standby engine indicator display will show the most important values of the engine to the pilot (appendix XVI). 1.2 Subsystems An engine is accommodated with several systems that make the engine operate more appropriate. The accessory gearbox placed at the engine drives the sub-systems for the engine and aircraft (1.2.1). This accessory gearbox also drives the hydraulic pumps for the hydraulic system (1.2.2). At this gearbox an engine oil pump is placed for the lubrication system (1.2.3). Furthermore, most systems use electric power (1.2.4). At last, the engine delivers bleed air to power the pneumatic system. Some of the engine systems use this bleed air to operate (1.2.5) Gearbox Every gas turbine has a gearbox (figure 1.1). This gearbox is driven by the engine by the use of the N2 shaft. Via a radial drive shaft (1) the horizontal drive shaft is driven (2). This shaft is connected to the accessory gearbox. At the accessory gearbox the engine oil pumps (3) are places for the lubrication system. Also the hydraulic pumps (4) are driven to ensure there is hydraulic pressure in the hydraulic system. There are low pressure fuel pumps (5) and high pressure fuel pumps (6) that are used for the fuel management system Radial drive shaft 2. Horizontal drive shaft 3. Engine oil pump 4. Hydraulic pump 5. Low pressure fuel pump 6. High pressure fuel pump 6 3 Fig 1.1 Accessory gearbox Hydraulic system The Fokker 100 is accommodated with a hydraulic system. This system is driven by engine driven pumps and electric driven pumps. The engine drives the accessory gearbox and the accessory gearbox will drive the hydraulic pumps. This pump will make sure that there is a hydraulic pressure in the system. The Fokker 100 has two hydraulic systems and at each engine two pumps are placed. This is done to create redundancy. One engine can pressurize both systems. It also has a standby hydraulic system. 4 5 Group 2A2O Amsterdam, January,

10 1.2.3 Lubrication system Oil is stored in oil tanks (figure 1.2) (1) (appendix XVII). To get this oil out of the tank a pump driven by the engine is used (2). This pump is placed on the gearbox and sucks the oil out of the tank to the bearings. After the oil passes the pump, the oil filter (3) makes sure that no wear will take place as a result of particles in the oil. This filter consists of a filter base and a cover. Between this filter base and the cover the filtering unit is placed. At this filtering unit a pre-blockage indicator is used to indicate that the filter is blocked partly or completely. When the filter is blocked a red pin, called the red indicator, will stick out of the pre-blockage indicator. As soon as the filter is blocked partly the bypass valve opens and the oil passes unfiltered through this valve. A check valve can prevent the oil from leaking to the bearings when the engine is idle and there is no oil pressure anymore. This check valve will open when the engine is started and the oil is pumped through the filter. Several magnetic plugs separate the magnetic parts from the oil. After the oil passes this filter it comes to heat exchanger (4) which is used to cool down the oil and heat the fuel to prevent icing. Once the oil gets mixed with air a centrifugal breather (5) is placed, which has the purpose to separate the oil from the air. A part of the oil will be separated from the air in the oil tank. The air, which still contains small particles of oil, goes to the gearbox and will be purified there. A rotating wheel makes sure that the particles with the greatest density will be separated from the air. From the gearbox the oil will flow back to the tank via a drain pump Oil tank 2. Pump 3. Oil filter 4. Heat exchanger 5. Centrifugal breather Fig. 1.2 Oil system Electric systems Some of the subsystems use electric power to operate. The fuel management system is one of these systems. The fuel management system uses electric power to make it able that the fuel is regulated electrically (1.2.4a). The second system that uses electric power is the ignition system. This system uses electric power from the APU for the ignition spark (1.2.4b) a Fuel Management System The Fuel Management System (FMS) has several tasks (figure 1.3), but its main task is to manage the quantity of fuel injected into the combustion chamber. In the process of injecting the fuel, the fuel gets filtered and heated and the pressure is also raised in stages. The system gets its fuel from the booster pumps which are part of the fuel system of the aircraft. The incoming fuel has a pressure of psi. The simplified version of the FMS consists of seven main parts: 1. Engine fuel pump 2. Fuel/oil heat exchanger 3. Fuel filter 4. Hydro mechanical unit 5. Fuel shut off valve 6. Fuel flow transmitter 7. Fuel nozzles Group 2A2O Amsterdam, January,

