**Meet the editor**

Dr. Ernesto Benini is currently a Research Professor of Fluid Machinery at the Department of Mechanical Engineering, University of Padova, Italy, where he first obtained his MSc in Mechanical Engineering and then his PhD in Energy Technology. His research experience is mainly focused on gas turbine engine design and optimization with particular emphasis on aero-thermodynamics

of turbomachinery. He authored more than 120 research articles on these topics, the largest part on international refereed journals and congresses.

Contents

**Preface IX** 

**Part 1 Aero and Marine Gas Turbines 1** 

Konstantinos G. Kyprianidis

Chapter 3 **Possible Efficiency Increasing** 

**Part 2 Gas Turbine Systems 69** 

Chapter 5 **Flexible Micro Gas Turbine Rig** 

Marek Dzida

Chapter 4 **Exergy Analysis of** 

Chapter 1 **Future Aero Engine Designs: An Evolving Vision 3** 

Chapter 2 **State-of-Art of Transonic Axial Compressors 25**  Roberto Biollo and Ernesto Benini

> **of Ship Propulsion and Marine Power Plant with the System Combined of Marine Diesel Engine, Gas Turbine and Steam Turbine 45**

**a Novel SOFC Hybrid System with Zero-CO2 Emission 71** 

Marco Antônio Rosa do Nascimento and Eraldo Cruz dos Santos

Liqiang Duan, Xiaoyuan Zhang and Yongping Yang

**for Tests on Advanced Energy Systems 89** 

**a Cogeneration Power Plant with Gas Turbine by Numerical and Experimental Analysis 139**  Ene Barbu, Valeriu Vilag, Jeni Popescu, Silviu Ionescu, Adina Ionescu, Romulus Petcu, Cleopatra Cuciumita, Mihaiella Cretu, Constantin Vilcu and Tudor Prisecaru

Mario L. Ferrari and Matteo Pascenti

Chapter 6 **Biofuel and Gas Turbine Engines 115** 

Chapter 7 **Afterburning Installation Integration into** 

## Contents

## **Preface** XIII


#### **Part 2 Gas Turbine Systems 69**

	- **Part 3 Heat Transfer 189**

Contents VII

Chapter 19 **Damageability of Gas Turbine Blades** 

Chapter 20 **New Non-Destructive Methods** 

**– Evaluation of Exhaust Gas Temperature**

Józef Błachnio and Wojciech Izydor Pawlak

**in Front of the Turbine Using a Non-Linear Observer 435**

**of Diagnosing Health of Gas Turbine Blades 465**  Józef Błachnio, Mariusz Bogdan and Artur Kułaszka

Chapter 21 **Repair of Turbine Blades Using Cold Spray Technique 499**

Kazuhiro Ogawa and Dowon Seo

	- **Part 4 Combustion 237**
	- **Part 5 Materials and Fabrication 291**

VI Contents

Chapter 8 **Application of Statistical Methods** 

Chapter 9 **Jet Impingement Cooling in Gas Turbines** 

**on Gas Turbine Performance 211**

Chapter 11 **Developments of Gas Turbine Combustors** 

Chapter 12 **Characterization of a Spray in the Combustion** 

Chapter 13 **Materials for Gas Turbines – An Overview 293** Nageswara Rao Muktinutalapati

Chapter 16 **Unidirectionally Solidified Eutectic Ceramic Composites** 

**for Ultra-High Efficiency Gas Turbine Systems 371**

Chapter 14 **Titanium in the Gas Turbine Engine 315**

Chapter 15 **Platinum-Based Alloys and Coatings: Materials for the Future? 337** Lesley A. Cornish and Lesley H. Chown

Chapter 17 **Study of a New Type High Strength** 

Enze Liu and Zhi Zheng

Antonio M. Mateo García

**Ni-Based Superalloy DZ468 with Good Hot Corrosion Resistance 399** 

Chapter 18 **BLISK Fabrication by Linear Friction Welding 411** 

Takeharu Hasegawa

Georges Descombes

Mark Whittaker

Yoshiharu Waku

**Part 5 Materials and Fabrication 291** 

Li Pan

**Part 3 Heat Transfer 189**

Chapter 10 **Influence of Heat Transfer** 

**Part 4 Combustion 237** 

**for Gas Turbine Plant Operation Monitoring 165**

Diango A., Périlhon C., Danho E. and Descombes G.

**for Air-Blown and Oxygen-Blown IGCC 239** 

**Chamber of a Low Emission Gas Turbine 267**

**for Improving Thermal Efficiency and Power Density 191** Luai M. Al-Hadhrami, S.M. Shaahid and Ali A. Al-Mubarak

> Chapter 21 **Repair of Turbine Blades Using Cold Spray Technique 499**  Kazuhiro Ogawa and Dowon Seo

Preface

generation equipment.

