project report power plant - portfolio anouk elsendoorn ... report power plant group 2a2o amsterdam,...

Project report power plant

Group 2A2O Amsterdam, January, 2010



Preface

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



Contents

Summary ..................................................................................................................................... 1

Introduction ................................................................................................................................. 2

1 Gas turbine research ............................................................................................................. 3

1.1 General gas turbine theory ..................................................................................................... 3

1.1.1 Types of engines .............................................................................................................. 3

1.1.2 Purpose of a gas turbine .................................................................................................. 4

1.1.3 Operation of a gas turbine .............................................................................................. 4

1.1.4 Indication ......................................................................................................................... 4

1.2 Subsystems .............................................................................................................................. 5

1.2.1 Gearbox ........................................................................................................................... 5

1.2.2 Hydraulic system ............................................................................................................. 5

1.2.3 Lubrication system .......................................................................................................... 6

1.2.4 Electric systems ............................................................................................................... 6

1.2.5 Bleed air ........................................................................................................................... 9

1.3 Gas turbine theory................................................................................................................. 11

1.3.1 Purpose .......................................................................................................................... 11

1.3.2 Inlet ................................................................................................................................ 12

1.3.3 Compressor ................................................................................................................... 12

1.3.4 Combustion ................................................................................................................... 14

1.3.5 Turbine .......................................................................................................................... 16

1.3.5a Components .................................................................................................................. 16

1.3.6 Exhaust Nozzle ............................................................................................................... 17

1.4 Engine Properties .................................................................................................................. 17

1.4.1 Performance .................................................................................................................. 18

1.4.2 Environment .................................................................................................................. 19

1.5 Demands ................................................................................................................................ 21

1.5.1 Client Demands ............................................................................................................. 21

1.5.2 Demands Authorities ..................................................................................................... 22

1.6 Next generation engines ....................................................................................................... 23

1.6.1 General gas turbine innovations ................................................................................... 24

1.6.2 Material innovations ..................................................................................................... 24

1.7 Function research .................................................................................................................. 25

2 Analysis gas turbine parts .................................................................................................... 26

2.1 Morphological overview .............................................................................................................. 26



2.1.1 Air intake .............................................................................................................................. 26

2.1.2 Fan blades ...................................................................................................................... 26

2.1.3 Bypass ............................................................................................................................ 27

2.1.4 Compressor ................................................................................................................... 27

2.1.5 Combustion ................................................................................................................... 28

2.1.6 Turbine .......................................................................................................................... 29

2.1.7 Exhaust nozzle ............................................................................................................... 30

2.2 Possible designs ..................................................................................................................... 31

2.2.1 Design One: Noise reduced engine ............................................................................... 31

2.2.2 Design Two: Low fuel consumption engine ................................................................... 32

2.2.3 Design three: Low emission engine ............................................................................... 32

2.3 Engine performance .............................................................................................................. 33

2.4 Advantages and disadvantages ............................................................................................. 34

2.4.1 The three designs .......................................................................................................... 34

2.4.2 Advantages and disadvantages overview ..................................................................... 35

2.5 Conclusion ............................................................................................................................. 36

3 Engine design ...................................................................................................................... 36

3.1 HVA CS-25E engine ................................................................................................................ 36

3.1.1 The engine specified ...................................................................................................... 36

3.1.2 Sub systems specified .................................................................................................... 37

3.1.2 HVA CS-25E layout ......................................................................................................... 38

3.2 Certification ........................................................................................................................... 38

3.2.1 Type certificate .............................................................................................................. 38

3.2.2 Airworthiness certificate ............................................................................................... 38

3.3 Maintenance research ........................................................................................................... 39

3.3.1 High and low maintenance parts................................................................................... 39

3.3.2 Engine inspection .......................................................................................................... 40

3.4 Financial aspects .................................................................................................................... 40

3.4.1 Design costs ................................................................................................................... 40

3.4.2 Maintenance costs ........................................................................................................ 40

3.4.3 Benefits .......................................................................................................................... 40

3.4.4 Breakeven point ............................................................................................................ 41

3.5 Final conclusion ..................................................................................................................... 41

3.6 Recommendation .................................................................................................................. 41

Reference list ............................................................................................................................. 42


Group 2A2O Amsterdam, January, 2010 1

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 183-03 MK 620-15 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 en-

gine. 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. Be-

sides the engine, there are several subsystems that making the operation of the aircraft more man-

ageable 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 desir-

able 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 pro-

pelling 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 51065.5 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 effi-

ciency. 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 spe-

cific 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 compo-

nent 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 com-

ponents 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 im-

proved in thrust and has a reduced emission. For Fokker 100 of nowadays the engine is not recom-

mended because the break-even point is at nineteen years. For the Fokker 100 of the next genera-

tion this design is recommended because it makes operation at 9000 ft on hot day conditions more

efficient.



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 effi-

cient, less noisy and have less emission. Because Fokker Aircraft operates in South-America, the per-

formance 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 writ-

ten 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 compo-

nents and their function of the gas turbine can be researched. These functions are: inlet, compres-

sion, 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 dis-

cuss is the laws and requirements. The design of the engine has to meet with these laws. Also re-

search 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 innova-

tions 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 op-

tion. 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 com-

pressor 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 main-

tenance 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 (appen-

dix III). Finally, for every abbreviation in this report the abbreviation list can be used (appendix IV).



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 air-

craft, 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 183-03 MK 620-15 turbine engine of the Fokker 100

will be done to obtain knowledge about the stages in the engine (1.3). Thereupon the theory, equa-

tions 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).

1.1.1 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 en-

gine is also a gas turbine engine but it is used for helicopters so this engine will not be discussed.

1.1.1a 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.

1.1.1b 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 dis-

turbance 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.

1.1.1c 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



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.

1.1.1d 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 produc-

tion so this design will not be used.

1.1.1e Turbofan engine

The efficiency of turbojet engines for supersonic flight speeds was excellent. However, for high sub-

sonic and transonic (400 – 540 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 ap-

proach 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 rela-

tive high amount of airflow which passes the engine by the bypass duct (appendix VIII).

1.1.2 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 for-

wards. 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.

1.1.3 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 out-

side 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.

1.1.4 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 (appen-

dix XII). Engine indications on the Fokker 100 are presented at two pages on the Multi Function Dis-

play System (MFDS) (appendix XIII). The primary page displays the primary parameters of both en-

gines, 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



due to aircraft movements. The TAT is the temperature of the air when it has been brought com-

pletely 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 moni-

tored 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 (ap-

pendix 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 gear-

box 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).

1.2.1 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.

Fig 1.1 Accessory gearbox

1.2.2 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 gear-

box 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.

1. Radial drive shaft

2. Horizontal drive shaft

3. Engine oil pump

4. Hydraulic pump

5. Low pressure fuel pump

6. High pressure fuel pump

1

2

6

3

4

5



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 indi-

cate 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.

Fig. 1.2 Oil system

1.2.4 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).

1.2.4a 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

20-50 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

1. Oil tank

2. Pump

3. Oil filter

4. Heat exchanger

5. Centrifugal breather

1

2

3

4

5



1

2

3

4

5

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 20-50 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 100-150 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 1000-1200 psi. This is the pressure which the fuel nozzles use to inject the fuel into the combus-

tion 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 tem-

perature 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 damag-

ing 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 gear-

box 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



data is used to analyze the engine thoroughly and by that improving the maintenance or even im-

prove the engine as a whole (appendix XVIII). The FADEC is able to control the rate fuel delivery elec-

tronically. 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 re-

turned 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 sig-

nal 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

2

3

4

5

1. Rich blow out area

2. Stall area

3. Overtemp area

4. Overspeed area

5. Lean die out area

Figure 1.4 Danger areas



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.

1.2.4b Ignition process The engine 620-15 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 trans-

ferred maximally. This turbine will run with an RPM rate of 50.000. 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 N1

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

2

3

4

5

1. Ignition start

2. Fuel injected/ignited

3. Self sustainable RPM rate N1/N2

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).

1.2.5 Bleed air The starter engine uses the bleed air from the high pressure compressor to power the pneumatic

systems with approximately 15-49 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).

1.2.5a 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 tem-

perature. In a level flight a high stage valve is closed and is used to power the pneumatic system. This



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 pres-

sure 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.

1.2.5b Bleed air system systems

As described above, the air of the pneumatic systems is supplied by the engines. Not only the sys-

tems 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.

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 cylin-

drical construction. This air flow has the purpose to cool down the rotor of the high pressure com-

pressor. 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.

1. Thermal anti-icing duct

2. Anti-icing pressure valve

3. Cowling

1

2

3



• 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 con-

trolled bleed air through every zone. The first zone is the zone with the accessory gearbox. This gear-

box consists of fuel and oil pumps. Because of the risk of flammable gases is high, the zone is venti-

lated using air that will enter the engine at the top and leave the engine via the air outlet at the bot-

tom. 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 op-

eration 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 noz-

zle (1.3.6).

1.3.1 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, addi-

tion 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, expan-

sion and exhaust. This process exists of a pressure-volume diagram and a temperature-entropy dia-

gram (figure 1.7). The surface under the graphic stands for the labour of the process.

1. 1-2 Compression

2. 2-3 Combustion

3. 3-4 Expansion

4. 4-1 Ambient air

Figure 1.7 p-V-diagram T-S-diagram



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 com-

pression, 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.

1.3.2 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).

1.3.2a 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, taxi-

ing, normal flight conditions, crosswind and landing. The inlet has a divergent lapse. The basic princi-

ple of the divergent inlet of the engine 620-15 is to decrease the speed. The speed has to be de-

creased 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 tempera-

ture will increase. In an uncompressible medium, the density will stay constant. However, in a me-

dium that is compressible, the density has to be taken into account because the density will be vari-

able. 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 (appen-

dix XXIV).

1.3.2b 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.

1.3.3 Compressor The compressor is the integral part of the gas turbine engine. The air can be compressed in two dif-

ferent 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 pres-

sure, the temperature will also increase. Because of the temperature rise, a good reflection is neces-

sary concerning the use of the materials (1.3.3c).

1.3.3a 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



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 ro-

tor blades under an acceptable angle. Because of the different angles, problems can occur as stall or

choke (appendix XXVI).

3

1

2

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 620-15 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 com-

pressor 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 sev-

enth stage to increase stall margin and improve engine operating characteristics.

1

1. 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 in-

crease by the rotor blades and will be pushed to the following stator vanes. An increase in the pres-

sure 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-



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 (ap-

pendix 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.

1.3.3c 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 alumin-

ium 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 tem-

perature 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 in-

crease 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 assem-

blage which can handle the forces. Because of this, titanium is a proper usable material besides its

high costs.

1.3.4 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 tem-

perature resistant and light material (1.3.4c).

1.3.4a 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 com-

bustion 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 620-15 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.



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 con-

ditions.

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 sec-

tion, 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 dilu-

tion zone (2).

1

2

1. Primary zone

2. Dilution zone

Figure 1.10 Combustion chamber

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



result will be incomplete. An electric spark from a lighting mechanism starts the combustion process

and the flame is further self sustained.

1.3.4c Materials

The walls and the intern parts of the combustion have to be resistant against the very high tempera-

tures 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 resis-

tant to corrosion that can be expelled during combustion. The materials also have to be resistant to

high heat and large vibration stresses.

1.3.5 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).

1.3.5a 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 tur-

bine 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.