11 1. Engine fuel pump 2. Fuel oil/heat exchanger 3. Fuel filter 4. Hydro mechanical unit 5. Fuel flow transmitter Figure 1.3 Fuel management system ad 1 Engine fuel pump The pump gets its fuel from the booster pumps. These pumps have raised the pressure to psi. The engine fuel pump raises the pressure even more by the use of stages. First there is the impeller; this raises the pressure level to psi. When the pressure is raised, the fuel is led through an interst strainer. This filters the fuel to prevent any particles from travelling further into the system and which may damage parts on the way. The fuel is led to the gear pump, which raises the pressure to psi. This is the pressure which the fuel nozzles use to inject the fuel into the combustion chamber. ad 2 Fuel/oil heat exchanger To prevent the fuel from clotting and clogging the filter, the fuel is heated. This is done by leading the fuel through a pipe system and on the outside of this pipe system hot air from the engine is blown. When the fuel leaves the pipe system it passes a Fuel Temperature Sensor Valve (FTSV). If the temperature of the fuel exceeds the limit the FTSV will close up. By doing that, the pressure will build in the oil bypass valve assembly. This valve will make it possible for the hot air to move directly from the inlet to the outlet and so the pressure and further rising of the heat is prevented. ad 3 Fuel filter The filters goal is to prevent any particles from travelling further into the system and thereby damaging any parts along the way. In case the filter is blocked by particles; the flow through the filter will diminish. This causes a pressure build up at the inlet side of the filter. When this happens a bypass valve will open so the fuel will move around the filter, proceeding to the fuel nozzle unfiltered. If the pressure difference of the filter is seven psi or more a fuel alert will go off in the cockpit. ad 4 Hydro Mechanical Unit The Hydro Mechanical Unit (HMU), also known as the fuel metering valve, is connected to the gearbox of the engine. The goal of the HMU is to measure and deliver the exact amount of fuel needed for combustion to the fuel nozzles. The HMU is controlled by a Full Authority Digital Engine (FADEC). The FADEC controls the Electronic Engine Control (EEC) to process all given variables and thereby measuring the right amount of fuel; but the most import function of the FADEC is trend monitoring. It monitors and saves engine data like the temperature in the different stages of the engine. This Group 2A2O Amsterdam, January,

12 data is used to analyze the engine thoroughly and by that improving the maintenance or even improve the engine as a whole (appendix XVIII). The FADEC is able to control the rate fuel delivery electronically. The Power Lever (PL) in the cockpit controls the RPM. The position of the PL is translated to an electric current by the resolver and transported to the EEC. The EEC translates this current into a suitable signal for the HMU. The fuel pump always delivers too much fuel. The excess fuel is returned to the pump by the bypass fuel return. ad 5 Fuel shut-off valve The fuel shut-off valve makes it possible to shut-off or open the fuel line to the system. This is also controlled electronically by the EEC and this switch is located directly beneath each PL. ad 6 Fuel flow transmitter The fuel flow transmitter measures the amount of fuel that goes to the fuel nozzles and sends a signal to the cockpit. This makes it visible for the pilot to see what the fuel consumption in kilograms per minute is. ad 7 Fuel nozzles Eventually the fuel is injected into the combustion chamber. There are several types of fuel nozzles which inject the fuel in different ways (appendix XIX). To minimize leakage and exposing the fuel to high temperatures which perhaps might lead to fire. A part of the pipe lines leading to the nozzles has double walled pipes. If any fuel would get in the second compartment of the lines this would leak back to the drain manifold. The EEC needs a reference value to calculate the position of the PL. This is the Steady State Line (SSL), this indicates which fuel/air ratio ( f/ a) is needed to run a certain RPM at a certain temperature. During acceleration or deceleration the f/ a does not follow the expected SSL but deviates from it. There are several extreme values for f/ a which cannot be exceeded without consequences (appendix XX). These five extreme values for f/ a are the rich blow out area, stall area, overtemp area, overspeed area and the Lean die out area (figure 1.4). 1. Rich blow out area 2. Stall area 3. Overtemp area 4. Overspeed area 5. Lean die out area Figure 1.4 Danger areas Group 2A2O Amsterdam, January,