World energy demand is likely to increase over the next 20 years, and it is well ascertained that fossil fuels will still be the dominant source for power generation all over the world. In this scenario, gas turbine (GT) engines will still represent a key technology, either in stand-alone applications or combined with other power

The challenges in GT technology today deal with several issues, such as increased on/off design efficiency, reduced performance degradation over time, and decreased pollutant emissions levels. Major research efforts and investments are being bestowed for the development of new, advanced GT technologies with superior performance, thus helping in the fulfillment of the Kyoto Protocol objectives for greenhouse gases reduction, as well as of many other transnational policies on sustainability and

The book intends to provide an updated picture, as well as a perspective view of some of the major improvements which characterize the GT technology in different applications, from marine and aircraft propulsion to industrial and stationary power generation. Therefore, the target audience for it involves design, analyst, materials and maintenance engineers. Manufacturers, researchers and scientists will also benefit

The book is organized into five main sections, which are comprised of 21 chapters: (I) Aero and Marine Gas Turbines, (II) Gas Turbine Systems, (III) Heat Transfer, (IV) Combustion, and (V) Materials and Fabrication. Starting from a general but evolving vision on GT and their components used in aircraft and marine engines in Section I, a systematic approach is used in Section II to describe the thermodynamic behavior of various GT-based technologies, including combined power, hybrid fuel cell/GT and small-scale GT plants. Then, in Sections III and IV, a total of four specific works (two articles each) are included, which present some of the latest developments in the internal heat transfer and combustion-related phenomena. Finally, a collection of nine papers in Section V. deal with the most recent advances regarding the new materials for GT component fabrication, including interesting suggestions on material damage

reduced environmental impact of energy technologies.

from the timely and accurate information provided.

prevention, diagnosis and repair.

## Preface

World energy demand is likely to increase over the next 20 years, and it is well ascertained that fossil fuels will still be the dominant source for power generation all over the world. In this scenario, gas turbine (GT) engines will still represent a key technology, either in stand-alone applications or combined with other power generation equipment.

The challenges in GT technology today deal with several issues, such as increased on/off design efficiency, reduced performance degradation over time, and decreased pollutant emissions levels. Major research efforts and investments are being bestowed for the development of new, advanced GT technologies with superior performance, thus helping in the fulfillment of the Kyoto Protocol objectives for greenhouse gases reduction, as well as of many other transnational policies on sustainability and reduced environmental impact of energy technologies.

The book intends to provide an updated picture, as well as a perspective view of some of the major improvements which characterize the GT technology in different applications, from marine and aircraft propulsion to industrial and stationary power generation. Therefore, the target audience for it involves design, analyst, materials and maintenance engineers. Manufacturers, researchers and scientists will also benefit from the timely and accurate information provided.

The book is organized into five main sections, which are comprised of 21 chapters: (I) Aero and Marine Gas Turbines, (II) Gas Turbine Systems, (III) Heat Transfer, (IV) Combustion, and (V) Materials and Fabrication. Starting from a general but evolving vision on GT and their components used in aircraft and marine engines in Section I, a systematic approach is used in Section II to describe the thermodynamic behavior of various GT-based technologies, including combined power, hybrid fuel cell/GT and small-scale GT plants. Then, in Sections III and IV, a total of four specific works (two articles each) are included, which present some of the latest developments in the internal heat transfer and combustion-related phenomena. Finally, a collection of nine papers in Section V. deal with the most recent advances regarding the new materials for GT component fabrication, including interesting suggestions on material damage prevention, diagnosis and repair.

#### XIV Preface

The editor is indebted to all the valuable contributions included in this book from the various GT experts spread worldwide, as well as to InTech Open Access Publisher for giving me the opportunity to edit this volume, and supporting me constantly during its preparation.

> **Dr. Ernesto Benini** Department of Mechanical Engineering University of Padova, Italy

**Part 1** 

**Aero and Marine Gas Turbines** 

**0**

**1**

*Sweden*

Konstantinos G. Kyprianidis *Chalmers University of Technology*

**Future Aero Engine Designs: An Evolving Vision**

Public awareness and political concern over the environmental impact of civil aviation growth has improved substantially during the past 30 years. As the environmental awareness increases, so does the effort associated with addressing NO*<sup>x</sup>* and CO2 emissions by all the parties involved. In the Vision 2020 report made by the Advisory Council for Aeronautical Research in Europe (2001), goals are set to reduce noise and emissions produced by the ever increasing global air traffic. Emissions legislation, set by the International Civil Aviation Organisation (ICAO) and it's Committee on Aviation Environmental Protection (CAEP), is becoming ever more stringent, creating a strong driver for investigating novel aero engine

On the other hand, airline companies need to continuously reduce their operating costs in order to increase, or at least maintain, their profitability. This introduces an additional design challenge as new aero engine designs need to be conceived for reduced environmental impact as well as direct operating costs. Decision making on optimal engine cycle selection needs to consider mission fuel burn, direct operating costs, engine and airframe noise, emissions and

CO2 emissions are directly proportional to fuel burn, and therefore any effort to reduce them needs to focus on improving fuel burn, by reducing engine Specific Fuel Consumption (SFC), weight and size. Reducing engine weight results in a lower aircraft maximum take-off weight, which in turn leads to reduced thrust requirements for a given aircraft lift to drag ratio. Reducing engine size – predominantly engine nacelle diameter and length – reduces nacelle drag and therefore also leads to reduced thrust requirements. For a given engine SFC, a reduction in thrust requirements essentially results in lower fuel burn. Lower engine SFC can be achieved by improving propulsive efficiency and thermal efficiency – either by reducing

Improvements in propulsive efficiency – and hence engine SFC at a given thermal efficiency – can be achieved by designing an engine at a lower specific thrust (i.e. net thrust divided by fan inlet mass flow). This results in a larger fan diameter, at a given thrust, and therefore in increased engine weight, which can partially, or even fully, negate any SFC benefits. Propulsive efficiency improvements at a constant weight are directly dependent on weight reduction technologies such as light weight fan designs and new shaft materials. Increasing engine bypass ratio aggravates the speed mismatch between the fan and the low pressure turbine. Introduction of a gearbox can relieve this issue by permitting the design of these two

**1. Introduction**

global warming impact.

designs that produce less CO2 and NO*<sup>x</sup>* emissions.

component losses or by improving the thermodynamic cycle.

