internship report david jimenez

7/29/2019 Internship Report David Jimenez

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Stage Master 2 GSATSAS

du 15 Mars au 30 Aout

2013

LMS, A Siemens Business

Dveloppement de modles dynamiques de PropulseursAronautiques

AS&D1-D Division

84, quai Charles de Gaulle

69006 Lyon (France)

Nom EtudiantM. David Jimnez Mena

Master 2 GSAT IMA2012-2013

Nom tuteur en entrepriseDr/Ing Loig ALLAIN

Titre / fonction : Product LineManager System SimulationAS&D

Nom tuteur IMADr. Anissa Meziane


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M2 David JIMENEZ MENA Dveloppement de modles de Propulseurs Aronautiques Page 2

SUMMARY

1. ACKNOWLEDGEMENT ............................................................................................................. 5

2. INTRODUCTION ......................................................................................................................... 6

2.1. Aviation Growth .................................................................................................................................. 6

2.2. First and main energy source: Gas Turbine ......................................................................................... 7

2.2.1. State of the Art Subsonic Engine SFC ................................................................................................ 8

2.3. Gas Turbine System Developer - Internship ......................................................................................... 8

3. CONTEXT ...................................................................................................................................... 9

3.1. Enterprise LMS Imagine .................................................................................................................... 9

3.2. Siemens Acquisition............................................................................................................................. 9

3.3. Organization Chart............................................................................................................................. 10

3.4. AS&D Team ....................................................................................................................................... 10

3.5. AMESIM ............................................................................................................................................. 10

3.5.1. Global Presentation ......................................................... ................................................................ 11

3.5.2. The Bond Graph Theory .................................................................................................................. 11

3.5.2.1. Causality ...................................................................................................................................... 12

3.5.3. Mass Spring system example ....................................................................................................... 14

4. JET ENGINE FUNDAMENTALS ............................................................................................ 16

4.1. Gas Turbine Engine ............................................................................................................................ 16

4.1.1. Thermodynamic Cycle Brayton Cycle ........................................................................................... 16

4.1.2. Jet Engine ........................................................................................................................................ 17

4.1.2.1. Jet Propulsion systems .................................................................................................................... 17

4.1.2.2. Air-breathing Engines ................................................................................................................. 18

4.1.2.3. Engine performances. Thermal and propulsive performance .................................................... 224.1.2.4. Technical limitations ................................................................................................................... 24

5. JET ENGINE MODELING ........................................................................................................ 27

5.1. Model already done - Analysis ........................................................................................................... 27

5.1.1. Main Component description ......................................................................................................... 27

5.1.1.1. Boundary Condition: Temperature/Pressure source ........................................................ .......... 27

5.1.1.2. Inlet and nozzle components ...................................................................................................... 29

5.1.1.3. Compressor and turbine components ........................................................................................ 29

5.1.1.4. Combustion chamber component .............................................................................................. 305.1.2. Hypothesis ....................................................................................................................................... 31

5.1.2.1. Hypotheses used Gas Mixture Library ..................................................................................... 31


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5.1.2.2. Hypothesis checks ....................................................................................................................... 31

5.1.3. Performance Analysis ...................................................................................................................... 34

5.1.4. Component Analysis ........................................................ ................................................................ 36

5.1.4.1. Inlet ............................................................................................................................................. 36

5.1.4.2. Compressor ................................................................................................................................. 39

5.1.4.3. Turbine ........................................................................................................................................ 435.1.4.4. Other critic points ....................................................................................................................... 45

6. A JET ENGINE LIBRARY ........................................................................................................ 45

6.1. Created Components ......................................................................................................................... 45

6.1.1. Inlet ................................................................................................................................................. 45

6.1.1.1. Element Test - Results ............................................................ ..................................................... 46

6.1.1.2. Modeler method ......................................................... ................................................................ 47

6.1.2. Compressor ..................................................................................................................................... 47

6.1.2.1. Mass flow rate problem .............................................................................................................. 47

6.1.2.2. IFP Library compressor ............................................................................................................... 48

6.1.2.3. Performance Compressor Map Data .......................................................................................... 49

6.1.2.4. Inertial Resistive component ...................................................................................................... 51

6.1.2.5. Calculation steps ......................................................................................................................... 53


6.1.2.7. Modeler method ......................................................... ................................................................ 57

6.1.3. Turbine ............................................................................................................................................ 57

6.1.3.1. Performance Turbine Map Data ................................................................................................. 59

6.1.3.2. Resistive component Calculation steps ................................................................................... 59


6.1.3.4. Modeler method ......................................................... ................................................................ 636.1.4. Air Split ............................................................................................................................................ 64


6.1.4.2. Modeler method ......................................................... ................................................................ 66

6.1.5. Nozzle Double Corp ...................................................................................................................... 67

6.1.5.1. Element Test Results ................................................................................................................ 67

6.1.5.2. Modeler method ......................................................... ................................................................ 69

6.1.6. Propeller .......................................................................................................................................... 69

6.1.6.1. First level of modeling................................................................................................................. 69

6.1.6.2. Second level of modeling ................................................................. ........................................... 71

6.1.6.3. Modeler method ......................................................... ................................................................ 72

7. SCHEDULE ................................................................................................................................. 73

7.1. V Cycle ............................................................................................................................................... 74

8. RISK ANALYZE ......................................................................................................................... 75

9. CONCLUSION ............................................................................................................................ 76

9.1. Project Synthesis ............................................................................................................................... 76

9.2. Personal Synthesis ............................................................................................................................. 76


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10. BIBLIOGRAPHY ................................................................................................................... 78

11. ANNEXES ............................................................................................................................... 79

11.1. LMS History ....................................................................................................................................... 79

11.1.1. Before 2007 ..................................................................................................................................... 79

11.2. Causality Rules................................................................................................................................... 80

11.2.1. I Element ......................................................................................................................................... 80

11.2.1.1. Hydraulic ..................................................................................................................................... 80

11.2.1.2. Mechanic .................................................................................................................................... 80

11.2.1.3. Electric ........................................................................................................................................ 80

11.2.2. C Element ........................................................................................................................................ 80

11.2.2.1. Hydraulic ..................................................................................................................................... 80

11.2.2.2. Mechanic .................................................................................................................................... 81

11.2.2.3. Electric ........................................................................................................................................ 81

11.2.3. R Element ........................................................................................................................................ 81

11.2.3.1. Hydraulic ..................................................................................................................................... 81

11.2.3.2. Mechanic .................................................................................................................................... 82

11.2.3.3. Electric ........................................................................................................................................ 82

11.3. Fundamental Equations ..................................................................................................................... 83

11.3.1. Conservation of matter ................................................................................................................... 83

11.3.2. Conservation of energy ................................................................................................................... 83

11.3.3. Jet Engine Thrust ............................................................................................................................. 84

11.4. Inlet/ Nozzle modeling: ..................................................................................................................... 84

11.5. Compressor/ Turbine modeling ......................................................................................................... 85

11.6. Axial compressor inertial equations................................................................................................... 86

12. VARIABLES GLOSSARY ..................................................................................................... 89

13. SYMBOLS GLOSSARY ......................................................................................................... 90

14. INDEX DES FIGURES .......................................................................................................... 90


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1. ACKNOWLEDGEMENTIt is very difficult to start an acknowledgement knowing that every person, I was working with, could expend

almost all its day long explaining me whatever, aiding me in whatever I had as problem. But anyway, I am going

to try it

First of all, I would like to express my biggest gratitude to my mentor Dr. Loig ALLAIN: thanks, at first,

for the opportunity and, above all, the trust you gave me at the starting and during this internship.