13 As mentioned before, the temperature influences the SSL. If the temperature drops the stall area will move down and the lean die out area will move upward. The EEC will use extreme values for each specific temperature b Ignition process The engine has a pneumatic starter. This is powered by a compressor which gets its electrical power from the APU. This compressor pumps the air through the starter air valve to the stator vanes. These vanes direct the air flow in such a way to the turbine vanes, that the impact energy is transferred maximally. This turbine will run with an RPM rate of It is connected to the gear train. The gear train converts this high rotation rate to a lower rate and that powers the clutch. The clutch makes sure that the starter drives the gas turbine and not the other way around. If the gas turbine would be directly linked to the starter turbine; the starter turbine would possibly explode because of the high RPM rates run by the gas turbine. There are several mechanical link possibilities to prevent this problem. When the gas turbine starts to run, the ignition (figure 1.5) is started after a couple of seconds (1). The exciter box produces the electricity needed for the ignition spark. The latest systems work with a voltage of 115 V 400Hz AC. The exciter is located on the outside of the gas turbine and is most of the times an air tight box. Usually only one of the ignition systems is started. After a couple of seconds the fuel is injected into the combustion chamber and the fuel is ignited (2). When this happens the Exhaust Fuel Temperature (EGT) starts to rise. The N2 axle will start to speed up and because the N 1 axle is aerodynamically linked to it, the N1 axle will also start to speed up. At a certain moment the both axles spin around fast enough to accelerate on their own (3). The ignition and the starter can be stopped then. The N2 axle will rise to its idle RPM. The EGT will reach a peak value (4) and eventually stabilize at the idle temperature (5). 1. Ignition start 2. Fuel injected/ignited 3. Self sustainable RPM rate N 1 /N 2 4. Peak value EGT 5. Stabilized EGT Figure 1.5 Starting cycle There are several problems that could occur by starting the engine (appendix XXI) Bleed air The starter engine uses the bleed air from the high pressure compressor to power the pneumatic systems with approximately psi (1.2.5a). Some of the systems in the engine also use this air; these systems are called the bleed air systems (1.2.5b) a Pneumatic power The gas turbine has an bleed air to the pneumatic system of the aircraft. At several stages air is drained from the high pressure compressor. At a certain stage the air has a higher pressure and temperature. In a level flight a high stage valve is closed and is used to power the pneumatic system. This Group 2A2O Amsterdam, January,

14 high stage valve will open when the engines have a lower speed. Then one of the last stages is also used to power the pneumatic system. So, this higher stage valve makes sure the pressure in the pneumatic system stays constant. This pressure will be between the 50 and 55 psi. When the pressure reaches 90 psi a pressure relief valve will open to lower this overpressure. The pressure will go to a pre-cooler. This pre-cooler cools the air temperature down to a temperature of 175 C b Bleed air system systems As described above, the air of the pneumatic systems is supplied by the engines. Not only the systems in the aircraft, but also the engine systems use this air for powering these systems. The systems used by the engine are: 1. Engine anti-icing 2. Internal cooling 3. Core compartment cooling 4. Turbine case cooling ad 1 Engine anti-icing The engine, especially the inlet, has to be free of ice. This is to prevent interference of the airflow and damage to the fan or compressor when the ice is released from the inlet. That is why the edge of the inlet and the spinner cone are heated (figure 1.6). For heating, air from the pneumatic systems is used. This air is led through a thermal anti-icing duct (1) and an anti-icing pressure valve (2) to an annular tube. This tube, called the engine cowling (3), has a large number of holes. From this tube the hot air flows to the cowling Thermal anti-icing duct 2. Anti-icing pressure valve 3. Cowling 2 3 Fig 1.6 Engine anti-icing system ad 2 Internal cooling This system uses pressure from the low pressure compressor to cool down and seal the bearings in the engine. With the use of cooling valves, the pressure will be conducted to seal the bearing spaces. Also a controlled air flow from one of the stages of the high pressure compressor is led to the cylindrical construction. This air flow has the purpose to cool down the rotor of the high pressure compressor. This air will flow to the turbine section to cool down the turbine wheels and blades also. There are different types of cooling (appendix XXII): Convection cooling is a type of cooling where cooling air flows into the turbine blade. By cooling the wall, heat is removed. In the gas turbines of nowadays convection cooling is widely used. Impingement cooling is a form of convection cooling. In this form of cooling, the cooling air jets have a high velocity in comparison to convection cooling. With impingement cooling it is possible to cool one section more than another section. Film cooling uses the cooling air to form an insulating layer. This layer is set between the walls of the blades and the hot air. Transpiration cooling makes use of the porous walls of the blades. Cooling air flows through these porous walls and heat is immediately removed by the cooling air. Group 2A2O Amsterdam, January,