## **Future Aero Engine Designs: An Evolving Vision**

Konstantinos G. Kyprianidis *Chalmers University of Technology Sweden*

#### **1. Introduction**

Public awareness and political concern over the environmental impact of civil aviation growth has improved substantially during the past 30 years. As the environmental awareness increases, so does the effort associated with addressing NO*<sup>x</sup>* and CO2 emissions by all the parties involved. In the Vision 2020 report made by the Advisory Council for Aeronautical Research in Europe (2001), goals are set to reduce noise and emissions produced by the ever increasing global air traffic. Emissions legislation, set by the International Civil Aviation Organisation (ICAO) and it's Committee on Aviation Environmental Protection (CAEP), is becoming ever more stringent, creating a strong driver for investigating novel aero engine designs that produce less CO2 and NO*<sup>x</sup>* emissions.

On the other hand, airline companies need to continuously reduce their operating costs in order to increase, or at least maintain, their profitability. This introduces an additional design challenge as new aero engine designs need to be conceived for reduced environmental impact as well as direct operating costs. Decision making on optimal engine cycle selection needs to consider mission fuel burn, direct operating costs, engine and airframe noise, emissions and global warming impact.

CO2 emissions are directly proportional to fuel burn, and therefore any effort to reduce them needs to focus on improving fuel burn, by reducing engine Specific Fuel Consumption (SFC), weight and size. Reducing engine weight results in a lower aircraft maximum take-off weight, which in turn leads to reduced thrust requirements for a given aircraft lift to drag ratio. Reducing engine size – predominantly engine nacelle diameter and length – reduces nacelle drag and therefore also leads to reduced thrust requirements. For a given engine SFC, a reduction in thrust requirements essentially results in lower fuel burn. Lower engine SFC can be achieved by improving propulsive efficiency and thermal efficiency – either by reducing component losses or by improving the thermodynamic cycle.

Improvements in propulsive efficiency – and hence engine SFC at a given thermal efficiency – can be achieved by designing an engine at a lower specific thrust (i.e. net thrust divided by fan inlet mass flow). This results in a larger fan diameter, at a given thrust, and therefore in increased engine weight, which can partially, or even fully, negate any SFC benefits. Propulsive efficiency improvements at a constant weight are directly dependent on weight reduction technologies such as light weight fan designs and new shaft materials. Increasing engine bypass ratio aggravates the speed mismatch between the fan and the low pressure turbine. Introduction of a gearbox can relieve this issue by permitting the design of these two

(1976); the author has also provided an update to that discussion based on current economical and technological projections (Jackson, 2009). Wilde (1978), Young (1979), and Pope (1979) provide a good reference on how the future for civil turbofan engines for medium and long range applications was envisaged in the late 70's. Some early discussions on future trends in commercial aviation from the aircraft manufacturer's perspective can be found in Swihart (1970) and Bates & Morris (1983), while Watts (1978) provides an airliner's view of the future. A review on the several technical and economic obstacles that were identified in the late 80's with respect to the realization of the Ultra-High Bypass Ratio (UHBR) turbofan concept is provided by Borradaile (1988) and by Zimbrick & Colehour (1988). Peacock & Sadler (1992) give an update on the subject, focusing further on engine design constraints and the technology advancements required for producing a competitive UHBR configuration. Potential year 2020 scenarios are explored by Birch (2000) while an overview of current aero engine technology and some insight on the future of aircraft propulsion is given by Ruffles (2000). Sieber (1991) and Schimming (2003) provide an excellent discussion on counter-rotating fan designs. Finally, for a review on the development of civil propulsion from the early 50's to recent years the interested reader is referred to Saravanamuttoo (2002). The focus of the next section will be given on recent European research initiatives on enabling

Future Aero Engine Designs: An Evolving Vision 5

Within the EU Framework Program 6 research project VITAL (enVIronmenTALly friendly aero engines, 2009) a number of low pressure system component technologies have been investigated (Korsia, 2009; Korsia & Guy, 2007). The emerging progress will allow the design of new powerplants capable of providing a step reduction in fuel consumption and generated

The VITAL project concentrated on new technologies for the low pressure system of the engine, which enable the development of low noise and low weight fan architectures for UHBR engines. To achieve these objectives, the VITAL project has investigated three different low pressure configurations, leading to low noise and high efficiency power plants. The three configurations are the DDTF (Direct Drive TurboFan) supported by Rolls-Royce, the GTF (Geared TurboFan) by MTU and the CRTF (Counter-Rotating TurboFan) by Snecma. The DDTF architecture offers a re-optimised trade-off between fan and turbine requirements considering the low weight technologies introduced by the VITAL programme. The GTF combines a fan with a reduction gear train, to allow different rotating speeds for the fan on one hand, and the booster and turbine on the other. The CRTF offers a configuration with two fans turning in opposite directions, allowing for lower rotational speeds, since the two fan

The technologies being built into the VITAL engines include (Korsia, 2009; Korsia & Guy,

• New fan concepts with the emphasis on two types: counter-rotating and lightweight fans. • New booster technologies for different operational requirements; low and high speed, associated aerodynamic technologies, new lightweight materials and associated coating

technologies relevant to the three research questions that have been set.