Thanks for all the challenges you proposed me, and for the encouragement you gave me every time I

need it, I really appreciate them.

Secondly, I would like to thank another person who has been supporting me as well during all this

period and who has solved a lot of my several questions about everything I had, Dr. Olivier BROCA.His help has been really important to me.

As well, I wish to express my appreciation to Mr. Djiby Toure, he is the person who has such good

knowledge of programming and AMESim field, that I could not find any question that he could not

answer. He spent, as everyone here, a good part of his time helping me.

My grateful thanks could not go anywhere else but toward Mr. Louis de Riberolles, thanks for the big

foreign conversations which had made me feel closer to my natal country. Thanks for being the

best desk neighbor.

Finally, thanks to Grgoire Grenoble and Stphane Mouvand for all such interesting conversationsand the good atmosphere that they create with their good humor.

In conclusion, I would like to thank all the LMS Imagine staff for making it a pleasant work place.


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2. IntroductionThe evolution announced by the aviation transport has been bringing environment concerns about

pollutant emission reduction and noises reduction. The aeronautic transports are focusing onreduction of the pollutant emissions and the noises of each flight phase amongst other points.

In this line, the two giants of the aircraft manufacturer Airbus and Boeing afford in order to take a

part of the market of the new less pollutant line aircraft: Airbus 350 XWB and Boeing 787 Dreamliner.

Figure1 : Boeing 787 Dreamliner / A350 XWB

2.1.Aviation Growth

Aviation is a critical aspect of modern society, moving people and goods throughout the world and

fostering economic growth. From 1980 to 2010, the demand for air transportation grew by a factorof four; while forecasts for the next 25 years vary, they present a strong message that this trend will

continue.

Figure 2 : Billions of passengers transported by air transportation

Growth in the total volume of air transportation has important environmental ramifications

associated with climate change and stratospheric ozone reduction on a global scale. On local to

regional scales, issues such as noise, decreased air quality and local water quality are recognized as

important consequences of air transportation.


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Aviation fuel burn is responsible for approximately two to three percent of global carbon dioxide

(CO2) emissions, and aviation it is considered to be a fastest growing, potentially significant source of

greenhouse gas emissions. Commercial aviation is increasingly being targeted by legislators for

mandatory carbon-trading schemes and limits on aircraft emissions.

In this framework, the aeronautic transports are fixing the objective of reduction the pollutantemissions.

2.2.First and main energy source: Gas Turbine

From the aircraft creation, the aeronautic transport has been dependent of its main source: the

engine. At this last period aircraft manufacturers are very interested in motorist manufacturer

evolution and one of causes is this pollution and high consumption problem.

Consequently, in the reduction of the pollutant emissions, aeronautic transports are very dependent

on the evolution of gas turbines engines.

The last example is the A350XWB aircraft. This aircraft delivers 25 per cent lower fuel burn per seat

when compared to the current competing jetliner. Of this performance, about one-third is due to its

Rolls Royces power-plant: Trent-XWB.

Figure 3 : Rolls Royce engine: Trent-XWB


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2.2.1. State of the Art Subsonic Engine SFC

The parameter which describes and let compare the different engines is the thrust-specific fuel

consumption (TSFC). This parameter has got a 50% reduction during the past 50 years. The reduction

in fuel burned can be observed by engine type categorized into turbojets followed by low, medium,

and high bypass ratio turbofans.

Figure 4 : State of the Art Subsonic Engine SFC - NASA

2.3.Gas Turbine System Developer - Internship

As you could realize, the reduction of the consumption in the aeronautic transport is a high interest

for aircraft constructors. Consequently, it is important for aircraft, equipment and system

manufacturer to work on the optimization of the energy consumption. Knowing when, why and howthe energy is consumed will allow for the 50% CO2 emission reduction objective.

The internship subject is based on the development of different tools for motorist developer system

engineers in order to help them at the engine system modeler during the gas turbine design phases

until the transient integration.

Focusing in the different component they are in the gas turbines. The behavior of each one

(compressor, turbine, combustion chamber) shall be studied to account the component on its own,

and also, studying his behavior in a global gas turbine system.

AMESim software and Python, C/C++ language were used.


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3. Context3.1.Enterprise LMS Imagine

LMS International is an engineering innovation partner for companies in the automotive, aerospaceand other advanced manufacturing industries. LMS enables its customers to get better products

faster to market, and to turn superior process efficiency to their strategic competitive advantage.

LMS offers a unique combination of virtual simulation software, testing systems and engineering

services.

LMS is focused on the mission critical performance attributes in key manufacturing industries,

including structural integrity, system dynamics, handling, safety, reliability, comfort and sound

quality. Through its technology, people and over 25 years of experience, LMS has become the

partner of choice for most of the leading discrete manufacturing companies worldwide. LMS is

certified to ISO9001: 2000 quality standards and operates through a network of subsidiaries andrepresentatives in key locations around the world.

The acquisition of Imagine enabled LMS to develop strategically its current wallet of solutions of

simulation of functional performances and physical tests, and thus to provide a complete range of

applications of modelling, simulation and test of the effective behavior of mechanical and

mechatronic intelligent systems.

Today, LMS Imagine grows up around two poles of complementary activities: software development

(LMS Imagine.Lab suite) and engineering services based on this platform as automobile and

aeronautic industries.

3.2.Siemens Acquisition

As of January 3, 2013, LMS was acquired by Siemens. As a business segment within Siemens PLM

Software, LMS will provide a portfolio of products and services for manufacturing companies to

manage the complexities of tomorrows product development by incorporating model -based

mechatronic simulation and advanced test in the product development process.


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3.3.Organization Chart

The organization chart of Siemens LMS Imagine is as follows:

Figure 5 : LMS organization

3.4.AS&D Team

Aircraft Space and Defense Team is composed by 5 engineers or doctors who develop aerospace

solutions. Its know-hows touch hydraulic, electronica, pneumatic, thermal, dynamic and

aerodynamic fields.

My internship has been developed in this department; their expertise lets extend my knowledge in

its different fields.

3.5.AMESIM

AMESim is a simulation software for modeling and analyzing multi-domain systems. It is part of

systems engineering domain and falls into the mechatronic engineering field.