1. 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 mount-

ing flange

9. Turbine rear bearing

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

3

8

2

4

5

9

6

7

1



propel the aircraft forward. In the case of engines which do not make use airflow for propulsion, the

turbine will provide power to a shaft for a propeller or rotor. This power is obtained by extracting

energy from the hot airflow coming from the combustion chamber which will spin the rotor blades in

the turbine. The flow of gasses going through the turbine is lowered in pressure and temperature,

making the components of the turbine suffers high stresses during this process. The turbine blade

tips may rotate at speeds over 1500 feet per second. The continuous flow of gas to which the turbine

is exposed may have an entry temperature between 850 and 1700 °C and may reach a velocity of

over 2500 feet per second in parts of the turbine. To produce the driving torque, the turbine may

consist of several stages each employing one row of stationary nozzle guide vanes and one row of

moving blades. The number of stages depends upon the relationship between the power required

from the gas flow, the rotational speed at which it must be produced and the diameter of turbine

permitted. The Fokker 100 gas turbine consists of five turbine stages. Modern gas turbines are also

rested with a twin spool configuration, meaning two pressure turbines. The high pressure turbine

mounted on the N2 spool powers the high pressure compressor and the low pressure turbine

mounted on the N1 spool powers the low pressure compressor on the front of the engine.

1.3.5c Materials

High temperature and the high speed rotation are the greatest limiting factor on the nozzle guides

vanes and the turbine blades. Making the nozzle guides vanes and the turbine blades more vulner-

able for fracture. The nozzle guide vanes are in static position making the high temperatures more

vulnerable for the metal. Therefore the vanes are made of nickel alloys with ceramic coating to pre-

vent the nickel from melting. Turbine discs, on the other hand, must rotate at high speeds, making

the metal more affective to fatigue. These discs are made out of a nickel based alloy, increasing the

life limits en fatigue resistance of the discs.

Turbine blades must be strong enough to carry the rotating loads and the great temperatures of the

gasses coming from the combustion chamber. These turbine blades where made out of high temper-

ature steel forgings, but these were rapidly replaced by cast nickel base alloys which give better

creep and fatigue properties.

1.3.6 Exhaust Nozzle The final stage of the gas turbine where the gas exits is called the exhaust. The exhaust has the task

to accelerate the gasses to the required velocity and direction before exiting the gas turbine (1.3.6a).

For a proper function of the exhaust a specific material must be chosen (1.3.6b).

1.3.6a Principle

After the gasses passes through the turbines, the gasses have to have a specific condition to have the

proper amount of thrust. The gasses coming out of the turbine have a speed of approximately 750 to

1250 feet per second almost 0.5 Mach. These gasses can also reach a temperature of 850 degrees

Celsius. The tasks of the exhaust are to make the gas flow in the required direction, with the required

velocity to create the thrust necessary to propel the aircraft. The exhaust consists of an exhaust

cone. This conical shaped device is positioned on the rear turbine disc, protecting the discs against

extreme heat. The task of the exhaust cone is to reduce the speed of the gasses coming out of the

turbine. There are also fairing inside of the exhaust, making the gas flow less turbulent when exiting.

At the end of the exhaust, the gas is compressed by the convergent propelling nozzle.

1.3.6b Material

The gasses entering the exhaust are approximately between 550 and 850 degree Celsius. Therefore,

a specific material has to be chosen that can resist these high temperatures and prevents heat con-

duction to the structure of the aircraft. The chosen material ideal for the exhaust is a nickel based

alloy. To prevent the heat conduction an insulated fibrous material sandwiched between thin layers

of stainless steel is mounted around the exhaust.

1.4 Engine Properties In the process of designing an engine, a clear explanation of the different performance properties is

needed (1.4.1). An engine is designed to have high efficiencies and efficient propulsion. Apart from



the equations used for calculating different efficiencies and forces of the engine, environmental fac-

tors will also be explained. This will be done by explaining the noise and emissions (1.4.2) and the

equations used while calculating.

1.4.1 Performance When a gas turbine engine is designed, the functions of the different parts are factors for creating an

engine with good or poor performance. To achieve high efficiencies, the stages of the engine will be

examined with the suited equation of efficiency (1.4.1a). For the propulsion of the engine and its

efficiency the speed and the mass of the air are primary (1.4.1b).

1.4.1a Efficiency

To have an engine with a high efficiency, each stage has to be improved and analysed. For each stage

an efficiency equation is designated. The equations are given per engine station. An example of en-

gine stations is visible in a stage overview figure (Calculation document). Each number indicates a

position inside the engine between the forward and aft. The efficiencies that will be handled are:

1. Inlet efficiency

2. Fan efficiency

3. Low/High pressure compressor efficiency

4. Combustion efficiency

5. High/Low pressure turbine efficiency

6. Nozzle efficiency

7. Thermal efficiency

8. Thermodynamic efficiency

9. Mechanical efficiency

10. Propulsion efficiency

ad 1 Inlet efficiency

For the inlet efficiency two types of efficiencies are used. One of these efficiencies are Isentropic

efficiency, which is used for temperature rises. This is the efficiency between the ambient tempera-

ture (Tam) and the fan inlet temperature (T’01). However, due to a pressure rise from the divergent

inlet, the actual fan inlet temperature is higher (T01). The ram efficiency is used for the relation of the

ambient pressure (Pam), the inlet pressure (P00) and the fan inlet pressure (P01) (Calculation document

equation 1-2).

ad 2 Fan efficiency

The fan efficiency can be calculated by using the relation between the fan inlet temperature (T01) and

the fan outlet temperature. The fan outlet temperature can be subdivided in an actual (T02) and

theoretical (T’02) temperature. The theoretical temperature can be calculated by multiplying the ac-

tual inlet temperature with the pressure ratio of the fan with the added factor of the Laplace con-

stant (Calculation document equations 3-4).

ad 3 Low/high pressure compressor efficiency

The low/high pressure compressor (LPC/HPC) efficiency can be determined by use of the relation

between the compressor inlet (Ta1/2) and outlet temperature. The compressor outlet temperature

can be subdivided in an actual (Tb1/2) and theoretical (T’b1/2) temperature. The theoretical tempera-

ture can be calculated by multiplying the compressor ratio with the compressor inlet temperature

(Calculation document equations 8-10).

ad 4 Combustion efficiency

Due to a lack of reliable sources an equation could not be found. Therefore, all combustion efficien-

cies that will be used will be between 90 percent and 98 percent, this is reasonable in modern air

transport.

ad 5 High/Low pressure turbine efficiency

For the high and low pressure turbine efficiency, temperature is the main factor. The relation be-

tween the turbine inlet temperature (TIT) (Ta3/4) and the outlet temperature are used. The outlet



temperature can also be subdivided into an actual temperature (Tb3/4) and the theoretical tempera-

ture (T’b3/4) (Calculation document equation 11-12).

ad 6 Nozzle efficiency

In this project the engine will only be exposed to unchoked conditions, thus the exhaust pressure is

always lower than the ambient pressure. The nozzle efficiency is subdivided into two parts, the fan

nozzle and core nozzle. The equation used for the fan nozzle gives the efficiency of the fan duct with

the relation between the fan outlet temperature (Ta5) and the duct exhaust temperature (Tb5). The

core nozzle efficiency is the relation between the nozzle inlet temperature (Tb6) and the core outlet

temperature (Tb6). The both outlets can be subdivided into an actual and a theoretical temperature

(Calculation document equation 13-14).

ad 7 Thermal efficiency

The efficiency of the entire engine can be seen as the thermal efficiency. A theoretical approach can

be made by not taking losses into account. This efficiency is determined by taking the compressor

ratio (�� = ��) and the Laplace constant into account (Calculation document equation 17).

ad 8 Thermodynamic efficiency

The thermodynamic efficiency is much like the thermal efficiency. However, the different losses are

taken into account. Here the efficiency of the turbine and compressors are used along with the com-

pressor and temperature ratio of the compressor and turbine. (Calculation document equation 18)

ad 9 Mechanical efficiency

Due to a lack of reliable sources an equation could not be found. Therefore, all mechanical efficien-

cies that will be used will be 65 percent, because there are also losses of the gear box.

ad 10 Propulsion efficiency

The propulsion efficiency can be defined as the relation between the exhaust speed or jet speed (Cj)

and the actual speed of the aircraft, which can be seen as the speed at the inlet of the engine (Ci).

The exhaust speed equation is used in the case of an unchoked situation. (Calculation document

equation 19-20)

1.4.1b Propulsion

To calculate the propulsion the fan thrust and core thrust are needed. On calculating the main fac-

tors are the speed of the exhaust and the speed of which air enters the engine. The mass flow of the

air through the engine will also play an important role. (Calculation document equations 22-23) By

added the two thrusts the total thrust of the engine is acquired. From this it is clear that the greater

the mass flow in the engine is the greater the thrust. The final factor in designing an engine is that of

its fuel consumption. To derive the fuel consumption it the fuel/air ratio has to be defined. (Calcula-

tion document equation 24) With the ratio it is clear to see the amount of fuel used per unit of air.

With the mass flow of the air it is then possible to determine the fuel consumption.

1.4.2 Environment An added aspect to an engine is that of its environmental influences. One being the generated noise

(1.4.2a) and one the gas emissions (1.4.2b). These influences will be examined and the coherent

equations will be given.

1.4.2a Noise

Noise is produced by practically all the components of the engine (appendix XXIX). This is prominent

in the four stages of the engine: the intake, compression, combustion and exhaust. Beginning with

the intake, as seen on the Fokker 100, a turbo fan is used. Noise is generated by the fan blades as

they rotate. By reason of the displacement of the air at the trailing edge of the blade and the tip

speed of the blade, noise is emitted. As a blade rotates it leaves a turbulent wake zone in which the

following blade will be propelled, this will in itself also produce noise. The noise generated at the

compressors are similar to that of the fan blades. In the combustion stage, noise is emitted as the air

is burnt. Because this noise is in the core of the engine, no necessary solutions for the noise is



needed. Finally, the largest noise producer of the engine is the exhaust. The noise can be traced to

the mixing of the fast jet air flow with the atmosphere.

Various solutions have been designed to reduce the fan made noise. Firstly, this is done by limiting

the amount of blades in the fan. With reduced blades the weight and cost will decrease. Also, the air

that is propelled backwards into the compressors will help in improving the characteristics of a com-

pressor surge. However, this will be at the cost of the fan efficiency, since more blades would answer

in a higher efficiency. Another solution is by increasing the space between the blades, by doing so the

turbulent wake of the one blade will have less effect on the upcoming blade. Yet another solution is

by increasing the spacing between the blades that rotate and the stationary blades, that are used for

the correction of air molecules so that these will enter the compressors at the right angle. This will,

however, lead to a longer engine. By lengthening the intake duct and using sound absorbing materi-

als, further noise reduction at the intake is received. Solution for the noise emitted by the compres-

sor is by increasing the spacing between the rows of blades and vanes. Another solution seen in

modern engines is the use of two compressor stages; a low pressure compressor and a high pressure

compressor. In the first compressor air is only partly compressed to a higher pressure and then at a

later stage compressed to a much higher pressure. The latter being done further in the core of the

engine. Finally, solutions for the exhaust noise have been made. The main way to reduce the noise is

by increasing the mixing rate of the exhaust air with the atmosphere. One way is by using High by-

pass engines. Seeing that the engine of the Fokker 100 is a low bypass engine other methods are

used. One is the Lobe-type nozzle (appendix XXX). This type increases the amount of jet exhaust

openings. Creating smaller jet openings and increasing the contact surface, the exhaust air will decay

at a much higher rate and mix with the atmosphere. The corrugated nozzle (appendix XXXI) is also a

manner of reducing noise from the exhaust. With this design air flows freely into the jet exhaust, this

brings rapid mixture of the jet air. However, by using these means the total thrust is reduced. An

overview of the methods used in the Fokker 100 to reduce noise is shown in (appendix XXXII). A

combination of a lobe typed exhaust, a longer intake duct and two compressor stages, are used to

reduce noise on the Fokker 100 aircraft.