**3. Enabling technologies and recent research**

**3.1 Propulsor technologies**

rotors split the loads involved.

and noise reduction design.

noise.

2007):

components at their optimal speeds, and can hence reduce engine weight, as well as improve component efficiency. The first research question therefore rises:

#### *How low can we really go on specific thrust?*

Improvements in thermal efficiency – and hence engine SFC at a given propulsive efficiency – can be achieved for conventional cores mainly by increasing engine Overall Pressure Ratio (OPR). At a given OPR there is an optimal level of combustor outlet temperature T4 for thermal efficiency. However, at a fixed specific thrust and engine thrust, an increase in T4 can result in a smaller core and therefore a higher engine bypass ratio; in some cases, a potential reduction in engine weight can more than compensate for a non-optimal thermal efficiency. Increasing OPR further than current engine designs is hindered by limitations in high pressure compressor delivery temperature at take-off. Increasing T4 is limited by maximum permissable high pressure turbine rotor metal temperatures at take-off and top of climb. Increasing turbine cooling flows for this purpose is also fairly limited as a strategy; cooling flows essentially represent losses in the thermodynamic cycle, and increasing them eventually leads to severe thermal efficiency deficits (Horlock et al., 2001; Wilcock et al., 2005). Designing a combustor at very low air to fuel ratio levels is also limited by the need for adequate combustor liner film-cooling air as well as maintaining an acceptable temperature traverse quality (Lefebvre, 1999). The second research question therefore rises:

#### *How high can we really go on OPR and T*4*?*

Aggressive turbofan designs that reduce CO2 emissions – such as increased OPR and T4 designs – can increase the production of NO*<sup>x</sup>* emissions due to higher flame temperatures. The third research question therefore rises:

#### *What is the trade-off between low CO*<sup>2</sup> *and NOx?*

The research work presented in this chapter will focus on identifying several novel engine cycles and technologies - currently under research - that can address the three research questions raised. These concepts will be evaluated based on their potential to reduce CO2 and NO*<sup>x</sup>* emissions for engine designs entering service between 2020 and 2025. Design constraints, material technology, customer requirements, noise and emissions legislation, technology risk and economic considerations and their effect on optimal concept selection will also be discussed in detail.

#### **2. An evolving vision**

Numerous feasibility studies have been published over the years focusing on future engine and aircraft designs that can reduce fuel consumption; a brief review of some of these publications will be carried out here.

One of the earliest discussions on the subject of improving engine fuel efficiency is provided by Gray & Witherspoon (1976), looking at conventional and heat exchanged cores, as well as non-steady flow combustion processes and open rotor configurations. A similar study focusing on geared and open rotor arrangements as well as heat exchanged cycles is presented by Hirschkron & Neitzel (1976).

An interesting discussion on how specific thrust levels were expected to evolve in the mid-70's based on the economic and technological projections of that time period is given by Jackson 2 Will-be-set-by-IN-TECH

components at their optimal speeds, and can hence reduce engine weight, as well as improve

Improvements in thermal efficiency – and hence engine SFC at a given propulsive efficiency – can be achieved for conventional cores mainly by increasing engine Overall Pressure Ratio (OPR). At a given OPR there is an optimal level of combustor outlet temperature T4 for thermal efficiency. However, at a fixed specific thrust and engine thrust, an increase in T4 can result in a smaller core and therefore a higher engine bypass ratio; in some cases, a potential reduction in engine weight can more than compensate for a non-optimal thermal efficiency. Increasing OPR further than current engine designs is hindered by limitations in high pressure compressor delivery temperature at take-off. Increasing T4 is limited by maximum permissable high pressure turbine rotor metal temperatures at take-off and top of climb. Increasing turbine cooling flows for this purpose is also fairly limited as a strategy; cooling flows essentially represent losses in the thermodynamic cycle, and increasing them eventually leads to severe thermal efficiency deficits (Horlock et al., 2001; Wilcock et al., 2005). Designing a combustor at very low air to fuel ratio levels is also limited by the need for adequate combustor liner film-cooling air as well as maintaining an acceptable temperature

traverse quality (Lefebvre, 1999). The second research question therefore rises:

Aggressive turbofan designs that reduce CO2 emissions – such as increased OPR and T4 designs – can increase the production of NO*<sup>x</sup>* emissions due to higher flame temperatures.

The research work presented in this chapter will focus on identifying several novel engine cycles and technologies - currently under research - that can address the three research questions raised. These concepts will be evaluated based on their potential to reduce CO2 and NO*<sup>x</sup>* emissions for engine designs entering service between 2020 and 2025. Design constraints, material technology, customer requirements, noise and emissions legislation, technology risk and economic considerations and their effect on optimal concept selection

Numerous feasibility studies have been published over the years focusing on future engine and aircraft designs that can reduce fuel consumption; a brief review of some of these

One of the earliest discussions on the subject of improving engine fuel efficiency is provided by Gray & Witherspoon (1976), looking at conventional and heat exchanged cores, as well as non-steady flow combustion processes and open rotor configurations. A similar study focusing on geared and open rotor arrangements as well as heat exchanged cycles is presented

An interesting discussion on how specific thrust levels were expected to evolve in the mid-70's based on the economic and technological projections of that time period is given by Jackson

component efficiency. The first research question therefore rises:

*How low can we really go on specific thrust?*

*How high can we really go on OPR and T*4*?*

*What is the trade-off between low CO*<sup>2</sup> *and NOx?*

The third research question therefore rises:

will also be discussed in detail.

publications will be carried out here.

by Hirschkron & Neitzel (1976).