AMESim is the acronym of Advanced Modelling Environnement for performing Simulations ofengineering systems. LMS Imagine develops a simulation platform in order to create 1D virtual

models (also known as System Model). It proposes a complete multi-domain approach for an

integration in the same environment. It allows you to analyze and to optimize the performance of

the complexes units, like for example an aircraft or a car.


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3.5.1. Global Presentation

A system is modeled in AMESim by a set of connected components.

These components represent several validated analytic sub-models, which could be represented in

an one-dimensional geometry, that represents either the hydraulic, pneumatic, electric or

mechanical physic behavior of the system.

The components of the same physic domain are gathered libraries allowing for a package.

This approach lets the possibility to simulate the behavior of intelligent systems independently of the

detailed CAO geometry.

The link between each sub-model is done by the component ports which are conditioned by the

causality concept inherited from the Bond Graph. A link is allowed if it respects the causality

between components.

3.5.2. The Bond Graph TheoryAMESim is based on the Bond Graph theory: It is a graphical representation of a physical dynamic

system. It represents exchange of physical energy, allowing the utilization of different physic

domains.

The fundamental idea of a bond graph is that power is transmitted between connected components

by a combination of effort and flow.

Systems Effort (e) Flow (f)

MechanicalForce (F) Velocity (v)Torque () Angular velocity ()

Electrical Voltage (V) Current (i)Hydraulic Pressure (P) Volume flow rate (dQ/dt)

Thermal Temperature (T) Entropy change rate (ds/dt)Pressure (P) Volume change rate (dV/dt)

ChemicalChemical potential () Mole flow rate (dN/dt)Enthalpy (h) Mass flow rate (dm/dt)

Magnetic Magneto-motive force (em

) Magnetic flux (

)


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Pneumatic

Temperature (T) Mass flow rate (dm/dt)

Pressure (P) Enthalpy flow rate (dm h/dt)

Each port of the AMESim has at least the 2 variables (effort and flow) representing its energy

domain.

The velocity (flow) and the force (effort) are transmitted by its ports in the follow examples:

Figure 6: A mechanical component

Figure 7: An electrical component

Figure 8: A pneumatic component

3.5.2.1. CausalityBond Graph has a notion of causality, indicating which side of a bond determines the instantaneous

effort and which determines the instantaneous flow.

As the word means, causality is the relation between an event (the cause) and a second event (the

effect), where the second event is understood as a consequence of the first.

Bond Graph uses the energy exchanges, and using the conservation of energy, in each component it

is calculated the determination either of flow or effort, depending if it has been given as input the

effort or flow.

Because of this causality, AMESim represent all the domain of physics using 3 main elements:

- Inertia element I: it is used to model inductance effects in electrical systems and mass or inertia

effects in mechanical or fluid systems.


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- Capacitive element C: is an element which uses a static constitutive relation exists between an

effort and a displacement. Such a device stores and gives up energy without loss. In bond graph

terminology, an element that relates effort to the generalized displacement (or time integral of flow)

is called a one port capacitor. In the physical terms, a capacitor is an idealization of devices like

springs, torsion bars, electrical capacitors, gravity tanks, accumulators, etc.

- Resistive element R: is an element in which the effort and flow variables at the single port are

related by a static function. Usually, resistors dissipate energy. This must be true for simple electrical

resistors, mechanical dampers or dashpots, porous plugs in fluid lines, and other analogous passive

elements.

It is exampled in the next figure C,I,R elements for different domain of physics (Causality rules

explained in Annexes):

Domain Inertia Capacitive Resistive

Hydraulic Hydraulic Inertia Volume Orifice

Mechanic Mass Stiffness Friction

Electric Inductance Capacitor Resistance


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3.5.3. Mass Spring system example

In order to model a mass spring system, we need to recover the needed components from the

mechanical library that we can see in the next figure.

Figure 9 : Different AMESim Component Figure 10 : Library Tree

In the figure 9 we can see different components of the mechanical library as well as the possibility of

see the component ports.

We can realize that the apple that represents the gravity does not have any port, because it defines

a system property. In the other way, we can realize that the components which have ports, they have

inputs (red arrows) and outputs (green arrows) of different units.

In order to respect the causality, it is necessary that every input has the same output corresponding.

The conexion between each component is done by the intermediary of its ports. You can realize that

the product of the inputs and outputs are equal to the Power. Indeed, the power let the cohesion of

all this physic domains and the Bond Graph guarantees the energy conservation.

Figure 11 : Example of a mass-spring sketch and the mass 4 temporal displacement


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The figure shows an example of mass-spring system and the visualization of the displacement of the

mass number 4.

Inside this sketch, AMESim has:

- Component 0 which imposes null displacement, speed and acceleration.- Component 1 which generates a force.

- Component 2 which generates a displacement, a speed and an acceleration.

- Component 3 which generates a force.

- Component 4 which generates a displacement, a speed and an acceleration.

- Component 5 which generates a null force.

- Component 6 which describes the gravity.

We can realize that the causality is respected in each link.

Looking at the last figure graph we can realize as well that the system is conservative. Indeed, they

are four big component categories:

- Inertial components: like masses.

- Capacitive: like springs.

- Dissipative: like dampers.

- Other: like the gravity.

The system does not interact with dissipative components, so it is a conservative system. If the

model is not corresponding to our needs, we can model a spring using a spring-dump for example.

Figure 12 : Example of a mass- dumper string sketch and the mass 4 temporal displacement

This time, we can see the dump effect: it absorbs the system energy, having by consequence adecreasing amplitude of displacement.


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4. Jet Engine FundamentalsIn order to understand how the aircraft gas turbine engine operates, one should be familiar with

some of the essential physics laws that govern the field of turbopropulsion.

4.1.Gas Turbine Engine

Gas Turbine is an industrial application of a specified thermodynamic cycle, the Brayton Cycle. This

chapter contains the description of this thermodynamic cycle, its aeronautic application (the jet

engine) and the explanation of the several kind of jet engine which are already developed.

4.1.1. Thermodynamic Cycle Brayton Cycle

Gas Turbine engine is a rotary thermodynamic machine which role is to deliver mechanical energy

(either increasing flow kinetic energy, or increasing mechanical shaft rotation) from calorific energy

generated at the hydrocarbon combustion.

This internal thermic combustion engine uses the Brayton cycle during its work life. This

thermodynamic cycle describes the evolution of the air flow state condition during this constant

pressure heat engine. This open system cycle makes different phases:

Flow is pressurized in an isentropic process (1-3). The pressurized flow follows an isobaric process; heat is added on (3-4). The flow is expanded in an isentropic process (4-6). Heat rejection is done following an isobaric process (6-1).

Figure 13 : Brayton Thermodynamic cycle

In a practical gas turbine, the main required elements are:

-Compressor elements

-Combustion chamber

-Turbine elements

These three elements compose all the gas generator systems.


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4.1.2. Jet Engine

In order to get this cycle in an aeronautic application, we will need to add specific components to get

a better performance of this cycle:

- Inlet: This element allows using the aircraft kinetic energy to compress the flow before enteringto compressor.