To convert the noise into a numeral factor the perceived noise level is used (appendix XXXIII). This

factor gives an indication of the noisiness of the engine. To calculate this, a factor of the perceived

noise (PN) of the engine is needed. The log of this factor is then divided by the log of two and multi-

plied by ten. In addition the standard noise factor is added, this being 40. This can be seen as the

average noise made by speech, resulting when PN=1.

The International Standardization Organisation (ISO) uses the Effective Perceived Noise Level (EPNL)

unit for noise created by aircraft on the ground. These factors are given in Decibel, thus resulting that

the current unit is the Effective perceived Noise in Decibel (EPNdB). The Fokker 100 has been certi-

fied with an EPNdB of less than 90dB. This would mean that the PN of the current Fokker 100 is less

than 32.

Another factor used to determine the noisiness of an aircraft, according to Lighthill’s theory, is the

Radiated Sound Intensity (appendix XXXIV). Here is seen that the intensity is dependent on the air

density, the diameter of the exhaust nozzle and the velocity of the outgoing gas.

1.4.2b Emissions

The gas emission of an engine is inevitable. This is the result of the chemical combustion in the com-

bustion chamber of the engine. With the ignition of the air with fuel, gasses are released: Nitric Ox-

ides (NOx), which are Nitrogen Oxide (NO) and Nitrogen di-Oxide (NO2), Carbon mon-Oxide (CO),

Carbon di-Oxide (CO2), Carbon (C), Hydrocarbon (CHn) and Water (H20). The reason for this gas emis-

sion falls outside this project and will not be handled. In the air industry a dimensionless term quanti-

fying smoke emissions is used to illustrate the amount of gas emissions. This is measured during the

testing of the engine in the different operating conditions, such as take-off, landing and taxiing. In

this project only a concept design is made and therefore, the smoke number will not be handled.

Various solutions have been made to decrease to output of these gasses. However, the gas emission

can never ultimately reach zero. One of the solutions is by injecting water or steam into the combus-

tion chamber. This will decrease the heat of the flame. This reduces NOx gasses by 40 percent. How-



ever, the Carbon gasses are increased. Another option is the dry low NOx design. This design is made

without the need of water, thus dry. With this design comes two possibilities: rich burning quick

quenching design and the lean burning design. With the rich burning design, in the primary stage the

fuel is burnt with the air, which may lead to smoke in the primary stage. Then the burnt air would be

mixed with large amounts of air in the second stage, diluting the combustion. Unfortunately, the

amount of carbon monoxides and dioxides will be greater. With the lean burning design air is mixed

with the fuel prior to combustion. This is widely used today. But with these designs the complexity of

the engine control is increased while the thrust received is lowered.

To estimate the gas emissions the following basic chemical equation is used (Equation 1).

FUEL +AIR → HEAT + WATER + CARBON DIOXIDE + NITROGEN 1

With this equation it is evident that if fuel mixed with air is burnt, water, carbon dioxide and nitrogen

are the direct products. But in reality, the entire first compound will never fully react in the combus-

tion. This will leave small traces of unburned carbon and even unburned fuel or hydrocarbon in the

combustion. The chemical equation is used to determine the amount of elements in a reaction (ap-

pendix XXXV). In the equation the small traces of Carbon and fuel are not taken up. With this equa-

tion a link to the emission of engines can be made. The molecular mass of the needed molecules are

given in (appendix XXXVI). Here the conclusion can be drawn that with each chemical combustion X

amount of CO2 and �� amount of water and 3.76 �� + �

�� amount of Nitrogen. With the amount of

Litres that will be used, the density of the fluid and the molarity of the compound, the amount of mol

can be calculated (Equation 2).

MOLARITY = AMOUNT OF MOL * (AMOUNT OF LITRES-1

) 2

By multiplying the amount of mol by the molecular mass of the element the amount of grams can be

calculated (Equation 3).

MOL = AMOUNT OF GRAMS * (MOLECULAR MASS-1

) 3

Once all the factors before each element is made equal one can see how much grams of an element

is emitted from the compound.

1.5 Demands

The new chosen engine must have certain specifications to meet with the demands of Fokker Aircraft

(1.5.1). However, the aircraft must be airworthiness; therefore the engine must satisfy by the de-

mands of International Civil Aviation Organization (ICAO) and European Aviation Safety Agency

(EASA) (1.5.2).

1.5.1 Client Demands Fokker Aircraft had given Engine Incorporated the assignment to design a new gas turbine engine for

the Fokker 100 next generation. But Fokker aircraft has given Engine Incorporated a few demands

that they have to bear in mind. Because Fokker aircraft wants to improve the properties of the en-

gine, for example the efficiency and the emission, the fowling requirements will be discussed: 1. Thrust

2. Efficiency fuel consumption

3. Certification

ad 1 Thrust

The new gas turbine engine must be able to give the Fokker 100 enough thrust to departure and land

from existing airports with a maximal elevation of 9000 ft. This must be done with different weather

conditions. The temperature of the OATlimitations(Outside Air Temperatures) for take-off must be 35 or

45 degrees celcius. Also the thrust on MSL at standard day conditions must be calculated (appendix

XXXVII). The minimum thrust at a elevation of 9000 ft is 51065,5 Newton. At MSL the TAS en ρ are



different, but the CL and CD are the same compared to the values at 9000 ft. Therefore the thrust at

MSL is the same like the thrust at 9000 ft.

ad 2 Efficiency fuel consumption

The current Fokker 100 has poor fuel consumption therefore; Fokker Aircraft demands that the new

engine has a fuel consumption with a requirement range of more than 1500nm without refueling.

This must be achieved with a maximum operating speed of 320 knots / M 0,75 and at 35000 feet

(appendix XXVIII).

ad 3 Certification

European Aviation Safety Agency (EASA) demands that the Fokker 100 next generation must be certi-

fied again, to be airworthy. The certification process of a gas turbine engine must be worked out and

which influence the process has on the total certification of the aircraft. Certification, emission and

maintenance must be worked out in the report.

1.5.2 Demands Authorities An aircraft must be airworthiness before it is allowed to fly, therefore the authorities has set up de-

mands for the aviation. For the certification of a new engine, ICAO controls the emission of the en-

gine (1.5.2a). EASA has also demands for designing a new engine and has that divided in; the installa-

tion of an engine, safety and emission and uses the documents CS-25, CS-E and CS-34 (1.5.2b).

1.5.2a International Civil Aviation Organization

The International Civil Aviation Organization (ICAO) states the regulations regarding the emissions of

subsonic jet aeroplanes in the chapter Environmental Protection of the Annex 16. The regulations are

based on the measurement of the noise emissions on several noise measurement points. According

to the Annex 16, there are three types of measurement points: Lateral Full power, Flyover and Ap-

proach measurement point.

The Lateral Full Power is the point on a line parallel to and 450 meters from the runway centre line,

where the noise levels are at the maximum during take-off. These noise levels may not exceed 95

EPNdB (Effective Perceived Noise decibels). Flyover measurement point is the point on the extended

centre line of the runway and at a distance of 6500 meters from the start of roll. Here must the noise

levels not exceed 90 EPNdB. The final measurement point is the Approach measurement point. This

is the point on the extended centre line of the runway, 200 meters from the threshold and in the air

with the position on the ground corresponding to 120 meters vertically and 300 meters beyond the

threshold below the three degree decent path. At this point the noise levels must stay below 100

EPNdB.

For the certification of a new engine, several emissions must be controlled. The controlled emissions

are the smoke and the gaseous emissions coming out of the engine. Gaseous emissions can be di-

vided in unburned Hydrocarbons, Carbon mon-Oxide and Oxides of Nitrogen. The smoke can be

measured in terms of Smoke Number and the gaseous emissions shall be measured in grams.

1.5.2b European Aviation Safety Agency

The EASA operates by the use of Certification Specifications (CS) for aircraft flying in European air-

space. For this report, the following Certifications will be used:

1. CS-25

2. CS-E

3. CS-34

The installation, the compatibility and cooling should operate safely. This is achieved by the use of

the CS-25. Furthermore, the accessories should be working on a safe manner which is connected to

the engine. This is established by the use of the CS-Engines. The engine is not allowed to exceed the

maximum amount of exhaust emission, this is done by the use of CS-34.

ad 1 CS -25

CS-25 says (appendix XXXIX), that every power plant installation should include each component

that is necessary for propulsion. Furthermore, the installation must be accessible for necessary in-



spections and maintenance. For safe manners, every power plant must be arranged and isolated

from each other. Because of this, failures or malfunctions of any engine will not prevent the contin-

ued safe operation of the remaining engines.

The power plant cooling provisions must be able to maintain the temperatures of power plant com-

ponents, and engine fluids within the temperature limits.

Every engine should contain of an air intake system. This system should supply air under each operat-

ing condition for which certification is requested. At the air intake there must be methods to prevent

hazardous quantities of fuel leakage or overflow from drains, vents or other components of flamma-

ble fluid systems from entering the engine air intake. Because of this, the aeroplane must also be

designed to prevent water or slush on the runway, taxiway or other airport operating surfaces from

being directed into the engine air intake.

For the exhaust system applies, that the system must ensure safe disposal of exhaust gases without

fire hazard or carbon monoxide contamination in any personnel compartment. Each exhaust system

component must be ventilated to prevent points of excessively high temperatures.

ad 2 CS-E

The Certification Specification for engines (appendix XL) contains airworthiness specifications for the

issue of type certificates, and changes to those certificates, for engines. CS-E contains the specifica-

tions for the approval for use of the Engine.

The engine, following by CS-E, should be free from dangerous surge and instability throughout its

operating range of ambient and running condition within the air intake pressure and temperature

conditions. All the Engines must be equipped with an igniter system, suitable for starting the Engine

on the ground and in flight, at all altitudes up to a declared altitude. The major rotating components

of the Engine must have adequate strength to withstand both the thermal and dynamic conditions of

normal operation and any excessive thermal or dynamic conditions that may result from abnormal

speeds, abnormal temperatures or abnormal vibration loads. Where bleed air is used to cool or to

pressurize areas of the Engine, the functions of which could be detrimentally affected by the ingress

of foreign matter, the design must be such that the passage of foreign matter of unacceptable quan-

tity or unacceptable size is precluded.

There are also several operating limitations of power, rotational speed, turbine entry temperature,

oil temperature, etc. This is established by means of the endurance tests.

ad 3 CS-34

CS-34 (appendix XLI) stands for Certification Specifications for Aircraft Engine Emissions and Fuel

Venting. This CS-34 references to Annex 16 volume II written by ICAO. The aircraft must, under CS-

34, be designed to comply with the applicable emission requirements. The emission of an engine

contains Hydrocarbons (HC), Nitrogen (NOx) and Carbon monoxide (CO). The output smoke will be

measured and reported in the form of a Smoke Number. Guidance material for the application of

the certification specification for aircraft engine emissions can be subdivided in:

- For instrumentation and measurement techniques for gaseous emissions, the attachments to

Appendix 3 of ICAO Annex 16, Volume II; and

- For instrumentation and measurement techniques for gaseous emissions from afterburning

gas turbine engines, the attachments to Appendix 5 of ICAO Annex 16, Volume II.

The mass, als known as Dp, of the gaseous emissions will be measured and reported in grams. Fur-

thermore, there is the rated output to concern about. This is the maximum thrust available for land-

ing and take-off under normal conditions at ICAO Standard Atmosphere (ISA) sea level. The rated

output is calculated in kilo Newton’s.