15 Water/steam-cooling uses tubes in the turbine blade. Water passes through these tubes and it will become steam when emitted to the tips of the blades. Steam has a higher capacity of transferring heat than air has. ad 3 Core compartment cooling Air from the fan or low pressure compressor is used to cool down the section between the engine and the plating. For this system the spaces will be divided from each other in isolated areas. At each zone the temperature is regulated and the formation of flammable gases will be prevented by a controlled bleed air through every zone. The first zone is the zone with the accessory gearbox. This gearbox consists of fuel and oil pumps. Because of the risk of flammable gases is high, the zone is ventilated using air that will enter the engine at the top and leave the engine via the air outlet at the bottom. Zone two only consists of oil ducts and is also cooled by bleed air. Zone three consists of the combustion chambers and turbines. This zone is cooled with fan air via two openings, the cooling air inlet and the cooling rear inlets. ad 4 Turbine case cooling The high pressure turbine as well as the low pressure turbine will be cooled with the purpose to limit expansion of the turbine stator case. The turbine case will be cooled by using air from behind the fan. This air is drained by using an extended piping system, called the bird cage. By using this bird cage the air is blown to the turbine stator case. 1.3 Gas turbine theory Before a new gas turbine for the Fokker 100 can be established, a research must be made of the operation of the gas turbine (1.3.1). The engine can be divided in five different stages with different tasks. The first stage of the gas turbine is the inlet (1.3.2). Following the inlet is the stage where the air gets compressed, called the compressor (1.3.3). The stage where the air gets combined with fuel is called the combustion chamber (1.3.4). After the hot gasses exits the combustion chamber, the gasses will pass through the turbine stage (1.3.5). The final stage of a gas turbine is the exhaust nozzle (1.3.6) Purpose The gas turbine engine is a heat engine using air as a working fluid to provide thrust. To achieve thrust, the air has to be accelerated. To achieve this, the velocity of the airflow or the kinetic energy of the airflow has to be increased. To obtain this, the pressure should increase first. Secondly, addition of heat energy and finally conversion back to kinetic energy should occur (appendix XXIII). These situations can be shown in the Brayton process. The Brayton process is a schematic figure from a two axis gas turbine. This process contains the following stages: inlet, compression, combustion, expansion and exhaust. This process exists of a pressure-volume diagram and a temperature-entropy diagram (figure 1.7). The surface under the graphic stands for the labour of the process Compression Combustion Expansion Ambient air Figure 1.7 p-v-diagram T-S-diagram Group 2A2O Amsterdam, January,

16 Point 1 refers the air by atmospheric pressure that is compressed following line 1-2. From point 2 till point 3, heat is added to the air by the use of burning fuel by constant pressure. Also the volume of the air will increase. Pressure will be lost in the combustion chamber as seen from point 2 to point 3. From 3 till 4, the gas, resulting from the combustion, will flow through the turbine and the pressure will flow back to the atmospheric pressure. Parts of the energy will be converted, by the use of the turbine, in mechanical power, resulting in thrust when approaching the atmosphere. During compression, the pressure will increase and the volume will decrease from the airflow. As a result of this, the temperature will increase. During combustion, when fuel is added to the airflow and will be burned, the temperature will decrease. As result of the temperature rising, the volume will increase and the pressure will be constant. During expansion, the temperature will decrease and the pressure and volume will increase Inlet The purpose of the inlet is to provide the compressor, of the engine, with air. The form and size of the inlet have been coordinated on the most common circumstances of the aircraft (1.3.2a). The inlet of an engine has to be made of a material that is strong but at the same time light (1.3.2b) a Principle There are different inlet possibilities when concerning the velocity of the aircraft. There are three different inlets: a subsonic, a supersonic or a static inlet. The air, that enters the inlet, must have an axial velocity and minimized camber. This applies for every stage in the engine during take-off, taxiing, normal flight conditions, crosswind and landing. The inlet has a divergent lapse. The basic principle of the divergent inlet of the engine is to decrease the speed. The speed has to be decreased to reach the desirable axial entrance speed for the compressor. When this speed is higher, it can cause damage to the compressor. Because of the decrease in speed, the pressure and temperature will increase. In an uncompressible medium, the density will stay constant. However, in a medium that is compressible, the density has to be taken into