- Nozzle: Being the main goal the increasing of moment quantity in order to get the maximum ofthrust to propel the aircraft, this machine should increase the jet engine exit speed. For this

reason, the nozzle element is used in the jet engine cycle.

We can see the differences between a car engine solution and an aircraft engine solution.

Figure 14 : Comparation between Jet Engine system and car engine system

4.1.2.1. Jet Propulsion systems

Jet Propulsion principle is based in the momentum change of a fluid by the propulsion system as

already described. The main goal of these systems is to produce a thrust force in order to propel

aeronautic vehicles.

The fluid may be either the gas used by the engine itself (e.g., turbojet), a fluid available in the

surrounding environment (e.g., air used by a propeller) or it may be stored in the vehicle and carried

it during the flight (e.g., rocket).

Jet propulsion systems can be subdivided into two broad categories: air-breathing and non-air-

breathing engines. Air-breathing propulsion systems include the reciprocating, turbojet, ramjet,

turboprop and turboshaft engines. Non-air-breathing engines include rocket motors, nuclear

propulsion systems, and electric propulsion systems.

Focus of this work is turbine propulsion systems (turbojet, turbofan, turboprop and turboshaftengines).


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Figure 15 : Jet Engine family

4.1.2.2. Air-breathing Engines4.1.2.2.1. Gas Generator

Gas generator is defined as the heartof gas turbines. A schematic diagram of a gas generator is

shown in the next figure.

Figure 16 : Gas Generator schema

Compressor, combustion chamber, and turbine are the main components of gas generators, being

this subsystem a common structure for turbojet, turbofan, turboprop and turboshaft engines. The

purpose of gas generators is supply high-temperature and high-pressure gas.

Figure 17 : Different gas turbines types using the same gas generator schema


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Turbojet

Turbojets provide thrust using the jet propulsion principle: the air is compressed through inlet and

compressor components, after that air is mixed with fuel and it is burned in the combustor chamber

and finally, they are expanded through the turbine and nozzle components.

The gas expansion through the turbines supplies the power to turn the compressor. The net thrustdelivered by the engine is the result of converting internal energy to kinetic energy.

Figure 18 : Turbojet schema

Pressure, temperature and velocity variations through a jet engine system are shown in the next

figure. In the compressor section, the pressure and temperature increase, the gas absorbs the energy

provided by the compressor. The temperature of the gas is further increased by the combustion

done through the combustion chamber. In the turbine section, the energy contained in the hot gas is

absorbed and converted to shaft power which is provided by a mechanical system to the

compressor. This absorption is done by an expansion process which results in decreasing the

temperature and pressure of the flow. In the nozzle, the gas stream is further expanded to produce ahigh exit kinetic energy.

Figure 19: Typical single-spool axial flow turbo-jet engine

Turbojets work is very performance at high speed; however, as aircraft speed gets lower values, its

performance decreases (Causes explained in the chapter: Engine Performance).

In order to work at this low aircraft speed, gas turbine types have been developed: turbofan,turboprop and turboshaft engines.


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4.1.2.2.2. Turbofan

Turbofan system uses this configuration: inlet, fan, gas generator, and nozzle. In turbofan engines, a

portion of the turbine work is used to supply the fan power. This element propels the incoming air

which is dividing in two flows: the core flow and the secondary one. The main flow passes through

the gas generator and the secondary flow goes through the long fan duct to the fan exhaust so as to

increase the mass flow and this way the engine trust is higher.

Figure 20 : Flow schema for turbofan systems

Inside a limited realm of flight, the turbofan engine is more economical and efficiently than the

turbojet engine. The thrust specific fuel consumption (TSFC or fuel mass flow rate per unit thrust) is

lower for turbofans and indicates a more economical operation.

Figure 21 : Turbofan system

The principal advantage of these systems is that in a low flight speed, this system accelerates a larger

mass of air than a turbojet, having a higher propulsive efficiency (this concept is explained in the next

chapter).


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4.1.2.2.3. Turboprop and Turboshaft

The gas generators that drive a propeller are called turboprop engines. The expansion of gas through

the turbine supplies the flow energy required to power the propeller. The turboshaft engine is similar

to the turboprop except that power is supplied to a shaft rather than a propeller. The turboshaft

engine is used quite extensively for supplying power for helicopters. The limitations and advantages

of turboprops are those of the propeller. For low-speed flight and short-field takeoff, the propeller

has a performance advantage. When speed approaches the sound speed, compressibility effects

produce loses in the aerodynamic propeller efficiency. Due to the rotation of the propeller, the pales

tip will approach the sound speed before the vehicle approaches to that speed.

Figure 22: Turboshaft engine


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4.1.2.3. Engine performances. Thermal and propulsive performanceEach kind of engine operates only inside a certain range of altitudes and Mach numbers.

In order to understand the operational and the best performance range of engines, the overall

engine system has been analyzed.

Jet Propulsion system should be studied in two parts or subsystems:

-Thermal or engine System:

It is the subsystem which transforms fuel energy to kinetic and calorific energy. The parameter which

describes its performance is the thermal efficiency . It is defined as the net rate of organizedenergy (shaft power or kinetic energy) out of the engine divided by the rate of thermal energy

available from the fuel in the engine, .Thermal efficiency can be written in an equation form as

For engines with shaft power output as turboshafts, is equal to out shaft power. For engineswith no shaft power output as turbojet engines, is equal to the rate of change of the kineticenergy of the fluid through the engine. The output power of a jet engine with a single inlet and a

single exhaust is given by the difference of the kinetic energy between the exit and entrance flows.

(Fundamental theories jet engine: see annexes Fundamental Equations)

[( ) ], is the air mass flow; is the fuel mass flow; is the flight speed; is the exit flow speed.-Propulsion System:

It is the subsystem which transforms the shaft power or kinetic energy to propulsive power.

Figure 23: Propulsion subsystem


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For a jet engine the movement quantity varies expanding the combustion products. For a turbofan

the variation of movement quantity will be produced by the propeller propulsion.

The parameter which describes its performance is the propulsive efficiency . It measures howeffectively the engine power

is used to power the aircraft. Propulsive efficiency is the ratio of

the aircraft power (thrust multiplied by velocity) to the power out of the engine, , in equationform, this is written as

* +

This equation shows that the propulsion performance increases with respect to the flight speed. This

allows the analysis of the various architecture of gas turbine:

-Turbojet engines have a high speed ratio at low speed, giving a low propulsive performance at

these flight conditions. Increasing the flight speed, propulsive performance increases proportionally.

-Turbofans propulsive performance in a low flight speed range is better than turbojets. Its

performance increases with respect the flight speed. The disadvantage of the propeller propulsive

engines at the higher aircraft speeds is its rapid fall off in efficiency, due to shock waves created

around the propeller as the blade tip speed approaches Mach 1.0.