**2. An evolving vision**

(1976); the author has also provided an update to that discussion based on current economical and technological projections (Jackson, 2009). Wilde (1978), Young (1979), and Pope (1979) provide a good reference on how the future for civil turbofan engines for medium and long range applications was envisaged in the late 70's. Some early discussions on future trends in commercial aviation from the aircraft manufacturer's perspective can be found in Swihart (1970) and Bates & Morris (1983), while Watts (1978) provides an airliner's view of the future. A review on the several technical and economic obstacles that were identified in the late 80's with respect to the realization of the Ultra-High Bypass Ratio (UHBR) turbofan concept is provided by Borradaile (1988) and by Zimbrick & Colehour (1988). Peacock & Sadler (1992) give an update on the subject, focusing further on engine design constraints and the technology advancements required for producing a competitive UHBR configuration. Potential year 2020 scenarios are explored by Birch (2000) while an overview of current aero engine technology and some insight on the future of aircraft propulsion is given by Ruffles (2000). Sieber (1991) and Schimming (2003) provide an excellent discussion on counter-rotating fan designs. Finally, for a review on the development of civil propulsion from the early 50's to recent years the interested reader is referred to Saravanamuttoo (2002). The focus of the next section will be given on recent European research initiatives on enabling technologies relevant to the three research questions that have been set.

#### **3. Enabling technologies and recent research**

#### **3.1 Propulsor technologies**

Within the EU Framework Program 6 research project VITAL (enVIronmenTALly friendly aero engines, 2009) a number of low pressure system component technologies have been investigated (Korsia, 2009; Korsia & Guy, 2007). The emerging progress will allow the design of new powerplants capable of providing a step reduction in fuel consumption and generated noise.

The VITAL project concentrated on new technologies for the low pressure system of the engine, which enable the development of low noise and low weight fan architectures for UHBR engines. To achieve these objectives, the VITAL project has investigated three different low pressure configurations, leading to low noise and high efficiency power plants. The three configurations are the DDTF (Direct Drive TurboFan) supported by Rolls-Royce, the GTF (Geared TurboFan) by MTU and the CRTF (Counter-Rotating TurboFan) by Snecma.

The DDTF architecture offers a re-optimised trade-off between fan and turbine requirements considering the low weight technologies introduced by the VITAL programme. The GTF combines a fan with a reduction gear train, to allow different rotating speeds for the fan on one hand, and the booster and turbine on the other. The CRTF offers a configuration with two fans turning in opposite directions, allowing for lower rotational speeds, since the two fan rotors split the loads involved.

The technologies being built into the VITAL engines include (Korsia, 2009; Korsia & Guy, 2007):


Fig. 2. Compressor efficiency improvement with year of entry into service.

very good axial compressor design is illustrated in Fig.2.

that include (Rolt & Kyprianidis, 2010; Wilfert et al., 2007):

• Active control of surge and tip clearance in compressors.

• Active control of a cooled cooling air system.

Standard, and Allison.

demonstrator.

**3.2 Core technologies**

propellers driven by a more conventional low pressure turbine through a differential planetary differential gearbox, was the result of a joint effort by Pratt & Whitney, Hamilton

Future Aero Engine Designs: An Evolving Vision 7

Both projects were eventually put on hold towards the end of the decade as fuel prices fell significantly. Nevertheless, the open rotor concept has now resurfaced within the EU Framework Program 7 research project DREAM (valiDation of Radical Engine Architecture systeMs, 2011) and the Clean Sky Joint Technology Initiative (2011). Within DREAM, the feasibility of two different open rotor architectures is evaluated including noise. Within Clean Sky, research work is being carried out by some of Europe's largest aero engine manufacturers, such as Rolls-Royce and Snecma, focused on designing, building and testing an open rotor

Improving core component efficiencies (including reducing losses in the cycle such as duct pressure losses) is one way of improving the engine thermal efficiency. Nevertheless, modern CFD-assisted designs are already quite aggressive and limited benefit may be envisaged by such future advancements (Kurzke, 2003); the increasing effort required to improve an already

Within the EU Framework Program 6 research project NEWAC (NEW Aero engine Core concepts, 2011) a number of advanced core component technologies have been investigated

• Improved high pressure compressor aero design and blade tip rub management.

• Flow control technologies including aspirated compression systems.

Fig. 1. Effect of fan tip pressure ratio and bypass duct pressure losses on fan equivalent polytropic efficiency


The open rotor engine concept, for high subsonic flight speeds, has also risen as a candidate for improving fuel consumption on several occasions since the advent of the first high bypass ratio turbofan engine. Such engine configurations, often refereed to as propfans in the literature, are direct competitors to ultra high bypass ratio turbofan engines. Their are located at the ultra-low specific thrust region of the design space, where propulsive efficiency benefits for turbofans are negated by very low transfer efficiencies. As illustrated in Fig. 1, this is due to the dominant effect on transfer efficiency that bypass duct pressure losses have when looking at low fan tip pressure ratio engine designs, i.e. low specific thrust. Open rotor engines do not suffer from bypass duct pressure losses and can therefore achieve a very high propulsive efficiency at a good level of transfer efficiency. Compared to turbofans, propfans also benefit from reduced nacelle drag and weight penalties.