1.6 Next generation engines The engine on the current Fokker 100 is not up to the standards. In the past years enormous innova-

tions have been done concerning the gas turbines. The turbine engines nowadays can stand a higher

temperature, produces less noise and has a low emission. For reaching these goals, changes have

been made in the properties of the gas turbine (1.6.1). Not only the components in general were

changed, also an innovation in materials choice took place (1.6.2).



1.6.1 General gas turbine innovations The next generation engines are designed in such a way that the environmental impact of the engine

is as low as possible. This means that during engineering a next generation engine low fuel consump-

tion (1.6.1a), emission ratings (1.6.1b) and noise nuisance (1.6.1c) are central issues.

1.6.1a Low fuel consumption

The fuel consumption is related to several things and one being the effectiveness of the fans. If the

fans can move a large a amount of air at a relatively slow speed contributes to a better fuel con-

sumption. These goals are accomplished by making the fans more aerodynamically enhanced and

materials are constantly improved in relation to the weight.

1.6.1b Emission ratings

To increase the performance of the engine, the engine is constantly improvements are made. These

improvements sometimes lead to a higher temperature in the combustion chamber. The material of

the combustion chamber limits the temperature. If the temperature is not high enough the fuel

combustion will not be total, resulting in a high emission. So some new engines are equipped with a

tile layer. This tile layer is able to cope with the higher temperatures and by using the tiles, the tem-

perature is able to increase. This ensures a better burning process. The disadvantage of the method

is that it substantially increases the weight of the engine.

1.6.1c Noise nuisance

Noise nuisance is a great factor in the aviation branch. The noise created by an aircraft is for a great

part accountable to the fans. There are some new technologies in the noise reduction area. The GE90

engine is currently installed on the Boeing 777.The engine fans are large in diameter that they propel

such an amount of air that the fans are able to spin at a lower rate and thereby lowering the amount

of noise created by the fans. The air flow through the engine also has a lower speed now and the

mixing of the hot air with the atmosphere will not go as roughly as before. This results also in lower-

ing the noise created by the mixing of the air.

1.6.2 Material innovations A change in material choice brings a lot of new properties for the components. The past decade sev-

eral innovations were done to improve the material properties such as temperature resistance. For

these material innovations three main stages of the gas turbine will be researched. The compressor

material significantly increases the strength (1.6.1a). The main reason to innovate new materials for

the combustion chamber is the increasing temperature (1.6.2b). Also the new materials used for the

turbine can stand way more temperature stresses than the material used in the past (1.6.2c).

1.6.2a Compressor

In the earlier days aluminum was used, but because of the high temperatures titanium was added.

Nowadays the material Ti6Al4V is used for compressor blades, which means that a titanium alloy is

used with six percent aluminum and four percent vanadium. The aluminum has the purpose to

strengthen the blades and the vanadium is there to moderate the high temperatures.

1.6.2b Combustion

For the combustion liner several changes are made. This is because the temperature in the combus-

tion chambers of nowadays is higher than a couple of years ago. Also the combustion has to meet

with several new environmental desires such as lower emission and fuel burn. The original combus-

tion liner was made of AISI 309 stainless steel. But this design cannot fit the desires of the engines

anymore. The temperature in the combustion liners has increased during a few years and that is why

HS-188 has recently been used in the combustion. HS-188 is an alloy of cobalt, nickel, chromium and

tungsten. The advantage of this alloy is that it is designed for high temperature resistant. Also this

material is oxidation resistant for temperatures under the 1095 degree Celsius and significantly in-

creases the creep rupture strength. Besides that, also a thermal barrier coating (TBC) is used. At

some gas turbines TBC is added and it showed a great performance.



1.6.2c Turbine

Because the turbine blades have to stand high temperatures, the material tantalum is used. In the

past the material titanium was used but tantalum has higher temperature strength and a better oxi-

dation resistance. Recently the alloy is used in the single crystal blade. This controls the elastic prop-

erties and therefore makes sure that the natural vibrations in the blade are controlled. Another ma-

jor change is the added tungsten and rhenium to nickel making it a nickel alloy. This ensures high

solution strength. Finally, ceramic coatings are used to operate at high temperatures and will im-

prove the engine performance. This thermally deposited ceramic coating improves propulsion and

power generation.

1.7 Function research The function of the gas turbine can be described in a function research. The function research is

based on how the gas turbine receives the proper amount of air and converts it to the necessary

thrust. The main function of the gas turbine is to create sufficient thrust to propel the aircraft for-

ward. This function can be described in several sub-functions. The gas turbine has eight types of sub-

functions:

1. Air intake

2. Fan blades

3. Bypass

4. Air compression

5. Gas combustion

6. Gas expansion

7. Gas acceleration

8. Engine control

ad 1 Air intake

The air intake is where the input starts of the gas turbine. The air entering the gas turbine is de-

creased in velocity and increased in pressure.

ad 2 Fan blades

The air in the air intakes is divided by the fan in the bypass air and in the compressor air. The func-

tions of the fan are to compress the air going to the compressor and tip of the fans increases the air

flow velocity of the air entering in de bypass.

ad 3 Bypass

The bypass functions are to provide cooling to the core and to mix the cold and hot air in the exhaust

to reduce the noise emissions. The bypass converts the increased pressure from the fan into kinetic

energy to provide the needed thrust.

ad 4 Air compression

The air coming in the gas turbine gets compressed by the compressor. The potential energy of the air

increases the pressure and temperature of the air.

ad 5 Gas combustion

The compressed air is mixed with fuel and ignited in the combustion chamber. Igniting this mixture

results in a greater pressure, that is needed to create the amount of thrust.

ad 6 Gas expansion

Gasses coming out of the combustion chamber will be lowered in pressure by the turbine. This re-

duction of pressure will produce kinetic energy to propel the compressor.

ad 7 Gas acceleration

The final stage and the output of the gas turbine is the exhaust. Here, the gasses decrease to the

atmospheric pressure, creating the thrust needed to propel the aircraft.



ad 8 Engine control

The function of the engine control is to monitor the engine performance and to prevent common

engine failures such as compressor stall.

2 Analysis gas turbine parts When designing a new gas turbine the different stages of the gas turbine must be analysed. The

analyses of these parts are been described in a morphological overview (2.1). After analyzing the

parts of the gas turbine, three types of gas turbines are chosen based on noise emissions, fuel con-

sumption and gas emissions (2.2). The calculation of the three types of gas turbines concerning the

efficiencies and performance in various situations will be described (2.3). Each chosen design will be

compared in a pro and con diagram (2.4). The design with the highest score from the pros and cons

diagram will become the new design of the gas turbine (2.5).

2.1 Morphological overview An engine depends on “two fuels” for the engine to work, air and the actual fuel. The engine consists

of five stages. The air passes the air intake (2.1.1) into the engine. From there on the air enters the

compression stage, air is compressed to a certain pressure (2.1.2), fuel is added to the air and this

mixture is ignited in the combustion chamber (2.1.3). This ignited air/fuel mixture powers the tur-

bines (2.1.4) and eventually leaves the engine through the exhaust nozzle (2.1.5).

2.1.1 Air intake The air intake is an opening in the cowling of the engine. Through this opening air enters the engine.

The purpose of the air intake is to minimize drag and to provide an undisturbed air flow to the com-

pressor. By doing this, the burning process of the engine is optimized. The airflow entering the en-

gine cannot flow at a supersonic speed, this would damage the fan blades. To keep the speed of the

air below Mach one the inlet is shaped divergent for subsonic flight.

The air flow that enters the engine is highly sensitive to changes in the shape of the inlet. Ice could

form on the edges of the inlet and this could lead to a sharp edged inlet. This would disturb the air-

flow. That is why the front of the air intake is heated to prevent ice from forming. The heating system

determines in some way the material of the inlet. The used material needs to be able to conduct

warmth. The front of the engine is made out of aluminium.

2.1.2 Fan blades Various suitable fan blades, that provide sufficient air flow to the compressor and provide the re-

quired thrust, will be given in this paragraph. This will be done by giving the different types of fan

blades (2.1.2a) and materials (2.1.2b).

2.1.2a Types of fan blades

There is a range of different fan blades used in a gas turbine. The following options will be described:

1. Snubber

2. Curved

3. Wide cord

ad 1 Snubber

This type of fan blade has a snubber incorporated in its design. The snubber is a support on the blade

surface to maintain the required distance between the fan blades. This support will prevent move-

ments of the blades that can cause aerodynamic instabilities. Unfortunately, the snubber adds

weight to the blade and creates a turbulent wake which can lead to airflow disturbances in de com-

pressor.

ad 2 Curved

The shape of the curved blade is made in such a way, that the lower end of the blade creates enough

pressure to enter the compressor. But the upper part of the blade provides enough airflow through

the bypass to create the thrust needed. These curved blades are also debris tolerant making the en-

gine more reliable.



ad 3 Wide cord

The wide cord fan blade is a fan blade with a stretched cord that gives the blade an added strength

with less weight. This stretched cord will ensure the stability of the blade by minimizing movement of

the blade during rotation. With this design the use of snubbers is obsolete.

2.1.2b Materials

The weight of the fan blade is essential for withstanding the out of balance forces which occur during

rotation and in a possible blade failure. The basic material used in fan blades is titanium. Titanium is a

metal which has a high strength and is relatively ductile, making this a suitable metal to be used in

the rotating fan blade. The use of Carbon is also an option; unfortunately the flexibility of the mate-

rial greatly reduces the effectiveness of the blade. A modern technique used in modern aircraft is a

blade made out of a honeycomb core which is sandwiched between two titanium plate skins.

2.1.3 Bypass

The bypass ratio of the engine depends on the thrust which is required by the engine. The bypass

ratio can be divided into three groups:

1. Low bypass ratio

2. Intermediate bypass ratio

3. High bypass ratio

ad 1 Low bypass ratio

An engine with a low bypass ratio, has a relative small amount of air which is bypassed around the

hot core of the engine. The low bypass ratio results that the core needs to produce less power to

drive the fan. An engine which has a low bypass ratio has a ratio which is lower than 3:1. This bypass

ratio means that 75 percent of the air particles are separated for bypass air and 25 percent of the

particles enter the core. This low bypass ratio however results in a high exhaust gas temperature.

ad 2 Medium bypass ratio

An engine which has been equipped with a medium bypass ratio has a ratio which lies between 3:1

and 5:1. The engine properties will lie between the low and high bypass engine.

ad 3 High bypass ratio

An engine with a high bypass ratio has a relatively high amount of airflow which passes the engine by

the bypass duct. A high bypass engine has a minimum bypass ratio of 5:1 (83 percent bypass air 17

core air). This high bypass ratio design will cool the exhaust flow behind the engine which will reduce

the noise. The diameter of the fan of a high bypass engine will be larger. This will result in higher

temperatures inside the engine and more fuel consumption but at the same time more thrust.

2.1.4 Compressor The various design options for a suitable compressor will be given in this paragraph. This will be done

by giving the different types of compressors (2.1.4a) and possible material use (2.1.4b).