Figure 24 : Propulsive efficiency


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4.1.2.4. Technical limitationsThe development of a jet engine model should not forget the technical limitation inherent to their

components. The high temperature and pressure conditions of work entail that in some occasions

theoretical solutions cannot succeed.

Because of that, the evolution of gas turbines has different important key technologies to bedeveloped at the same time than jet engines.

4.1.2.4.1. Materials

The evolution of the gas turbine appropriately began with the achievements of the materials and

manufacturing process engineers. A chronological progress of turbine airfoil material capability over

the past 50 years shows an improvement exceeding 260 C (500 F).

Figure 25 : Metal Temperature Capability

4.1.2.4.2. Turbine Airfoil Cooling

In order to protect of the high temperature of the combustion chamber exit gas, turbine blades have

a system which creates a cool film air on the airfoil leading edge.

Figure 26 : Turbine solutions

The jet engine turbine blade is the most sophisticated heat exchanger. Therefore, turbines work at aninlet gas temperature higher than its material melt temperature.


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Figure 27 : Rotor inlet gas temperature Vs. Cooling effectiveness

4.1.2.4.3. Compressor Design and Engine ConfigurationThe compressor has often been referred to as the heart of the engines. The normal compressor

operating line and stall line (a compression limit) is shown as a function of pressure ratio and mass

flow rate.

Figure 28 : Performance Compressor Map

As shown in the figure, there are several factors that have the potential for causing a flow

breakdown. The compressor stall margin remaining (SMR) is a key for having a safe engine operation.


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The engine cycle efficiency has achieved higher values increasing compressor pressure ratios.

Pressure ratios for subsonic aircraft applications have increased by a factor of 20 during the past 50

years. The more recent engines have reached pressure ratios of 40 at sea level.

Figure 29 : Compressor Pressure ratio evolution


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5. Jet Engine Modeling5.1.Model already done - Analysis

It was a first level of turboshaft engine model, a gas turbine created to drive helicopter propellerswas already developed.

Figure 30: functional model of turboshaft engineThanks to the multi-physic capability of AMESim, mechanical and pneumatic models are connected

together in order to simulate this system. A description of the components from those systems is

done in this chapter.

5.1.1. Main Component description

5.1.1.1.

Boundary Condition: Temperature/Pressure sourceThis component is created in order to simulate any operational environment conditions, in this case,

for the gas turbine. Determining pressure and temperature air condition, the take-off, cruise or

landing phase can be simulated.


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Figure 31: source componentTo provide the accurate work conditions, this model uses the atmosphere component. Giving as

input the flight altitude and flight Mach to the atmosphere component, it provides the pressure and

temperature corresponding to that flight condition. These two parameters, pressure at temperature,

are calculated using a standard definition of atmosphere (ISA 1976).

Figure 32: atmosphere component

As this engine model is working at different altitudes, the temperature and pressure will vary in

function of this variable. The next figures show the evolution of the altitude during the simulation,

and how the temperature and pressure modify its state conditions:

Figure 33: altitude versus time

Figure 34: air temperature and pressure versus time


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5.1.1.2. Inlet and nozzle componentsThese elements are modeled as a resistive element, and there is a dissipative effect which models

the resistance to flow.

There are modeled as pneumatique pipe submodels which have a progressive expansion /

contraction of its section.

This resistive element is going to calculate de mass flow rate and the enthalpy flow rate using the

pressure and temperature as input.

Figure 35: diffuser component

The explanation of how these variables are determined is described in Annexes.

5.1.1.3. Compressor and turbine componentsThis component simulates a gas compressor. This element is a multi-physic component which is

linking mechanical system with pneumatic system.

Figure 36 : GM compressor

This element uses tabulated data imposed by the user to describe the component behavior. These

tabulated data relates pressure ratio (ratio between output pressure and input pressure), air mass

flow and rotatory shaft speed,


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As this element is modeled as a resistive component, it needs pressures of its two pneumatic ports as

input. Using this pressure ratio, the rotatory shaft speed and the tabulated data, the compressor

defines the mass flow rate.

The enthalpy and the mechanical torque are then calculated from the isentropic efficiency of

compressors and turbines which is defined by the user.

Figure 37: Compressor isentropic and non-isentropic transformation

The procedure to calculate its outputs: mass flow rate, the enthalpy rate and the mechanical torque

are described further in the annexes.

5.1.1.4. Combustion chamber componentThis functional model is going to add heat energy flow to the gas flow.

Figure 38 : Combustion chamber component

Using the conservation of energy, the temperature is going to rise because of this heat flow added.


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5.1.2. Hypothesis

5.1.2.1. Hypotheses used Gas Mixture LibraryIn order to develop jet engine models, several pneumatic components were used from the library

called Gas Mixture Library.

The advantage of this library is the capability to use different types of gases, having the possibility to

model systems which use gas mixtures as jet engine combustions.

The different components available from this library are classified as follows:

Figure 39 : Gas Mixture Library schema

The resistive submodels can be described as steady-state submodels. It means that they are assumedto react instantaneously to the temperatures, pressures and fractions applied to them so that they

are always in an equilibrium state.

The different models developed in this work were based using the follow assumptions:

-The gas is following the perfect gas equation:

- Gas behaving as a semi-perfect gas: its specific heat capacities Cp and Cv are only function of gas

temperature.

-The gas density works as an incompressible flow.

5.1.2.2. Hypothesis checksThe first activity done was checking the different hypothesis already defined.

For this mission, AMESim platform allows the user to put several sensors in different system parts.

The different checks done were:

-Checking matter conservation:


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Mass flow rate is constant in all the pneumatic circuit accomplishing matter conservation.

-Checking energy conservation:

Energy absorbed by turbines is powered to compressors, having mechanical losses because of the

inertial masses.

Energy added by the combustion chamber is manifested in the air flow enthalpy.

After the addition of the different sensors, the system became as follows:

Figure 40 : AMESim model


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An energy flow chart schema was developed in order to check visually this energy exchange:

Figure 41 : Jet Engine Dashboard

The system has been classified in different sub-systems:

- External System: it is the boundary air condition. The initial energy that the air has beforeentering to the engine.

- Thermal System: it takes part the combustion chamber which introduces a heat flow of energy tothe pneumatic system.

- Fluid System: it is the pneumatic system; it is related to external system, mechanical system andthermal system.

- Mechanic System: this system is composed of the mechanical support system, which is in chargeof giving the initial mechanical power to start the engine up, the propulsion system, which

provides power to the external unit (the propeller in a helicopter case) and the system which is in

charge of transmit the turbine energy to the compressor.


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5.1.3. Performance Analysis

After hypotheses were checked, the performance system was analyzed.

As seen in previous chapters, the performance of the thermal system is defined as follows:

In order to know how much energy is recovered by the free turbine, they are two ways to calculate

it:

-Using the pneumatic system: to deduce how much energy the air flow is losing through the turbine,

the differences of total enthalpy from both of its pneumatic ports should be calculated.

-Using the mechanical system: the energy the turbine recovers is transformed in mechanical energy.