account because the density will be variable. This can be shown with the equation of continuity (appendix XXIV). When the inlet divergence, the area increases the velocity decreases. When the speed is decreasing, the pressure in the inlet must rise, as shown in the first law of Poisson (appendix XXIV). When the pressure is rising inside the inlet, the air temperature will automatically increase, as shown with the third law of Poisson (appendix XXIV) b Materials The most common materials for the inlet section are aluminium alloys. Aluminium has been chosen because it is a material with good properties. Principally, it is a light but still strong material Compressor The compressor is the integral part of the gas turbine engine. The air can be compressed in two different ways: by means of centrifugal flow or by means of axial flow. Both types are driven by the engine turbine and are mostly directly coupled to the turbine shaft. The engine of the Fokker 100 works with the use of an axial flow compressor (1.3.3a). The compressor exists of different stages. During operation, each stage will increase in pressure (1.3.3b). As a reaction of the increase of pressure, the temperature will also increase. Because of the temperature rise, a good reflection is necessary concerning the use of the materials (1.3.3c) a Components The axial flow compressor of the Fokker 100 is a multi-stage unit with sixteen stages where the air will be compressed in axial lengths. No change in direction will take place of the air flow. Because of this, a high efficiency can be reached. The axial flow compressor consists of two main parts (figure 1.8): Rotor blades Stator vanes Group 2A2O Amsterdam, January,

17 The axial flow compressor consists out of one or more rotor assemblies that consist out of blades. These assemblies are located between bearings in the casing that hold the stator vanes in the right position. The increase of pressure in every stage is very low. Each stage consists of a row of rotor blades (1) followed by a row of stator vanes (2). When several rows of stages are following each other (3), it is necessary to place the stator vanes under a certain angle. Because of this, a velocity triangle will occur (appendix XXV). Variation in the angle of the stator vanes is necessary to let the compressor work below certain speeds of the design conditions. Since a higher pressure ratio will be reached, the change of the stator vanes angles ensures that the air flow reaches the next row of rotor blades under an acceptable angle. Because of the different angles, problems can occur as stall or choke (appendix XXVI). 1. Nine rotor blades 2. Nine stator vanes 3. Rows of stators and rotors Figure 1.8 Rotor blades and stator vanes The axial flow compressor of the engine is equipped with a twin spool engine (figure 1.9). This means, the compressor has a low pressure spool (1) and a high pressure spool (2). The low pressure spool consists of a three stage intermediate pressure compressor. This compressor is driven by a three stage low pressure turbine and is fitted with a bleed valve. A bleed valve is used to prevent compressor stall during fast deceleration. Only a percentage of the air from the low pressure compressor goes into the high pressure compressor. The high pressure spool consists of a twelve stage high pressure compressor. This compressor is driven by a two stage high pressure turbine. The high pressure compressor is fitted with variable inlet guide vanes and an annular bleed valve at the seventh stage to increase stall margin and improve engine operating characteristics Low pressure spool 2. High pressure spool 2 Figure 1.9 Twin-spool compressor 1.3.3b Operation During operation the rotor will rotate with high speeds with the use of the turbine. The rotor will be rotating so the air will be pushed continuously in the compressor. The speed of the air will now increase by the rotor blades and will be pushed to the following stator vanes. An increase in the pressure will occur because energy will get into the rotor that increases the air velocity. The air will now slow down in the following stator vanes and the kinetic energy, which is produced, will now convert into potential energy (appendix XXIV). The vanes of the stators are also used to correct the airflow. The function of the last row of stator vanes is to straighten the air flow and to make the airflow uniform with an axial velocity before en- Group 2A2O Amsterdam, January,