Several open rotor programs took place during the 80's, resulting in engine demonstrators and flight tests. The purpose of these projects was to develop propfan concepts that could fly efficiently at speeds comparable to high bypass ratio turbofans, i.e. close to Mach 0.8. General Electric proposed the UDF (UnDucted Fan), a pusher configuration with counter-rotating propellers driven by a counter-rotating low pressure turbine (GE36 Design and Systems Engineering, 1987). The 578-DX, a pusher configuration with counter-rotating 4 Will-be-set-by-IN-TECH

Fig. 1. Effect of fan tip pressure ratio and bypass duct pressure losses on fan equivalent

• Polymer composites and corresponding structural design and manufacturing techniques are studied in parallel with advances in metallic materials and manufacturing processes. • Shaft torque density capabilities through the development of metal matrix composites and

• Low pressure turbine weight savings through ultra high lift airfoil design, ultra high stage

• Technologies for light weight and low drag installation of high bypass ratio engines related

The open rotor engine concept, for high subsonic flight speeds, has also risen as a candidate for improving fuel consumption on several occasions since the advent of the first high bypass ratio turbofan engine. Such engine configurations, often refereed to as propfans in the literature, are direct competitors to ultra high bypass ratio turbofan engines. Their are located at the ultra-low specific thrust region of the design space, where propulsive efficiency benefits for turbofans are negated by very low transfer efficiencies. As illustrated in Fig. 1, this is due to the dominant effect on transfer efficiency that bypass duct pressure losses have when looking at low fan tip pressure ratio engine designs, i.e. low specific thrust. Open rotor engines do not suffer from bypass duct pressure losses and can therefore achieve a very high propulsive efficiency at a good level of transfer efficiency. Compared to turbofans, propfans also benefit

Several open rotor programs took place during the 80's, resulting in engine demonstrators and flight tests. The purpose of these projects was to develop propfan concepts that could fly efficiently at speeds comparable to high bypass ratio turbofans, i.e. close to Mach 0.8. General Electric proposed the UDF (UnDucted Fan), a pusher configuration with counter-rotating propellers driven by a counter-rotating low pressure turbine (GE36 Design and Systems Engineering, 1987). The 578-DX, a pusher configuration with counter-rotating

polytropic efficiency

multi metallic shafts.

loading, lightweight materials and design solutions.

to nozzle, nacelle and thrust reverser.

from reduced nacelle drag and weight penalties.

Fig. 2. Compressor efficiency improvement with year of entry into service.

propellers driven by a more conventional low pressure turbine through a differential planetary differential gearbox, was the result of a joint effort by Pratt & Whitney, Hamilton Standard, and Allison.

Both projects were eventually put on hold towards the end of the decade as fuel prices fell significantly. Nevertheless, the open rotor concept has now resurfaced within the EU Framework Program 7 research project DREAM (valiDation of Radical Engine Architecture systeMs, 2011) and the Clean Sky Joint Technology Initiative (2011). Within DREAM, the feasibility of two different open rotor architectures is evaluated including noise. Within Clean Sky, research work is being carried out by some of Europe's largest aero engine manufacturers, such as Rolls-Royce and Snecma, focused on designing, building and testing an open rotor demonstrator.

#### **3.2 Core technologies**

Improving core component efficiencies (including reducing losses in the cycle such as duct pressure losses) is one way of improving the engine thermal efficiency. Nevertheless, modern CFD-assisted designs are already quite aggressive and limited benefit may be envisaged by such future advancements (Kurzke, 2003); the increasing effort required to improve an already very good axial compressor design is illustrated in Fig.2.

Within the EU Framework Program 6 research project NEWAC (NEW Aero engine Core concepts, 2011) a number of advanced core component technologies have been investigated that include (Rolt & Kyprianidis, 2010; Wilfert et al., 2007):


Fig. 4. Evolution of turbine entry temperature and future trend.

specific thrust less marked.

intercooler are presented by Walker et al. (2009).

core's for a fixed OPR and specific thrust (Saravanamuttoo et al., 2001). The argument behind this is that the heat removed by the intercooler will largely need to be reintroduced in the combustor by burning more fuel, while the reduction in compression work and increase in bypass stream thrust (due to the heat rejection) will only partially compensate for the loss in cycle efficiency, at a fixed specific thrust and T4. Adding the expected intercooler pressure losses in the cycle calculations would further worsen the SFC deficit and make the increase in

Future Aero Engine Designs: An Evolving Vision 9

However, cycle calculations based on half-ideal gas properties and no dissociation (i.e. isobaric heat capacity dependent on temperature), presented by Walsh & Fletcher (1998), give a slightly different picture on intercooling. For a given T4, the optimal OPR for an intercooled core will be much higher than that for a conventional core. Comparing the two concepts at their optimal OPR levels, for a given technology level, can make the intercooled core more attractive with respect to thermal efficiency and not just specific thrust. Canière et al. (2006) and da Cunha Alves et al. (2001) also reached the same conclusion about the thermal efficiency of the intercooled cycle while studying this concept for gas turbines used in power generation. Papadopoulos & Pilidis (2000) worked on the introduction of intercooling, by means of heat pipes, in an aero engine design for long haul applications. Xu et al. (2007) performed a mission optimization to assess the potential of a tubular intercooler. Recent work by Xu & Grönstedt (2010) presents a refined tubular configuration estimating a potential block fuel benefit of 3.4%. The work addresses the limitation that short high pressure compressor blade lengths and related low compression efficiencies may impose on engines designed for short range missions, and suggest a novel gas path layout as a remedy to this constraint. A design study of a high OPR intercooled aero engine is described in Rolt & Baker (2009), while details on the aerodynamic challenges in designing a duct system to transfer the core air into and out of the

The introduction of recuperation in an aero engine, for high thermal efficiency at low OPR, has also been the focus of different researchers. Lundbladh & Sjunnesson (2003) performed a feasibility study for InterCooled (IC) and Intercooled Recuperated Aero engines (IRA) that

Fig. 3. Evolution of turbine material capability and future trend.