2.1.4a Types of compressors

The gas turbine engine has a range of different compressor types for its operation. The following

optional compressor designs will be enlightened:

1. Single spool axial compressor

2. Twin spool axial compressors

3. Triple spool axial compressor

ad 1 Single spool axial compressor

The axial flow compressor compresses the air in axial lengths. This is done in several compression

stages, the amount relying on the choice of the designer and the expectations. With this compressor

the design will have the following benefits: a small frontal area, multistage possibility which increases



Compressibility Ratio (CPR), straight-through flow and allowing high ram efficiency. Unfortunately,

this design is heavy and requires a high starting power to initiate rotation.

ad 2 Twin spool axial compressor

The double axial compressor is much like the single, only the compression stage is divided into a low

and a high compressor stage. With this design the air is first partially compressed in the low pressure

compressor. After the air is compressed it travels to the high pressure compressor, here the air is

compressed again into a higher pressure. By the use of these two stages the air is gradually com-

pressed, this reduces the heat and the speed of the air during the compression stage. The air that

enters the combustion chamber is at a lower speed and at a lower temperature, thus resulting in

better combustion efficiency, seeing that the chance of ‘flame out’ is reduced.

ad 3 Triple spool axial compressor

The triple axial flow compressor can be seen as an extended double axial flow compressor. Here the

compression stage is divided into a low, intermediate and a high pressure compressor. The air is

pressurized at a lower rate increasing the combustion efficiency. The more rotating parts in the en-

gine, the more power is necessary to drive these parts. The power used to drive these parts is seen

as power lost in the thrust.

2.1.4b Materials

Because the compressor works under high pressures and sometimes high temperatures, durable

materials are appropriate. A usable material is a nickel based alloy. This alloy has similar properties as

steel while the relative weight of lower. Finally, the use of titanium alloys has grown immense in the

air transport. Titanium alloys have the same properties as that of steel, while having less than half its

weight.

2.1.5 Combustion For the combustion chamber of the new engine several types of combustion chambers can be

used (2.1.5a). The combustion chamber can be built up out of several materials (2.1.5b).

2.1.5a Types of combustors

In aircraft engine engineering there are four main types of combustion chambers in use for gas tur-

bine engines. These types of combustion chambers are:

1. Multiple combustion chamber

2. The tubo-annular combustor

3. Annular combustor

4. Double annular combustor

ad 1 Multiple combustion chamber

The multiple combustion chamber (appendix XLII) is used on centrifugal compressor engines and the

earlier types of axial flow compressor engines. The multiple combustor chamber engines can be de-

signed as straight-through or reverse-flow designs. However, reverse flow multiple combustion

chambers create a considerable pressure loss so straight-through multiple chamber engines are

used. The combustion chambers are disposed around the engine near the compressor. The delivery

of the air from the compressor is directed by ducts to pass into the individual chambers (1). Each

chamber has an inner flame tube around which is placed in an air casing (2). The air passes through

the flame tube narrowing and also between the tube and the outer casing. The separate flame tubes

are all interconnected (3). This allows each tube to operate at the same pressure and also allows

combustion to propagate around the flame tubes during engine starting.

ad 2 Tubo-annular

The tubo-annular combustion chamber (appendix XLIII) bridges a gap between the multiple and an-

nular types. The combustor works with a number of flame tubes (1) which are fitted inside an air

casing (2). The airflow is similar to the multiple combustion chamber. These combustors have the



advantage of the ease of maintainability of the multiple system and the compactness of the annular

system. The tubo-annular combustor also has a better temperature distribution and can be of the

straight-through or reverse-flow design. In most aircraft engines the tubo-annular combustors are of

the straight through flow type. The straight-through flow type tubo-annular combustor requires a

much smaller frontal area than the reverse-flow type tubo-annular combustor. The tubo-annular

combustor also requires more cooling air flow than a multiple or annular combustor because the

surface area of the tubo-annular combustor is much greater. The tubo-annular combustor has a

more even combustion because each can has its own nozzle and a smaller combustion zone, result-

ing in a much more even flow.

ad 3 Annular combustors

Annular combustors (appendix XLIV) consist out of a single flame tube (1) which is completely annu-

lar in form and contains in an inner (2) and outer casing (3). The chamber is open at the front to the

compressor and at the rear to the turbine nozzles. While the annular combustor provides a specific

amount of power output, the length of the combustion chamber can be reduced up to 75 percent of

that of a tubo-annular system of the same diameter, this results in savings in weight and production

cost. Another advantage is the elimination of combustion propagation problems from chamber to

chamber. In reference to the tubo–annular combustion system the wall area of the annular chamber

is less and so the amount of cooling air required to prevent the burning of the flame tube wall is less,

by approximately fifteen percent. This reduction in cooling air raises the combustion efficiency which

contributes to virtually eliminate unburnt fuel and oxidizes the carbon monoxide to non-toxic carbon

dioxide, thus reducing air pollution. On the other hand, the annular combustor is much harder to get

to for maintenance and tends to produce a less favourable radial and circumferential profile as com-

pared to the tubo-annular combustors. The introduction of the air spray type fuel spray nozzle to this

type of combustion chamber also greatly improves the preparation of fuel for combustion by aerat-

ing the over-rich pockets of fuel vapours close to the spray nozzle; this results in a large reduction in

initial carbon formation.

ad 4 Double-annular combustor

The next step in combustors may be a double-annular chamber (appendix XLV). This chamber con-

sists out of two concentric flame tubes (1) and a double-annular shaped combustion chamber (2).

This dual annular design is pursued in order to obtain lower emissions. These lower emissions are

achieved by using two rows of burners. The double-annular combustor can use just a single row dur-

ing low power (idle, descent) conditions, while both sets are lit under high power (take-off, climbout)

conditions. This considerably reduces NOx emissions by 40 to 50 percent. With the double-annular

combustion chamber it is also possible to lower the emission of carbon monoxide (CO). With a

greater volume of the combustion chamber, the burning process can be more rapid and efficient.

2.1.5b Materials

There are many materials which can be used for the combustion chamber such as: SS309 (stainless

steel), Hast X (iron, chromium, molybdenum alloy), N-263 and HA-188. Because of investigations and

studies to materials, one material has properties which are better in reference to other materials.

The material which can best be used in the combustion liner is HA-188. This material is a Chromium

(CR) – Ni-based alloy which improves creep rupture strength and allows higher temperatures. The

material in the combustions chamber is also provided with a Thermal Barrier Coating (TBC), it pro-

vides the liner with an insulation layer which has a thickness of about 0.4-0.6mm. This layer can re-

duce the metal temperatures by 50-150°C. The primary benefit of TBC is to provide an insulating

layer that reduces the underlying base material temperature and mitigate the effects of hot streaking

or uneven gas temperature distributions. The HA-188 alloy and the thermal barrier coating contrib-

ute to a higher reliability of the combustor.

2.1.6 Turbine The turbine gets air from the combustion chamber with a high temperature. In the turbine mechani-

cal energy is obtained and used to drive the compressor. In the turbine the velocity of the air will



increase. One of the turbine types is the radial inflow turbine. This turbine has to types; cantilever

turbine and the mixed flow turbine. These types will not be discussed because of their low efficiency.

Another type of turbine is the axial flow turbine. These types of turbines will be discussed (2.1.6a).

At last, a few innovations, as discussed already, have been done concerning the materials. These

materials make sure the turbine blades can stand the high forces and temperatures (2.1.6b).

2.1.6a Types of turbines

As said before, only the types of the axial flow turbines will be discussed. There are three types of

axial flow turbines:

1. Impulse turbine

2. Reaction turbine

3. Impulse-reaction turbine

ad 1 Impulse turbine

The fixed nozzle guide vanes of the turbine will make the pressure drop at each stage. These fixed

nozzle guide vanes are shaped convergent. Therefore, the velocity is very high when entering the

rotor. The direction of the gas will be changed and the gas will go to the turbine blades. An impulse

force is created by the gas at the blades.

ad 2 Reaction turbine

In this type of turbine the blades are airfoil shaped. When the air enters the blade section there will

be at the upper side of the blade a decrease in pressure. At the lower surface of the blade an in-

crease of pressure will occur. This will cause a lift force at the blades. This force will make the blades

rotate. By decreasing the pressure at the upper surface of the blade, the velocity will be increased.

Also at these blades the air will cause an impulse force on the blades.

ad 3 Impulse-reaction turbine

The impulse turbine blade and the reaction turbine blade can be combined in an impulse-reaction

turbine. At the root of the blade the turbine blade is shaped like an impulse turbine blade and at the

tip the shape of the blade is the same as the shape of the reaction turbine blade. Mostly the gas tur-

bine then consists of 50 percent impulse and 50 percent reaction.

2.1.6b Materials

For turbine blades nickel based alloys are used in combination with a ceramic coating. This ensures

that the turbine blade can stand high temperatures. One of the last innovations in the materials for

turbine blades is the replacement of titanium by tantalum. In the material IN 738 and GTD 111 this

tantalum is used. Both materials have nickel as main material and a high percentage of chrome.

2.1.7 Exhaust nozzle The exhaust nozzle is an important part of the engine. It controls the airflow which leaves the engine.

There are different types of configurations of exhaust nozzles. For this report the project group only

concentrates on a main type (2.1.7a) and second on the group of ‘low noise’ types (2.1.7b) and the

materials that could be possibly used (2.1.7c).

2.1.7a Main type

This type is the most common. This type has a convergent nozzle that accelerates the air coming

from the turbines. There are no limitations concerning the noise or emission; so the fuel efficiency is

not affected.

2.1.7b Low noise type

There are multiple ways to reduce the noise created by the accelerated air. Only two are explained

here:

1. Chevron

2. Notched



ad 1 Chevron

A way to reduce noise nuisance is by using a Chevron nozzle. When the hot air from the core engine

mixes with the ambient air, this causes friction. This friction leads to noise nuisance. This nozzle

mixes the cold bypass air with the hot air from the core engine. By mixing the air; the temperature

difference between the mixed exhaust flow and the ambient airflow is smaller thus creating less

noise nuisance.

ad 2 Notched

The notched exhaust nozzle also mixes the air of the bypass flow and the core engine flow. The

notched nozzle is not as effective as the chevron. The notched is less effective in noise reduction but

the thrust capability of the notched nozzle is better.

2.1.7c Materials

The material used for any exhaust nozzle must be able to cope with high temperatures. The core

engine air that enters the exhaust nozzle is around 950 Kelvin. One material that is able to cope with

that temperature is a nickel and tungsten alloy.

2.1.8 Engine control

To control the thrust there are several devices that are installed into the engine. To control the en-

gine there are two ways, both ways use an EEC. Only two are explained here:

1. HMU

2. FADEC

ad 1 HMU

This system uses an EEC that controls the HMU. The HMU measures the right amount of fuel to keep

the fuel air ratio good. If during operation a fault occurs, the HMU switches to manual control.

ad 2 FADEC

This system also uses an EEC but it does not have a HMU. The FADEC has full control over the fuel

management system during all types of operations modes. It monitors and saves engine data to pro-

vide a better overall engine efficiency.

2.2 Possible designs The engine parts which can be used for every stage of the turbofan engine are described and illus-

trated in the morphological overview. The project group has investigated which parts of the engine

contributes to noise reduction, which parts contribute to fuel efficiency and which parts contribute

to lower emissions. With this in mind an engine with reduced noise is composed (2.2.1) an engine

which has low fuel consumption is composed (2.2.2) and an engine which has low emissions is com-

posed (2.2.3).

2.2.1 Design One: Noise reduced engine The first design which will be taken into account will be the noise reduced engine (appendix XLVI).