To deduce the shaft energy, the torque and the rotatory speed are used.

The heat energy flow added is:

Figure 42 : Heat flow added

When we start the adding of heat flow, the quantity of power added is:


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Using the mechanical energy, the quantity of power recovered by the turbine is:

Figure 43 : Turbine Shaft Power

So the performance during the simulation is:

Figure 44 : Thermal Engine Performance


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5.1.4. Component Analysis

An analysis of the behavior of each component has been done in order to understand and find points

to improve the system and getting better global results.

5.1.4.1. InletAs it has been explained in the Main Component Description part and in the annexes, this element ismodeled as a resistive component. It uses the aircraft flight condition (pressure and temperature)

and using the defined geometry of the inlet, it calculates the friction factor using empirical values.

Figure 45 : Inlet parameters

Accounting for the geometry configuration of this element, the inlet calculates output pressure and

temperature.

The first thing done was the verification of the different conditions which this component could work

and, then, a comparison was done between numerical behavior and literature sources.

The parameter that describes the behavior of inlets is the pressure recovery. This parameter links its

output pressure with the stagnation pressure at the intake.

Figure 46 : Inlet stations

In order to check the evolution of this parameter, subsonic and sonic boundary conditions have been

used to compare them.


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Using this inlet geometry:

Diameter at Port 1 (Contraction Side): 500 mm

Diameter at Port 2 (Expansion Side): 600 mm.

Length of element: 500 mm.

Using these boundary conditions,

The follow result comes out of this,

Figure 47 : Inlet Pressure Recovery

This figure shows that the pressure recovery results of the these three cases are .Regarding the literature about jet engine inlets, it can be concluded that this component works

different than the jet engines.

A graphic comparison between the behavior of a real jet engine intake and the behavior defined by

the military standard are showed in the next figure,

Figure 48 : The Intake Pressure Recovery using different hypothesys

Boundary Condition Case 1 Case Case 3

Altitude (m) 1000 1000 1000

Mach 1 2 3


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The pressure recovery falls down its value when the boundary condition gets sonic and supersonic

condition because of compressibility problems.

This element has other points to account for:

- The exterior geometry of the nozzle is a feature that is important for the jet engine globalperformance: the inlet external lips could affect the conditions which the flow could has at the intake

of the compressor. Inlet needs to have an aerodynamical geometric characteristic adapted to the

flight speed of the engine.

Figure 49: Comparison of the properties of the NACA 1- Series and Kchemann class A, B and C lip

fairing

- The drag created by the element is a feature that could be taken into account for the thrust global

gas turbine results.

Figure 50: Different intake configurations (Kchemann y Weber)

Figure 51: Intake flow field at high speed (cruise)


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5.1.4.2. CompressorCompressors are elements defined by performance maps already described in previous chapters. For

that, the real compressors are tested in test rigs, in order to know its performance outside the design

point. This test is based in an actuator who is going to modify the exhaust area in order to modify the

mass flow rate of the jet engine. There are two pressure sensors at each side of this compressor test

rig. This mass flow rate modification allows to check out the performance of the element at different

conditions, at the off-design points.

Figure 52: Compressor speed curve


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A model has been created that represent this test rig,

Figure 53 : Compressor virtual test rig

A linear variation of pressure is applied at the exit of the compressor, as shown in the next figure,

varying the pressure ratio in consequently.

Figure 54 : Condition used for compressor test

Four different cases were executed, changing the revolution number of the compressor shaft.

Case 1 Case 2 Case 3 Case 4

Number of revolutions (rpm) 12000 14000 16000 18000


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The result of this test is showed as follows:

Figure 55: Pressure ratio vs. Mass flow rate

Several things have been realized during this test:

-The tabulated data which is feeding this element does not account compressibility problems like

surge and stall conditions, while this is really important in the jet engine behavior.

Figure 56: Comparison between a typical compressor performance map and the compressor

performance map had as output from my virtual bank test.

-Results show the mass flow rate variable are really low. This element makes the air flow rate raises

its pressure several times so the mass flow should be high. The results show values which are not

larger than 1 kg/s, which are lower to the real values. For example, the average mass flow rate of a

commercial jet engine is about 40 kg/s and for military jet engines these values could get the order of

100 kg/s (or even more). The next figures show a comparison turbojet mass flow rates,


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Engine

Primary

Mass Flow

rate (Kg/s)

Secondary

Mass flow rate

(Kg/s)

JT8D 68,0 74,7

JT3D 88,3 120,0

CFM56-7 55,5 251

Figure 57 : Mass flow rate comparison


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5.1.4.3. TurbineIn order to analyze the behavior of AMESim turbines, a test rig was developed as for compressors.

Figure 58 : Turbine virtual test

Turbine was imposed to four virtual cases. The revolution number of the turbine shaft were defined

as follows,

The pressure ratio was imposed and it was linearly varied during each case,


Number of revolutions (rpm) 12000 14000 16000 18000


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Mass flow rate and recovery power were analyzed in these 4 cases.

The next figure shows the mass flow rate results from theses test cases. The variation of the pressure

ratio and the shaft revolution number do not change the mass flow rate results as is shown in the

figure,

The next figure shows an example of how the mass flow rate is in function of the shaft revolution

number.

Figure 60: Pressure Ratio Vs. Mass flow rate for a

typical Jet Engine turbine

In the rig cases, mass flow rate increases with pressure ratio. However, in the jet engine turbine

example, the mass flow rate behaves in a different way from the virtual one. One of the causes is

that, as for compressor cases, the tabulated data which feeding the turbine does not account

aerodynamical causes.

Figure 59: Pressure Ratio Vs. Mass flow rate for the 4 cases


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5.1.4.4. Other critic pointsOne of the goals of this internship is to develop other gas turbine models, like turbo-propellers or

turbofans.

Several component developments are needed:

-Air-Split: it splits the gas flow in two flows. This component is used in two corps engine like turbo-

props.

-Two corps engine nozzle: it works as a nozzle for engines whose has two air flows.

-Fan: low pressure compressor which works for two corps engines. It is situated before the air split.

-Propeller: it works in turboprop models. It is the element which propels the aircraft using the gas

turbine engine shaft power.

6. A Jet Engine LibraryThe development of the needed component has entailed the creation of a new library dedicated to

gas turbine application: the Jet Engine Library.

6.1.Created Components

6.1.1. Inlet

The intake element of gas turbine was one of the elements to improve in order to get better results.

As it has been described in the previous chapters, this component model has different points to be

improved:

-Pressure Recovery for different Mach speed

-the exterior geometry of the nozzle

-the drag created by the element

For a first level of modeling, the jet engine inlet solves the problem about its pressure recovery

behavior.

Figure 61 : Inlet components

In order to characterize the behavior of the inlet pressure recovery, American Military Standard

named MIL-E-5007D ([2], [5], [7]) are used as reference. These military specifications describe the

evolution of this variable, at different limit conditions, algorithmically.