18 tering the combustion. Changes in pressure and velocity will occur during the air flow through the compressor (appendix XXVII). The changes in pressure and velocity do also have influence on the air temperature. Through every stage the proportion from the total pressure from the output flow, in comparison to the input air flow is between the 1:1 and the 1:2. The reason for the small pressure increase through every stage is that the inflection of the air in every stage has to be limited to avoid subsequent blade stall. Although, the increase in pressure in every stage is low, the increase of the output pressure of a next stage is higher than the stage before. For example, the first stage has a pressure increase of three to four psi and the eighth stage of a thirty to one compressor system will now have an increase of eighty psi. The increase in pressure can now be calculated per stage (appendix XXIV). Because of the possibility of a multi-stage compressor, with a controlled air velocity and an axial flow, the losses of air will be minimized and makes high efficiencies possible and therefore a lower fuel usage c Materials The materials which are used for a compressor are chosen to compose an efficient design. There has to be thought about a light material that can handle high forces and temperatures. For casings the need of a material that is light and stiff is necessary. These needs are reached by the use of aluminium at the front side of the compressor system following by steel alloys because of the increase of the temperature. In the rest of the compressor, nickel alloys will be used because of the high temperature that will occur. The stator vanes are made out of steel or nickel alloys. Titanium can be used for the stators in the low pressure areas but is not capable for the stator vanes because of the increase in pressure and temperature. During the process of making rotor discs, drums and blades, a metal is needed with a high ratio and strong density. This results in the lightest possible rotor assemblage which can handle the forces. Because of this, titanium is a proper usable material besides its high costs Combustion The purpose of the combustion chamber is to convert chemical energy that is located in the added fuel, in heat. In the Fokker 100 this conversion is possible by the use of an annular combustion chamber (1.3.4a). The exerted product of the combustion chamber is a homogenous heat gas flow with an acceptable temperature for the turbine (1.3.4b). Because of the high temperatures in the combustion chamber, proper consideration for the manufacturers is necessary to choose a high temperature resistant and light material (1.3.4c) a Components The annular of the compressor connects flush to the input of this type of combustion chamber. Also the output of the combustion chamber connects flush to the annular input of the turbine. The combustion chamber consists of two concentric placed annular. Inside these annular the combustion process will take place. These annular are also called the inner and outer casing (appendix XXVIII). On the front side of the combustion chamber several spray nozzles are placed to press fuel, under high pressure, into the combustion. The walls inside the combustion chamber are equipped with several holes, to receive cool air for cooling the walls after a combustion process. The combustion chamber starts with a row of stator vanes. These vanes belong to the last row of the compressor and makes sure the air will enter the combustion chamber with a purely and axial flow. By this way, the losses of air will be minimized. These stator vanes, including the diffuser, form the integral part of the combustion section. In the outside wall of the combustion chamber, holes are made to attach the fuel spray nozzles. The engine is equipped with ten fuel spray nozzles. The most important advantage of the annular combustion chamber is that, with the same power output, the length of the chamber is 75 percent in comparison to the tubo-annular combustion chamber with the same frontal diameter. This makes a saving on weight and production costs possible. Group 2A2O Amsterdam, January,

19 1.3.4b Operation In a gas turbine the combustion takes place under almost constant pressure. The pressure changes will be so small that this can be neglected. The combustion chamber burns large amounts of fuel that enter the chamber by the use of fuel spray nozzles. At the same time, high amounts of air volumes, which are coming from the compressor, enter the combustion chamber. The air will now expand and an increase of the velocity will take place, forming a uniform heated gas that will flow to the turbine. This has to be reached with a minimum loss of pressure and a maximum heat release. The amount of fuel that is added to the air depends on the temperature rise. The maximum temperature that can be reached, in contrast to the materials in the turbine and in the fuel nozzles, is around the 850 and the 1700 degrees Celsius. In the compressor, the temperature will rise till 200 to 550 degrees Celsius. Then the combustion chamber ensures a temperature increase of 650 to 1150 degrees Celsius. The air temperature varies with the engine thrust. The combustion chamber has to be well functioned under a wide range of engine operating conditions. The air from the engine compressor flows through the combustion chamber with a velocity of 500 feet per second. This velocity of the air is too high for combustion and therefore the air has to be diffused. This is done by the use of slowing down the air and increase the static pressure. The speed of burning fuel, during normal conditions, is a few feet per second. This flow has now a velocity of