The type of inlet which is used for the engine is a subsonic inlet. This inlet is specially created to ab-

sorb the noise emitted from the front of the engine. This is done by using a honeycomb material that

absorbs the noise. After entering the inlet, the air will first meet the fan which uses a wide chord

blade. With this, the air is slowed down substantially, while still increasing the pressure. By doing this

the air that enters the engine has a lower velocity which can provided a better fuel burn. This means

that the fan blade can rotate at a lower speed to generate the amount of pressure. This reduces the

friction caused by the fan and reducing noise. The air is then pressurized and separated. A large per-

centage of the air enters the bypass and the rest enters the core of the engine. This means that the

engine is equipped with a high bypass ratio and increases the cooling of the engine and the mixture

rate of the exhaust. This reduces the noise emission. The air that enters the core engine will then

enter the compressor stage. This is separated in a low and a high compressor stage. Here the air will



be further pressurized. Because the air is not compressed immediately over both compressor stages,

the heat is reduced and the flow of the air has less turbulence. The air will then enter the double

annular combustion chamber. In this chamber, the air is burnt with fuel which will increase the vol-

ume. Because of the pre-mixture of the air and gas, this combustion chamber will provide optimal

efficiency. The gas will then leave the combustion chamber and flow through the turbines. The axial

impulse reaction turbines will be used. These turbines use a fifty percent reaction force of the gas to

rotate and a fifty percent of the impulse force to rotate. Hereby, the rotation of the turbine is opti-

mal. After the air leaves the turbine it goes to the chevron nozzle. The chevron nozzle reduces noise

by mixing cold bypass air with the hot air from the core engine. This is done to lower the overall

temperature difference of the exhaust air. This creates less friction and thereby less nuisance. The

entire engine will be monitored by a FADEC system, providing failure prevention.

2.2.2 Design Two: Low fuel consumption engine

Secondly, the low fuel consumption design will be described (appendix XLVI). The type of inlet which

is used for the engines will be a laminar flow intake. With this intake, turbulent flows of the air are at

a minimum. After passing the inlet, the air passes the fan. Here the air will be compressed and sepa-

rated, where a percentage will enter the bypass and the rest into the core engine. The fan blade that

will be used is the curved blade. With this design the air is compressed efficiently to provide enough

pressure to the compressor and to the bypass. The bypass of this design will be an intermediate by-

pass ratio. This will ensure that the air in the bypass is efficiently used for cooling along with thrust.

The air will then be compressed by an axial flow compressor. The compression will be done in two

stages, the low and high pressure stage. This compressor provides the best flow of air to the combus-

tion chamber, which will bring about a better air- fuel mixture. Therefore, less fuel has to be supplied

to acquire the required thrust setting; the mixture is then burnt efficiently. After the air passes the

compressor it will enter the double annular combustor. This double annular combustor uses two

rows of burners which can be used as a single and dual burning row. The single row is used during

idle and descending. With the use of one row instead of two rows it will result in lower fuel con-

sumption. The two rows are mainly used during take-off and climb-out. When the air leaves the

combustor it goes to the impulse –reaction turbine. Because the impulse-reaction turbine provides a

better bending of the airflow due to a better shape of the turbine blade more thrust can be created

with a fuel setting, so less fuel will be used. The air then passes the notched nozzle. The notched

nozzle reduces the noise but makes sure there will be a minimum thrust loss. The entire engine will

be monitored by a FADEC system, providing failure prevention.

2.2.3 Design three: Low emission engine The third and final design will be the low emission engine (appendix XLVI). The type of inlet which is

used for the engines is a subsonic inlet. The FOD tolerant inlet will be used. When the air passes the

inlet and meets the fan, the air will be pressurized and separated. A percentage of the air will enter

the bypass and the rest will enter the core of the engine. The fan blade that will be used is the curved

blade. This blade optimizes the pressure increase of the fan the fuel efficiency is increased. The air

will then be compressed by an axial flow compressor. This compression will take place in two stages,

a low and a high pressure compression. This compressor will provide the best flow of air to the com-

bustion chamber which causes the fuel in the combustion chamber to burn completely and reduces

the amount CO. Next a double annular combustor will be used, this double annular combustor uses

one or two rows of burners as described. By using one instead of two sets of burners 40 to 50 per-

cent of the NOx emissions is reduced. With the use of the double annular combustion chamber it is

also possible to lower the emission of CO. With this larger combustion chamber, the burning process

can be more rapid and efficient. When the air leaves the combustor an impulse reaction turbine is

used, because the fuel usage of this reaction turbine is less, the engine will emit less NOx and CO.

The air then passes the notched nozzle where the air of the core flow gets mixed with the air of the

bypass flow. The entire engine will be monitored by a FADEC system, providing failure prevention.



2.3 Engine performance To know which engine has the best properties of the three chosen engines (noise, fuel efficiency and

emission), a calculation will be made. For each engine an excel document will be made about the

categories of the engine (Calculation document). Each engine will be divided in three conditions:

1. ISA conditions

2. Elevation at 9000ft and hot day conditions

3. Cruise conditions

4. Calculation overview

ad 1 ISA conditions

The calculations are made on the basis of the ISA conditions. These are theoretical atmospheric con-

ditions which are commonly accepted.

The emission engine type has the highest thrust of 76696,02 N and the highest TSFC of 0,6001 lb/lbh.

The noise reduction engine type has a lower EGT rate to further decrease sound ratings. The noise

reduction type has the lowest TSFC, 0,3818 lb/lbh and a thrust of 62004,01 N.

The fuel efficient engine type is the one with the lowest TSFC, namely 0,5445 lb/lbh. The thrust of

this type of engine is 75091,95 N.

ad 2 Elevation at 9000 ft and hot day conditions

The given altitude at which the aircraft have to take off is 9000 ft in hot day conditions with a tem-

perature of 318 Kelvin. By influences of the weather conditions the pressure drops to 0,75 bar. These

two factors both influence the overall thrust.

Also in this situation the emission engine type has the highest thrust of 53204,31 N and the highest

TSFC of 0,5578 lb/lbh.

The noise engine type thrust is at this altitude not sufficient enough because this thrust is lower than

the required minimum thrust. The TSFC is 0,4881 lb/lbh and is the lowest of the three conditions.

At this height the fuel efficient condition has al lower thrust than at MSL, the thrust is 52194.40 N

and is lower because of the conditions at 9000 ft. The TSFC for a fuel efficient engine type at 9000 ft

is 0,5025 lb/lbh.

ad 3 Cruise conditions

Cruise conditions are at an altitude of 35000 ft and are combined with ISA atmospheric conditions at

that altitude. The pressure drops to 0.235 bar and the temperature drops to roughly 219 Kelvin.

The thrust rates of the three engines are all close in this situation. But again the emission engine type

is the one with the highest thrust, 14386.22 N. The noise reduction engine type has the lowest thrust

but excels at TSFC, range and the endurance.

ad 4 Calculations overview

An overview of the calculation results, including the results of thrust, range, endurance and TSFC of

the types of engines in three conditions.

Emission Noise reduction Fuel efficient MSL

Take off

9000

ft/Hotday

conditions

Cruise

height

MSL

Take off

9000

ft/Hotday

conditions

Cruise

height

MSL

Take off

9000

ft/Hotday

conditions

Cruise

height

Thrust (N) 76696.02 53204.31 14386.22 62004.01 30395.60 12126.72 75091.95 52194.40 13833.69

Range

(Nm)

1785.17 1747.28 1769.95 2093.35 1978.92 1956.58 1821.80 1779.97 1806.68

Endurance

(Hr)

14.88 14.56 5.53 17.44 16.49 6.11 15.18 14.83 5.65

TSFC

(lb/lbh)

0.6001 0.5578 1.0488 0.3818 0.4881 0.6708 0.5445 0.5025 0.9690



2.4 Advantages and disadvantages With the concept of three new designs an overview of the pros and cons of each design will be made.

This will filter out one final design, which will be recommended. First a short summary of each design

will be made (2.4.1) and then the pros and cons will be placed in a table to give an overview (2.4.2).

Finally, the design which is most beneficial will be chosen.

2.4.1 The three designs Here a short summary of the specifications of each design will be given. The three design types made

are:

1. Noise reduced engine

2. Low fuel consumption engine

3. Low gas emission engine

ad 1 Noise reduced engine

The noise reduced engine design is a design that focuses on the noise emission of the engine. It uses

noise absorbent materials and a fan blade that ensures an efficient rotation speed. A high bypass

ratio that cools and reduces the core noise emission and a chevron nozzle which gives a high mixing

rate at the exhaust. During ISA conditions, the engine is capable to produce 13921.79 pounds of

thrust at take-off, with a total efficiency of 15 percent, a range of 2093.35 NM and an endurance of

17.44 hours. This design produces 6824.74 pounds of thrust at 9000ft elevation in hot-day conditions

with a total efficiency of 14 percent, a range of 1978.92 NM and endurance of 16.49 hours. During

cruise flight, a maximum thrust can be reached of 2722.82 pound with a total efficiency of 31 per-

cent, a range of 1956.58 NM endurance 6.11 hours.

ad 2 Low fuel consumption engine

The low fuel consumption engine design focuses on fuel efficiency. It uses a laminar inlet to create an

air inlet flow with minimum turbulence. A curved fan blade that optimizes the pressure build up for

the core engine and the intermediate bypass which allows proper cooling in and around the core. A

double annular combustor, that uses pre-mixing of the air and fuel, adds to the efficiency of the

combustion. Also the ability to chose the amount of burners increases its efficiency. This design uses

a notched nozzle at the exhaust which reduces noise while maintaining most of the thrust. This en-

gine is capable to produce 16860.42 pounds of thrust in standard ISA condition at take-off, with a

total efficiency of 13 percent, a range of 1821.80 NM and an endurance of 15.18 hours. Furthermore,

it is capable of producing 11719.23 pounds of thrust at 9000ft elevation in hot-day conditions with a

total efficiency of 13 percent, a range of 1779.97 NM and endurance of 14.83 hours. In cruise flight,

the engine can reach a thrust of 3106.08 pounds with a total efficiency of 28 percent, a range of

1806.68 NM and an endurance of 5.65 hours.

ad 3 Low gas emission engine

The low gas emission engine design focuses on the gas emission of the engine. It uses a FOD inlet

that decreases the chances of foreign objects in the engine. A intermediate bypass is used for effi-

cient cooling in and around the core engine. A curved fan blade ensures an efficient pressurization of

the air for the bypass and the core engine. The double annular combustion chamber is used to burn

the air and fuel mixture efficiently reducing NOx emission. Also by choosing the amount of burners

the amount of emissions is also influenced. A reaction turbine is specially used for having the ability

to be used with “low power”. A notched nozzle is used to ensure a rapid mixture of the bypass air

with the gas air. This design is capable of producing 17220.59 pounds of thrust in standard ISA condi-

tion at take-off, with a total efficiency of 13 percent, a range of 1785.17 NM and an endurance of

14.88 hours. At 9000 feet under hot-day conditions, this engine produces 11945.99 pounds of thrust

with a total efficiency of 12 percent, a range of 1747.28 NM and endurance of 14.56 hours. The

thrust in cruise is 3230.14 pounds with a total efficiency of 28 percent, a range of 1769.95 NM and an

endurance of 5.53 hours.



2.4.2 Advantages and disadvantages overview A table can be made for the design advantages and disadvantages. This overview will be done by

giving each design a rating expressed in the amount of plusses (+), the more plusses the higher the

rating. Each design will be judged on various properties that are essential for the success of the en-

gine, not only on the market but also for the airliner and maintenance crew. The properties that will

be used have been chosen as the dogma properties for the success of the engine. The engine must

give a sufficient amount of reliability to the users while remaining maintainable for workers. It must

be compatible with the current and next gen Fokker100 for installation. The fuel efficiency of each

design, the noise and gas emissions are also key factors for the design. Because the engine will be

exposed to different operation environments, the characteristics of each design are key. The thrust in

the various conditions has been set to a minimum of 11.251,8 pounds. The range that can be

achieved without refuelling has been set to a minimum of 1500 NM. The TSFC is also important; this

has been set to a maximum of 0,951 lb/lbh. These amounts are required by client demands. All

properties that have a minimum or maximum, required by the client or which are properties chosen

as congenital will have a factor of two and be highlighted blue. Below, an overview is made contain-

ing the properties.