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The military specifications provide equations which describe the pressure recovery at different Mach

conditions:

Jet engine inlet element has as inputs the aircraft flight conditions:

-Flight altitude

-Mach speed

-Composition gas

The intake element uses the flight conditions (Flight altitude and Mach speed) to calculate the

boundary condition of the inlet. Therefore, it calculates the exit total pressures using the pressure

recovery equations.

6.1.1.1. Element Test - ResultsA model has been developed for test component results.

Figure 62: Inlet Test

In the rig test, they are two components, the created inlet and the atmosphere component. The

same conditions are imposed for both components. The atmosphere component calculates the

stagnation condition that calculates its pressure recovery and is used by the inlet.


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The pressure recovery results, showed in the next figure, describe the expressions based in the

military standard already referenced.

Figure 63: Inlet Pressure Recovery Vs. Mach Number

6.1.1.2. Modeler methodThe method used to model this element was C code, using AMESet to create the .make.

Figure 64: Method used in order to create the inlet AMESim component.

6.1.2.Compressor

6.1.2.1. Mass flow rate problemOne of the problems of the gas turbine model has is the low mass flow rate of the system. As the

system is composed by several components and they contribute to the model results, it is

complicated to find which component affects the most to the mass flow rate of the system.

The first hypothesis was that modifying the performance map of compressors, this variable shouldraise. To check it, the first work done was the modification of its performance map and the testing

the system results.

Figure 65: Compressor performance map


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The results were satisfactory and then, the work was focalized on having the most real

characterization of this component by the performance maps customization.

In the other way, as the final goal is the user utilization, the development of a tool to import the

performance map from other sources or the possibility to modify it in a more visual way are points

that should be considered for this component.

6.1.2.2. IFP Library compressorTo resolve the points related to the compressor performance map, a component which was already

been developed from other library has been used: IFP engine. This component describes a

compressor of an automobile turbo-compressor.

Figure 66: compressor from IFP library

This element was utilized because of its advantage points:

-It uses the standard SAE format to read the performance maps for compressor. SAE (Society

of Automotive Engineers) formats are used in the compressor design industry and it is important to

use the same format as manufacturers do.

-It had already developed a turbo preprocessing tool which, using the performance map input data,

extend the range of the map to the critical regions (low and high speeds, surge, choke).


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Figure 67 : Compressor preprocessing interface

- The element is modeled as an Inertial Resistive element (see Bond Graph theory), so it takes into

account the inertial effect of the exit air flow.

6.1.2.3. Performance Compressor Map DataThe main advantage of this component is the preprocessing and reading performance map

capability. However, in order to verify the adaptation of this component in an aeronautic system it is

needed examples of real performance maps.

These performance maps were found in a gas turbine design program called GasTurb. This software

allows the static simulations of different aero gas turbine models.

It is software very used in this field, so the possibility to work along the same line of this program

makes the jet engine library more interesting for users.

Figure 68 : Gas Turb software


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This software and its derived software (Smooth C and Smooth T) give some performance map

examples from some real compressors and let the user the possibility to create compressor or

turbine performance map (using Smooth C and Smooth T) by oneself.

Figure 69 : GasTurb compressor performance map example

However, GasTurb uses a different format of AMESim to describe numerically performance

compressor maps. In order to be able to use its maps, a preprocessing script has been created. This

script allows adapting its format to AMESim format (SAE format for this case).

Used Python language for developing this script.

Figure 70: Script created in order to convert performance compressor maps using Python language


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6.1.2.4. Inertial Resistive componentThis component is defined as an inertial resistive component. It takes into account the inertial effect

of the air flow: The volume of gas between the diffuser outlet and the downstream volume induces a

difference between the actual pressure ration and the pressure ratio read in the table during

transients.

6.1.2.4.1. Centrifugal compressor

As its compressor is defined as a centrifugal compressor, in order to take this inertial effect into

account, it uses specific geometric characteristic in the derivative equation.

Using this compressor geometry characteristic:

Figure 71 : Centrifuge compressor schema

Rimpeller: impeller outer radius [mm], Ddiffuser: diffuser mean diameter [mm], Ldiff: equivalent length of the diffuser as a function of the shape of the compressor:

o vaneless: Ldiff = 2 Rimpeller 4.3o vaned: Ldiff = 2 Rimpeller 5.3

This element was defined as a centrifugal compressor. In order to model an axial compressor, the

adaptation of the compressor geometric characteristic and the derivative equations are necessary.


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6.1.2.4.2. Axial compressor

The axial compressor has different geometric characteristic, as is shown as follows,

Figure 72: axial compressor for Jet Engine

The development of the equations which describe the axial compressor inertial behavior should be

modified. This development was done for the jet engine library (development explained in the

annexes Compressor/ Turbine modeling), being the specified equation for axial compressor as

follows,

, : Absolute total pressure at diffuser outlet,: Actual downstream absolute total pressure.L : axial length of last stage stator

S: Axial section of the stator station

SRSR

Rint

Rext

L


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6.1.2.5. Calculation stepsIn order to calculate its output (enthalpy flow and mass flow rate), and example for axial

compressors is developed:

-Taking the axial compressor example:

-The pressure at diffuser outlet is defined as follow,

-Pressure ratio is calculated using the tabulated data and the mass flow rate so,

Using these 3 equations we have: This expression is an explicit equation, so, in order to calculate it in AMESim, the mass flow rate is

declared as an explicit variable.


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6.1.2.6. Element Test - ResultsThe first test done was the converter script check. For this test, it was used a Gas Turb compressor

example, Turbine NASA TM 101433, a HP compressor,

Figure 73: 99 NASA TM 101433 performance map data

The converter script analyses the compressor maps and converts it in the AMESim format. It shows

graphically the conversion it does so as to make it clear to the user.

Figure 74: Script check


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After that, the preprocessing interface inhered to the compressor component is launched in order to

check if its data are the same of the data imported from GasTurb model,

Figure 75: Performance Map check

When the preprocessing tools are validated, the component global test should be created. For that, a

test has been created varying the revolution number of the compressor between 7000 and 12000

rpm imposed as follows,

Figure 76: Revolution number imposed


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During each test case the pressure at both ports are imposed to simulate the high pressure

compressor limit condition. The different pressures imposed are described as follows,


Enter Pressure (barA) 5 5 5 5

Output Pressure (barA) 8 10 12 14

Figure 77: Compressor test rig

The mass flow rate evolutions in function of the revolution number imposed. This last variable

decreases linearly, and for each case, there is a condition where the compressor does not work

correctly and the mass flow inverses its direction. There are the same consequences as whencompressor works in a surge or stall condition.

Figure 78: Compressor mass flow from test rig


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The conclusion of this isolate test rig is that the compressor behaves better than the last ones. The

mass flow is increased and its values are coherent to real jet engine mass flows.


Picture 79: phases of compressor creation

6.1.3. Turbine

As the compressor, there was a turbine component already developed for automobile turbo-

compressor solutions.