As mentioned earlier another way of improving engine thermal efficiency is to raise the cycle OPR. For conventional cores, increasing OPR and T4 depends primarily on future advancements in material and cooling technology. The evolution of turbine material capability over a period of 50 years is illustrated in Fig. 3. As can be observed, only mild improvements have been achieved so far and this seems to be a continuing trend; the potential introduction of ceramics would form a major improvement in the field, but substantially more research is still required before realising this. Despite the low improvement rate in turbine material technology (roughly 3 [K/year]) aero engine designs have seen substantial increases in T4 over the last 60 years (roughly 10 [K/year]); this is illustrated in Fig. 4 for engines designed for long-haul applications. The main reason behind these improvements in T4 has been the introduction of cooling and Thermal Barrier Coatings (TBC) in turbine designs; the interested reader is referred to Downs & Kenneth (2009) for a good overview of the evolution of turbine cooling systems design.

It is perhaps debatable whether an improvement rate of 10 [K/year] in T4 can be maintained in the future, and for that reason the design focus for more aggressive thermal efficiency improvements could very well be redirected to the introduction of heat-exchanged cores and advanced compressor technologies for future turbofan designs. In that respect, some of the technologies researched under the NEWAC project can be perceived as intermediate enabling steps for realising new engine core concepts that could improve the core thermal efficiency. These new core concepts comprise of:


When considering intercooling for an aero engine design, a common textbook misconception is that the thermal efficiency of an intercooled core will always be lower than a conventional 6 Will-be-set-by-IN-TECH

As mentioned earlier another way of improving engine thermal efficiency is to raise the cycle OPR. For conventional cores, increasing OPR and T4 depends primarily on future advancements in material and cooling technology. The evolution of turbine material capability over a period of 50 years is illustrated in Fig. 3. As can be observed, only mild improvements have been achieved so far and this seems to be a continuing trend; the potential introduction of ceramics would form a major improvement in the field, but substantially more research is still required before realising this. Despite the low improvement rate in turbine material technology (roughly 3 [K/year]) aero engine designs have seen substantial increases in T4 over the last 60 years (roughly 10 [K/year]); this is illustrated in Fig. 4 for engines designed for long-haul applications. The main reason behind these improvements in T4 has been the introduction of cooling and Thermal Barrier Coatings (TBC) in turbine designs; the interested reader is referred to Downs & Kenneth (2009) for a good overview of the evolution

It is perhaps debatable whether an improvement rate of 10 [K/year] in T4 can be maintained in the future, and for that reason the design focus for more aggressive thermal efficiency improvements could very well be redirected to the introduction of heat-exchanged cores and advanced compressor technologies for future turbofan designs. In that respect, some of the technologies researched under the NEWAC project can be perceived as intermediate enabling steps for realising new engine core concepts that could improve the core thermal efficiency.

When considering intercooling for an aero engine design, a common textbook misconception is that the thermal efficiency of an intercooled core will always be lower than a conventional

Fig. 3. Evolution of turbine material capability and future trend.

of turbine cooling systems design.

These new core concepts comprise of:

• High OPR flow controlled core.

• Ultra-high OPR core with intercooling. • Medium OPR intercooled recuperated core.

• High OPR active core including active cooling air cooling.

Fig. 4. Evolution of turbine entry temperature and future trend.

core's for a fixed OPR and specific thrust (Saravanamuttoo et al., 2001). The argument behind this is that the heat removed by the intercooler will largely need to be reintroduced in the combustor by burning more fuel, while the reduction in compression work and increase in bypass stream thrust (due to the heat rejection) will only partially compensate for the loss in cycle efficiency, at a fixed specific thrust and T4. Adding the expected intercooler pressure losses in the cycle calculations would further worsen the SFC deficit and make the increase in specific thrust less marked.

However, cycle calculations based on half-ideal gas properties and no dissociation (i.e. isobaric heat capacity dependent on temperature), presented by Walsh & Fletcher (1998), give a slightly different picture on intercooling. For a given T4, the optimal OPR for an intercooled core will be much higher than that for a conventional core. Comparing the two concepts at their optimal OPR levels, for a given technology level, can make the intercooled core more attractive with respect to thermal efficiency and not just specific thrust. Canière et al. (2006) and da Cunha Alves et al. (2001) also reached the same conclusion about the thermal efficiency of the intercooled cycle while studying this concept for gas turbines used in power generation. Papadopoulos & Pilidis (2000) worked on the introduction of intercooling, by means of heat pipes, in an aero engine design for long haul applications. Xu et al. (2007) performed a mission optimization to assess the potential of a tubular intercooler. Recent work by Xu & Grönstedt (2010) presents a refined tubular configuration estimating a potential block fuel benefit of 3.4%. The work addresses the limitation that short high pressure compressor blade lengths and related low compression efficiencies may impose on engines designed for short range missions, and suggest a novel gas path layout as a remedy to this constraint. A design study of a high OPR intercooled aero engine is described in Rolt & Baker (2009), while details on the aerodynamic challenges in designing a duct system to transfer the core air into and out of the intercooler are presented by Walker et al. (2009).