Properties Noise reduced engine Low gas emission Low fuel consumption

Overall conditions

Reliability -+++ ++++ -+++

Noise emission ++++ -+++ -+++

Gas emission -+++ ++++ ++++

Compatibility ---+ ---+ --++

Fuel efficiency -+++ -+++ ++++

Maintainability --++ -+++ --++

Take-off in ISA conditions

Endurance ++++ -+++ --++

Range ++++ --++ -+++

Thrust --++ ++++ -+++

TSFC ++++ -+++ -+++

Total efficiency --++ --++ --++

At 9000ft elevation in hot-day conditions

Endurance ++++ -+++ --++

Range ++++ -+++ -+++

Thrust ---- ++++ -+++

TSFC ++++ ++++ ++++

Total efficiency --++ --++ --++

Cruise flight

Endurance -+++ --++ --++

Range ++++ --++ -+++

Thrust --++ ++++ -+++

TSFC -+++ ---+ --++

Total efficiency ++++ -+++ -+++

Each design is capable of receiving a maximum of 104 plusses. By adding all the plusses the following

is obtained:

Gradation of each design

Noise reduced engine 72/104

Low gas emission engine 72/104

Low fuel consumption engine 71/104

This will mean that the low gas emission and the noise reduced engine designs are the most benefi-

cial designs, seeing that these designs have the most advantages.



2.5 Conclusion From the advantages and disadvantages research it is clear that the low gas emission and the noise

reduced engine designs have the most benefits. Unfortunately, the noise reduced engine design is

insufficient to allow an aircraft to take off from the runway in Quito (at 9000 feet) in hot day condi-

tions. This is the greatest loss to this design. This design will further complicate the installation on the

Fokker 100, because it uses a high bypass ratio in comparison to the intermediate bypass in the cur-

rent Fokker 100. Because of these disadvantages, the low gas emission design will better meet the

demands of the client. For this reason, the low gas emission engine will be chosen by Engine Incorpo-

rated. The low gas emission engine gives benefits to its high overall efficiency, range, TSFC and en-

durance in all flight phases. It has a gas emission rating that proves that the gas emission are mini-

mal, seeing that the design is incorporated with a double annular combustion chamber that also uses

pre mixing of the air and fuel. This design is relatively maintainable, because of its simple design and

reliable parts.

3 Engine design With the selection of the new engine model, the low emission reduced engine design (HVA CS-25E),

the final configurations can be made. This will give the client an overview on the build up of the en-

gine. Before the design will be drawn up, the engine specifications should be reviewed (3.1). All

stages, systems and allocations on the engine will be defined. Once this is done the certification

process will be dealt (3.2). The process that will be taken on certifying the engine or, if proven neces-

sary, the aircraft will be explained. Furthermore, a maintenance scheme will be made for the engine

(3.3). This scheme will focus on the parts of the engine that will need maintenance within a specific

timeframe. The following paragraph will contain the financial overview of the engine (3.4). Here, the

costs that will be incorporated for the development and maintenance of the engine, will be dis-

cussed. A conclusion will then be drawn from the entire project (3.5). Finally, a recommendation of

the engine or a will be given (3.6).

3.1 HVA CS-25E engine The HVA CS-25E engine will now be further configured with specific allocations. This will be done by

defining the engine specifications (3.1.1), the sub systems (3.1.2) and a drawn layout (3.1.3).

3.1.1 The engine specified The HVA CS-25E engine is a design with many benefits. The following stages in the design will be

specified:

1. Inlet specification

2. Fan specification

3. Bypass specification

4. Compressor specification

5. Combustor specification

6. Turbine specification

7. Exhaust specification

ad 1 Inlet specification

The HVA CS-25E design will use a laminar inlet. This will minimize turbulence in the airflow. The inlet

will have a divergent shape that will slow down the air flow as it enters the engine. The diameter of

the inlet will be equal to the fan diameter.

ad 2 Fan specification

The fan will use curved blades that will give a total diameter of 1.176 m. This fan design will have a

fan pressure ratio of 1.6 and will being made of a honeycomb core which is sandwiched between two

titanium plate skins. The fan will have a total weight of 174.41 Kg and will be seen as the first, single

compressor stage.



ad 3 Bypass specification

This design has a high bypass ratio of 3.04. This will allow 24.75 percent of the air from the fan to

enter the core of the engine and 75.25 percent that will pass through the bypass. The air from the

bypass will help cool the engine and reduce noise.

ad 4 Compressor specification

The engine will have two compressor stages, a twin spool compressor type. The low pressure stage

will consist out of five compression stages, which will gradually compress the air. Then the high pres-

sure stage will consist of 14 compression stages, which will further increase the pressure. The com-

pressors will ensure a compression ratio of 2.22 and a dynamic ratio of 4.84. The compressors will be

made of a titanium, aluminium vanadium alloy (Ti6Al4V).

ad 5 Combustor specification

The double annular combustion chamber will be used and will deliver most of the engines weight.

The combustion chamber will be made of a cobalt, nickel, chromium and tungsten alloy (HA-188).

ad 6 Turbine specification

The engine will use impulse reaction turbines. It will consist of a two stage high pressure turbine. This

will power the high pressure compressor. The shaft that will be used for rotation will be called the N2

shaft. Then a three stage low pressure turbine will be added. This will power the low pressure com-

pressor and the fan, this will be the N1 shaft. The turbine blades will be made of nickel, tungsten and

rhenium alloys. The turbines will be cooled using turbine case cooling.

ad 7 Exhaust specification

This design will use a notched nozzle, this ensures rapid mixture of the core and bypass airflow, also

lowering the noise nuisance. The total diameter of the exhaust will be 0,32 meters. The exhaust will

be made of a nickel, tungsten alloy.

3.1.2 Sub systems specified

Apart from the main stages of the engine, the sub systems carry out an important role. The ignition

process will be in like a manner as to the original Fokker 100 engine design. The other subsystems

help monitor and maintain operation of the engine itself and other systems throughout the entire

aircraft. The engine will be monitored by the FADEC system. There are four main sub systems that

will be accommodated in the engine:

1. Accessory gearbox

2. Fire protection system

3. Fuel management system

4. Bleed air system

ad 1 Accessory gearbox

Attached to the N2 shaft an accessory gear box will be added. This gear box will power the engine oil

pump, the engine driven hydraulic pump, the engine driven lubrication pump, the low pressure fuel

pump and the high pressure fuel pump which build up the pressure for the FMS.

ad 2 Fire protection system

In case there is a fire, a fire protection system is installed. The system used is a continuous fire detec-

tion system. This uses a tube of eutectic salt with inside a thin wire. If the temperature of the engine

rises the resistance of the wire decreases and if the voltage rises to a certain level an alarm switch

will lit up. There are two warnings which indicate two stages. The first stage is the overheat warning

and the second is the fire warning.

ad 3 Fuel management system

The fuel management system that will be used will be similar to that of the Fokker 100. This system

will consist of an engine fuel pump, a fuel/oil heat exchanger, a fuel filter, a hydro mechanical unit, a



fuel shut off valve, a fuel flow transmitter and the fuel nozzles. This system will be the driving head of

the fuel that is inserted into the combustion chamber.

ad 4 Bleed air system

The HVA CS-25E engine will have an bleed air system similar to the current Fokker 100. The bleed air

system will not only be used for the pneumatic system of the aircraft, but will also be used for the

engine anti-icing, internal cooling, core compartment cooling and the turbine case cooling.

3.1.2 HVA CS-25E layout The HVA CS-25E will be constructed onto the Fokker 100 (appendix XLVII). Although the engine de-

sign uses a high by-pass, since the fan blade has a relatively similar diameter to that of the current

Fokker 100, the mounting of the engine will not prove problematic. The engine will have a diameter

of 1,176 meters and a length of 2,4 meters.

3.2 Certification When designing an engine, the engine has to be certificated by aviation authorities. The HVA CS-25E

engine has to be type certified (3.2.1). In some cases the design only has to be supplemental type

certified. This is a type of certificate that will be given when a modification is done without extensive

changes. The engine HVA CS-25E does incorporate extensive changes, so this engine has to be type

certified. When the engine is certified, also the aircraft has to be certified again to ensure safety. This

type of certificate is called an airworthiness certificate (3.2.2).

3.2.1 Type certificate

If an engine is designed in such a way that extensive changes are done concerning the design and the

performance, a type certificate has to be applied. In Europe the type certificate will be issued by the

EASA. When applying for a type certificate a three-view drawing has to be delivered together with

some design specifications and operating specifications. These operating specifications will contain

the performing limits of the engine. Also the regulations applied during designing and the Type Cer-

tificate Data Sheet (TCDS) has to be included. In this type certificate datasheet the holder of the cer-

tificate and operation specifications like ratings and limitations are included. This sheet will be made

by the applicant in cooperation with an EASA engineer.

The application for a type certificate is valid for five years. Within that time the designer can prove

that longer time is needed to finish the design. Then the authorities can allow the designer to take

more time for designing, developing or testing. During the application the regulations of that mo-

ment will be valid for the design period. This prevents that the designer has to change the design

because of changing regulations.

When the design is completed on paper, a prototype will be made. Mostly, a few prototypes are

made so that every prototype can be tested in a different way. One of these tests will be for example

a construction test where the design will be put under several stresses. Also system tests will be

done. After ground testing the engine will be subjected at flight test. In this stage of certifying the

applicant is asked to set up a maintenance program. When passing all these tests correctly the appli-

cant becomes the type certificate holder. The only condition is that the engine coming out of the

factory has to be exactly the same as the prototype.

3.2.2 Airworthiness certificate When an engine is type certified the aircraft will be airworthiness certified. This certificate is only

issued when the aircraft is a type certified design and thus is equal with the TCDS and operation spe-

cifications. This certificate gives the aircraft flight authority as long as the aircraft meets with the

regulations for safe operations and maintenance. When the aircraft is not conforming to these re-

quirements anymore the aircraft immediately lose its flight authority. In case the maintenance is not

done in accordance with the maintenance regulations of the relevant country, the flight authority

will be repealed until the overdue maintenance is done and registered.



This airworthiness certificate has to be stored in the aircraft and the operator has to be able to show

the certificate anytime when required by the authorities.

3.3 Maintenance research The engine of the Fokker 100 needs maintenance to ensure safe and smooth operation. The mainte-

nance and checks which need to be performed on the engine depend on the normal wear of the en-

gine and the unexpected wear which can occur to the engine. The damage which can occur to the

different stages of the engines is examined (3.3.1). During maintenance there are several methods of

engine inspection (3.3.2).

3.3.1 High and low maintenance parts In this part the different types of damages which can occur on the HVA CS25E engine are discussed.