eighty feet per second. A region of low axial velocity will be created in the chamber, so the flame will be continuing burning during engine operating conditions. In normal operations, the proportion between air and fuel in the combustion chamber is 45:1 and varies to 130:1. The fuel is most efficient used at a proportion of 15:1 combustion, so the fuel is only burnt with a small amount of air that is entering the combustion chamber (figure 1.10). This is also called: primary combustion zone (1). This zone is reached with the use of a flame tube to measure the airflow alongside the chamber. About twenty percent of the airflow is absorbed by the entry section. This air will be flown into the primary combustion zone. The arisen upstream air flows from the centre of the flame tube to the desired recirculation. The air, which is not taken by the entry section, will flow to the annular space between the flame tube and the air casing. Through the walls of the flame tube, besides a combustion process, a selector process takes place. The most important twenty percent of the airflow will flow through the secondary holes of the primary zone to the dilution zone (2). 1. Primary zone 2. Dilution zone 1 Figure 1.10 Combustion chamber 2 The vortex airflow and the airflow of the secondary air holes are working together so a recirculation will occur. This recirculation will stabilize the flame of the combustion chamber. The recirculation of the gasses will increase the combustion process of the new fuel, because this new fuel will reach its ignition temperature fast. The temperatures of the released gasses of the combustion are around 1800 up to 2000 degrees Celsius. This temperature is too high for entering the nozzles of the turbine. The air that is not used for the combustion process, around the sixty percent of the airflow, will now be added to the flame tube. One third of this air will be used in the dilution zone to decrease the temperature of the gas flow before entering the turbine. The rest of the airflow, about two third, will now be used to cool the walls of the flame tube after combustion. This is reached because cooling air is flown along the inside wall of the flame tube, isolated from the heat of the combustion gasses. The combustion has to be complete before the dilution air will cool the flame. Otherwise, the combustion Group 2A2O Amsterdam, January,

20 result will be incomplete. An electric spark from a lighting mechanism starts the combustion process and the flame is further self sustained c Materials The walls and the intern parts of the combustion have to be resistant against the very high temperatures in the primary zone. This is reached with the use of heat resistant material, heat durable covers and to isolate the inside walls by cooling during the flame. The combustion chamber has to be resistant to corrosion that can be expelled during combustion. The materials also have to be resistant to high heat and large vibration stresses Turbine The second to last stage of the gas turbine is the turbine. The turbine consists of several components that are reliable for the function of the turbine (1.3.5a). There are several tasks that the turbine has to perform (1.3.5b). The turbine is situated after the combustion chamber, making the components vulnerable for high temperatures (1.3.5c) a Components The basic components of the turbine (figure 1.11) are the combustion discharge nozzles mounted on the combustion flange (1), the nozzle guide vanes (2), the turbine discs and the turbine blades (3). The rotating assembly is carried on bearings (4) mounted in the turbine casing and the turbine shaft (5) mechanically linked to the compressor shaft (6). The nozzle guide vanes are made with an aerofoil shape so the passage between the vanes forms a convergent duct. The vanes are located in the turbine casing in a manner that allows for the air to expand. The nozzle guide vanes are most likely to be hollow, so the nozzles can be cooled. This cooling is done by passing cold air from the low pressure compressor through the hollow parts of the vanes, this will reduce the effects of high thermal stresses and gas loads. The turbine blades are made of an aerofoil shape, designed to provide the gas to pass between adjacent blades that give a steady acceleration of the flow up to the exhaust. Here the area is smallest and the velocity reaches the required speed to exit the engine and produce the required amount of trust Combustion system mounting flange. 2. High pressure nozzle guide vane 3. Single stage high pressure turbine 4. High pressure turbine bearing 5. High pressure turbine shaft (N2) 6. Low pressure turbine shaft (N1) 7. Three-stage low pressure turbine 8. Exhaust unit mounting flange 9. Turbine rear bearing 1 Figure 1.11 Twin shaft turbine 1.3.5b Operation In a gas turbine engine, the turbine has the task of providing the power to drive the compressor and other accessories mounted on the gearbox of the engine. The turbine makes use of this power to Group 2A2O Amsterdam, January,

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