Picture 80: Jet Engine Turbine

As the compressor, this component has its two advantage points:

-It uses the standard SAE format to read the performance maps for turbine. This standard SAE

performance map follows the SAE J922 specifications as compressors.

-It had already developed a turbo preprocessing tool which, using the performance map input data,

extend the range of the map to the critical regions (low and high speeds, surge, choke).


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Figure 81: Turbine Preprocessing Interface


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6.1.3.1. Performance Turbine Map DataThis component was developed using the same procedures as for compressors. Another Python

script was developed in order to adapt the performance turbine map data from GasTurb to AMESim

format.

Figure 82: GasTurb example turbine map

6.1.3.2. Resistive component Calculation stepsThis element is considerate as a resistive component (see Bond Graph section). The steps used to

calculate this output variables (mass flow rate and enthalpy flow) are:

- Using the pressure at both ports, it calculates the Pressure Ratio.

- Using the PR and the tabulated data, it calculates mass flow rate and the efficacy.

- The enthalpy is calculated using turbine performance and mass flow rate (as described in GM

Turbine component).


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6.1.3.3. Element Test - ResultsThe first test done was the converter script check.

For this test, it was used a Gas Turb compressor example, Turbine NASA TM83655, a HP turbine:

Figure 83: NASA TM83655 performance map data

The converter script analyses the compressor maps and converts it in an AMESim format. It shows

graphically the conversion done by the Python script to the user.

Figure 84: Script check


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After that, the preprocessing interface inhered to the compressor component is launched in order to

check if its performance maps correspond to the data imported from GasTurb model.

Figure 85: Performance Map check

When the preprocessing tools are validated, the component global test should be created. For that, a

test has been created varying the revolution number of the turbine between 7000 and 12000 rpm

imposed as follows,

Figure 86: Revolution number imposed


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During each test case the pressure at both ports are imposed to simulate the high pressure

compressor limit condition. The different pressures imposed are described as follows,


Enter Pressure (barA) 8 10 12 14

Output Pressure (barA) 5 5 5 5

Figure 87: Compressor test rig

The mass flow rate evolutions in function of the revolution number imposed. This last variabledecreases linearly. The mass flow rate increases while the revolution number increases as its

performance map describes.

Figure 88: Compressor mass flow from test rig


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Figure 89: phases of turbine creation


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6.1.4.Air Split

The air split is the component which splits the flow in two parts: the main and the secondary flow.

This component is used in the two corps gas turbine engines as turboprops (see Jet Propulsion

systems chapter)

In order to model this functionality some gas mixture components have been used to create a super-component.

Figure 90: Air Split component

The component developed to model this function needs the geometry section of its different exits.

Using these geometry characteristics, the mass flow rate of each flow will vary.

Figure 91: Air Split parts

The variation of these sections will influence a parameter significant for the engine performance, the

By-Pass Ratio (BPR): it relates the secondary and the primary mass flows.


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6.1.4.1. Element Test - ResultsAn isolate test to check the split behavior was developed. A source is used in this test rig. They

provide a range of mass flow (from 0 to 10 kg/s) and at the component exits they are two sources

which impose a see level boundary condition.

Figure 92 : Air Split test

The Air Split geometry characteristics are defined as follows:

Secondary Area= 0.7 Primary Area= 1 The boundary conditions are defined as follows:

Enter Primary exit Secondary exit

Temperature (K) 290 290 290

Pressure (barA) 1.013 1.013 1.013

Looking at the follow graphics, one can realizes that the mass flow is divided in two parts. One of

both parts is higher than the other because the section is bigger.

Figure 93: results from air split test


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6.1.4.2. Modeler methodThe method used to model this element was the creation of a super-component. AMESim allows to

link some components in order to simulate the element functionality interested to model.

In this case, one can see how these three components will create the air-split function.

Figure 94: component to create air-split function

When the components, which one wants to link, are correctly linked, one can call the AMESim

function Create supercomponent.

Figure 95: AMESim options

Automatically, AMESim will create a supercomponent using the configuration selected:

Figure 96: Air-split supercomponent


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6.1.5. Nozzle Double Corp

The double flow nozzle is an element which models the nozzle for double flow engines. This

component is used in the two corps gas turbine engines as turboprops (see Jet Propulsion systems

chapter).

Figure 97: jet engine nozzle component

Like air split component, a supercomponent is used to model its functionality.

Figure 98: nozzle component

The user has to set the geometry characteristic of both nozzles: enter and exit sections.

6.1.5.1. Element Test ResultsAn isolate test is developed to check the nozzle behavior. For it, two sources are created which

provide two constant mass flows and at the component exits, a source imposes the exterior

conditions.

Figure 99: Nozzle test


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The nozzle geometry characteristics are defined as follows:

Core nozzle Fan nozzle

Enter diameter (mm) 500 500

Exit diameter (mm) 300 300

Length (mm) 2000 2000

The boundary conditions are defined as follows:

Primary flow Secondary flow External condition

Mass Flow rate (Kg/s) 0 to 80 0 to 50 -

Temperature (K) 700 350 293

Pressure (barA) - - 1.013

As one can see, looking at the follow graphics, the mass flow is divided in two parts. In order to

simulate the startup of the engine, both mass flow rates, the primary flow and the secondary one,

has been linearly increased during the simulation.

Figure 100: mass flow rates imposed during the test

The result, in the next figure, shows the mass flow rate at the exit of the double nozzle component.


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Figure 101: results from double flow nozzle test

6.1.5.2. Modeler methodThe method used to create the nozzle model is the same as the air-split model.

Figure 102: nozzle supercomponent

6.1.6. Propeller

In order to model the turbine gas which neeed a propulsion component to propel the aircraft as

turboprops, it is mandatory to create a propeller component.

Propeller is a propulsive element which is used to create aerodynamic traction using mechanical

energy.

During this internship two different levels of propeller modeling have been developed.

6.1.6.1. First level of modelingThis first level of modeling uses the blade element theory to calculate the thrust and the resistive

force it is going to develop at a given shaft revolution number.

Figure 103: propeller model


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6.1.6.1.1. Blade element theory

The blade element theory let you to calculate the lift and the drag that an aerodynamic profile will

provide at a defined boundary condition.

This theory uses the geometrical properties of the blade to determine the forces exerted by a

propeller on the flow-field.

The equations used for this component are described as follow:

Figure 104: Blade element theory schema

6.1.6.1.1.1. Aerodynamic information needed

In order to calculate the lift and drag provided by the blade, this method needs its aerodynamic

behavior information. It needs the evolution of Cl and Cd which are function of the attack angle of

the pale.

In order to supply this information, we used a software called XFoil which provides the information

of a large kind of profile which have already been developed.

Figure 105: Airfoil software


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The problem found at that moment was that the software format to export its results was not the

same of AMESim format. In order to adapt this one to the AMESim format, in the propeller code has

been added a part which modifies the data file.

6.1.6.1.2. Model Limitations

Tangential speed is varying linearly in function of the blade radium. The m

internship report david jimenez

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