The types of damages will be described according to the different stages of the engine:

1. Air intake

2. Compressor

3. Combustion chamber

4. Turbine

5. Exhaust

ad 1 Air intake

The air intake of the engine is very sensitive to damages. The damages to the inlet can be caused by

Foreign Object Damage (FOD), this means for example objects on the runway which can be sucked

into the engine and cause damage to the intake. Also if the aircraft has flown into a hail shower nicks

and dents in the fairing of the inlet occur. The inlet can also be damaged by bird strike and ice. The

intake of the engine is a part which needs regular maintenance.

ad 2 Compressor

The compressor of the engine is also vulnerable to different damages. For instance the compressor

blades can be exposed to FOD. This causes missing pieces in the tip area and the tip corners of the

leading edge and trailing edge of the blades. A maximum of 20 damaged blades is allowed in the axial

or radial length and must be less than 6,5mm. Other damages are worn areas or local distortion at

the blade tip corner on the leading and trailing edges, but also nicks, dents and scratches in the

area near the blade root and nicks. But also dents and scratches on the airfoil surfaces that are

not near the blade foot and not near the leading and trailing edges, furthermore distortion to

the leading and trailing edges and not in the blade foot area. Deposits and corrosion can also oc-

cur to the compressor blades. Regular inspection of the blades is necessary to make sure a blade

does not break off and cause further damage to the engine or the blade is in such a way that it pro-

vides a disturbed airflow.

ad 3 Combustion chamber

The combustion chamber can also exhibit damages. For instance the dome of the combustion cham-

ber can exhibit cracks; these cracks are allowed to about 51mm in length before repairing is neces-

sary. Around the plate the cracks are allowed to be 80mm. Missing of material is also allowed to

certain levels. The deflector of the combustion chamber can also be distorted.

ad 4 Turbine

The turbine of the engine can also be exposed to damage. For instance burns and perforations in the

high pressure turbine blades and shrouds, furthermore cracks in the leading and trailing edges of the

high pressure turbine blades as well as nicks, marks, scratches or dents in the concave or convex sur-

faces and the leading or trailing edges, as well as burns and erosion on the shrouds. The turbine area

is a part of the engine which needs regular inspections to be sure of well turbine operation.



ad 5 Exhaust

The damage which can occur to the exhaust system is overheating. Overheating can occur due to a

hot start. This means that during engine start-up, it is possible for temperatures inside the combus-

tion section to get hot enough to cause internal damage to the engine and the exhaust. The pilot has

to monitor the temperature of the engine during start-up to ensure the engine does not exceed the

limit. If the engine does exceed the limit a thorough check up of the engine may be necessary. The

engine needs to be fully removed from service and tear-down to provide an internal inspection to

determine if there was damage and repairing can be made.

3.3.2 Engine inspection During the life span of the engine, the engine needs to be inspected during maintenance. In the en-

gine several borescope ports are made to inspect in the compressor section the combustion chamber

(appendix XLVIII) and the turbine section (appendix XLIX). Another method of engine testing is the

Spectrometric Oil Analysis Program (SOAP). This method is used for the oil of the lubrication system

of the engine. This test measures the amount of metal particles which are present in the sample

which has been taken from the oil of the engine after the engine has run. The amount and type of

metal particles in the sample can give an indication of the wear of the engine and the necessary steps

which can be taken to prevent the engine from failing. Other engine testing methods are chip detec-

tor analysis, trend monitoring and vibration survey.

3.4 Financial aspects To give a supporting advice concerning the new gas turbine, the costs and the benefits of the next

generation engine must be investigated. The costs related to the design of a new gas turbine are the

design costs (3.4.1). Maintaining the engine airworthy will cost the airliner on ground time (3.4.2).

The next generation gas turbine will provide sufficient power with less fuel consumption to replace

the old gas turbines of the Fokker 100 (3.4.3). The costs and the benefits of the next generation gas

turbine will be compared in a breakeven point diagram (3.4.4).

3.4.1 Design costs When designing a new gas turbine, different cost must be taken into count. The costs for designing

this engine are onetime costs, meaning these costs are made only once in the design stage. Building

and testing the limits of the next generation engine will costs approximately € 25.000.000. With the

next generation engine design and operating specifications will the engine be certified according to

CS-25 and the CS-E of EASA. These costs will also be considered as design costs.

3.4.2 Maintenance costs Maintaining the engine airworthy will cost the maintenance crew man hours to perform the mainte-

nance program on the engine. The maintenance crew must first be trained to perform maintenance

on the next generation engine. Training the maintenance crew will cost the airliner approximately

€100.000. The maintenance programs on the engine are related with the amount of hours the engine

has performed or in a certain amount of days or if the pilot has any concern with the engine. The

average maintenance programs cost per year will be €400.000.

3.4.3 Benefits The new design of the engine will provide the airliner with less cost on the terms of emissions and

fuel consumption. The next generation engine innovative design will save up to €10.000 of the emis-

sions costs. The fuel consumption of the next generation engine compared with the old Rolls Roys

620-15 engine will provide 0.009 Kg/N*hours. The cost of fuel is € 0,55, meaning that the new engine

will save up to € 0,00495 for the amount of kilogram used per Newton hour. The average thrust of

the new engine is 48095,52 Newton. Multiplying the thrust with € 0,00495 will result in € 238,07 per

hour. Considering the engine is performing 300 days cycle, will result in a cost saving of €

1.714.124,214. The new engine will save up to €1.724.124,214 per year.



3.4.4 Breakeven point The next generation engine is a big investment for an airliner, however the breakeven point is not

reached until nineteen years. As shown in this breakeven point diagram.

3.5 Final conclusion The HVA CS-25E engine is designed to be an emission reduced engine with an improved thrust of

76696.02 N at MSL and a lower fuel consumption of 0.06001 Kgfuel/N*u at MSL, compared with the

old 620-15 engine (62820 N Thrust and 0.069 Kgfuel/N*u at MSL). The new engine has 13876.02 N

more thrust in comparison to the Tay 620-15, meaning that the engine will reach the 2.4 percent

climb radiant during a climb. The TSFC of the HVA CS-25E compared with the old 620-15 engine will

be around 0.00889 Kgfuel/N*u at MSL. Lower fuel consumption means that the new engine will pro-

duce lower noise levels, lower emissions, a greater range and will save the airliner up to €

1.714.124,214 a year.

The maintainability of the engine has been improved. This improvement has been done by choosing

super alloys for the high maintenance components, saving the airliner an amount of €400.000 a year

and time with the maintenance of the engine. When the costs of building and testing of the new

engine are compared with the benefits of the new engine, this will result in a break-even point that

will be reached in about nineteen years.

The new engine has a greater reliability compared to the new engine, due to its innovated design.

The new engine has an integrated FADEC system, this systems improves the fuel flow and the reliabil-

ity of the engine. The laminar air intake will improve the stall en surge characteristic of the engine,

making the engine more reliable.

3.6 Recommendation The HVA CS-25E engine is the next generation engine for all Fokker 100 aircraft, that is capable of

producing 53204.31 N thrust on 9000ft altitude with an average range of approximately 1767.47 Nm.

The new engine complies to the clients demand, these demands were that the engine must have a

minimum thrust (51065,5 N) at a elevation of 9000ft and a minimum range of 1500nm.

Despite all these good qualities of the engine, the break-even point of the engine shall not be

reached until nineteen years in business. Therefore, the recommendation of Engine incorporated B.V

for the airliner, is not purchasing the engine for the Fokker 100 aircraft of nowadays. This is because

of the break-even point not being reached until nineteen years. Several high maintenance compo-

nents in the turbine have a lifespan of twenty years. This means that the airliner after reaching the

break-even point will spend a great amount of money on replacing these components in the engine.

For the next generation Fokker 100 this engine is recommended. This is because taking off at an ele-

vation of 9000 ft with a maximum pay load is possible with this engine.

€ 0

€ 5.000.000

€ 10.000.000

€ 15.000.000

€ 20.000.000

€ 25.000.000

€ 30.000.000

€ 35.000.000

€ 40.000.000

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22

Costs

Benefits



Reference list

Books: Asselman, L.

Gasturbines algemeen

Bleskensgraaf, 2008.

Boyce, Meherwan P.

Gas Turbine Engineering Handbook Second Edition

Oxford, 2002

Cohen, H., G.F.C. Rogers, H.I.H. Saravanamutoo.

Gas Turbine Technology

4th edition

Harlow, 1996

Dijk, B.H. van

Dictaat Stromingsleer

Amsterdam, 2009

Hogeschool van Amsterdam

Amsterdamse Hogeschool voor Techniek

EASA

CS-25 Certifications for large aeroplanes

Koeln, 2004

EASA

CS-E

Koeln, 2004

EASA

CS-34

Koeln, 2004

EU

European Commission, Aviation and climate change On: http://ec.europa.eu/ 18-09-2009

http://ec.europa.eu/environment/climat/aviation/index_en.htm

03-11-2009

Frenchez, Pietersz

Projectboek periode 6 Powerplant

Amsterdam, 2009.

Hogeschool van Amsterdam

Domein Techniek

General electric

Aircraft gas turbine guide

Ohio, 1997

Horlock, J.H

Advanced gas turbine cycles

2003 Cambridge



Hüncke, K.

Jet engines

6th edition

Shrewsbury, 2003

Houtman, C.J.

Vliegtuiggasturbines

Technische Universiteit Delft

Delft, 1987

KLM

Aircraft Operations Manual Fokker 100

Amstelveen, 1988

KLM

Aircraft Operations Manual Fokker 100

Amstelveen, 1989

Kimmenade, Ir. A.J.M.van

Warmteleer voor Technici

6de druk

Haarlem 1976

Jeppesen

Aircraft general knowledge 3

Oxford 2002

Kroes, Michael J. & Thomas W. Wild

Aircraft Powerplants

7th edition

New York, 1995

Mattingly, J.

Elements of Gas Turbine Propulsion

2nd edition

Singapore 2006

Mattingly, J.

Aircraft engine design

2nd edition

Virginia 2002

Pol, Prof Dr. Ir. E. v.d

Gasturbines

Technische Universiteit Delft

Delft, 1985

Rolls Royce

The Jet engine

5th edition

Derby, 1996



Ruijgrok, G.J.J.

Elements of Airplane Performance

4de druk

Delft 2004

Scholder, Robert H.G.

Vliegtuiggasturbines

Amsterdam, 2005

Hogeschool van Amsterdam, Domein Techniek

Schornagel, C.

Motoren

Eelde, 1991

Siers F.J.

Methodisch ontwerpen: volgens H.H. van den Kroonenberg

3e druk

Groningen, 2004

Torenbeek, E.

Flight Phisics

Delft, 2009

Treager, Irwin E.

Aircraft Gas Turbine Engine Technology

3rd Edition Westerville

Ohio, 1996

Wentzel, Tilly

Het projectgroepsverslag

Amsterdam, 2008

Hogeschool van Amsterdam, Domein Techniek

Wild, Thomas W.

Transport Category Aircaft systems

Jeppeson

Englewood Col. 1996

Internet sources:

EASA

EASA, Agency Measures - Certification specifications On: http://www.easa.eu.int 03-11-2009

http://www.easa.eu.int/ws_prod/g/rg_certspecs.php

03-11-2009

EASA

EASA, Certification - Product Certification On: http://www.easa.eu.int 03-11-2009

http://www.easa.eu.int/ws_prod/c/c_da_main.php

03-11-2009

FAA

Type certificate data sheet on: http://www.airweb.faa.gov 2009

http://www.airweb.faa.gov/Regulatory_and_Guidance_Library/rgMakeModel.nsf/MainFrame?Open

FrameSet



02-11-2009

International Civil Aviation Organization

Publications On: www.icao.int 03-11-2009

http://www.icao.int/icao/en/m_publications.html

03-11-2009

NASA

The Evolution of Modern Aircraft on: http://www.hq.nasa.gov 17-08-2004

http://www.hq.nasa.gov/office/pao/History/SP-468/cover.htm

03-11-2009

Rekkof

Fokker Aircraft next generation

On: http://www.rekkof.nl/ 11-09-2008

http://www.rekkof.nl/index.html

03-11-2009

Rolls Royce

Civil Aerospace on: http://www.rolls-royce.com 02-11-2009

http://www.rolls-royce.com/civil/index.jsp

03-11-2009

Stork Fokker, Fokker Services

Fokker 100 On: http://www.fokker.com/ 21-10-2009

http://www.fokker.com/Stork/2077/Fokker_100.htmll

03-11-2009

project report power plant - portfolio anouk elsendoorn ... report power plant group 2a2o amsterdam,...

Documents