The introduction of recuperation in an aero engine, for high thermal efficiency at low OPR, has also been the focus of different researchers. Lundbladh & Sjunnesson (2003) performed a feasibility study for InterCooled (IC) and Intercooled Recuperated Aero engines (IRA) that

Fig. 5. Conceptual design tool algorithm (Kyprianidis, 2010).

• Cumulative EPNL vs. ICAO certification limits.

medium and long term goals.

• Engine time between overhaul.

aerodynamic) considerations.

• LTO (Landing and Take-Off) cycle D*p*NO*x*/F*oo* vs. ICAO certification limits and CAEP

Future Aero Engine Designs: An Evolving Vision 11

Where component design is concerned, for a conventional core the High Pressure Compressor (HPC) delivery temperature, and hence the engine OPR, is typically constrained by the mechanical properties of the HPC disc or HPC rear drive cone or High Pressure Turbine (HPT) disc material (Rolt & Baker, 2009). For an intercooled core, the OPR value is no longer constrained by a maximum allowable HPC delivery temperature. Nevertheless, the intercooling process increases the air density in the gas path and as a result the compressor blades tend to become smaller. Losses from tip clearances become increasingly important and a minimum compressor blade height limitation needs to be applied to maintain state of the art compressor efficiency. Core architecture selections for the conventional core set an upper limit to the HPC design pressure ratio that can achieved when driven by a single-stage HPT. A transonic single-stage HPT design can allow for relatively higher HPC pressure ratios at the expense of a lower polytropic efficiency. A two-stage HPT can offer high HPC pressure ratios at a high polytropic efficiency but a trade-off arises with respect to the need for more cooling air and increased engine length associated with the introduction of a second row of vanes and blades. With respect to the intercooled core, the minimum design pressure ratio for the Intermediate Pressure Compressor (IPC) can in some cases be limited by icing considerations during the descent flight phase. The maximum area variation that may be achieved by the variable area auxiliary nozzle is also constrained by mechanical (and

consider cycle benefits, weights and direct operating costs. Boggia & Rud (2005) provide an extended discussion on the thermodynamic cycle and the technological innovations necessary for realizing the intercooled recuperated core concept. Various aspects of the thermo-mechanical design of a compact heat exchanger have been presented by Pellischek & Kumpf (1991) and Schoenenborn et al. (2006). For a comprehensive review on the development activities for recuperated aero engines since the late 60's the interested reader can refer to McDonald et al. (2008a;b;c).

Finally, three different types of lean-burn combustor technology were also researched within NEWAC with the objective of reducing emissions of oxides of nitrogen (NO*x*):


#### **4. Design space exploration**

#### **4.1 Methodology, design feasibility and constraints**

To effectively explore the design space a tool is required that can consider the main disciplines typically encountered in conceptual design. The prediction of engine performance, aircraft design and performance, direct operating costs, and emissions for the concepts analysed in this study was made using the EVA code (Kyprianidis et al., 2008). Another code, WeiCo, was also used for carrying out mechanical and aerodynamic design in order to derive engine component weight and dimensions. The two tools have been integrated together within an optimiser environment as illustrated in Fig. 5, based on lessons learned from the development of the TERA2020 tool (Kyprianidis et al., 2011). This integration allows for multi-objective optimization, design studies, parametric studies, and sensitivity analysis. In order to speed up the execution of individual engine designs, the conceptual design tool attempts to minimize internal iterations in the calculation sequence through the use of an explicit algorithm, as described in detail by Kyprianidis (2010).

Aero-engine designs are subject to a large number of constraints and these need to be considered during conceptual design. Constraints can be applied within the optimiser environment at the end of the calculation sequence i.e., after the last design module has been executed. During a numerical optimisation, the optimiser will select a new set of input design parameters for every iteration and the resulting combination of aircraft and engine will be assessed. Using user specified objective functions the optimiser will home in on the best engine designs, determining the acceptability/feasibility of each design through the constraints set by the user. Infeasible designs will be ruled out, while non-optimum design values will result in engine designs with non-optimum values for the objective function selected. The optimiser will therefore avoid regions in the design pool that result in infeasible or non-optimum engine designs.

Design constraints set by the user include among others:


8 Will-be-set-by-IN-TECH

consider cycle benefits, weights and direct operating costs. Boggia & Rud (2005) provide an extended discussion on the thermodynamic cycle and the technological innovations necessary for realizing the intercooled recuperated core concept. Various aspects of the thermo-mechanical design of a compact heat exchanger have been presented by Pellischek & Kumpf (1991) and Schoenenborn et al. (2006). For a comprehensive review on the development activities for recuperated aero engines since the late 60's the interested reader

Finally, three different types of lean-burn combustor technology were also researched within

To effectively explore the design space a tool is required that can consider the main disciplines typically encountered in conceptual design. The prediction of engine performance, aircraft design and performance, direct operating costs, and emissions for the concepts analysed in this study was made using the EVA code (Kyprianidis et al., 2008). Another code, WeiCo, was also used for carrying out mechanical and aerodynamic design in order to derive engine component weight and dimensions. The two tools have been integrated together within an optimiser environment as illustrated in Fig. 5, based on lessons learned from the development of the TERA2020 tool (Kyprianidis et al., 2011). This integration allows for multi-objective optimization, design studies, parametric studies, and sensitivity analysis. In order to speed up the execution of individual engine designs, the conceptual design tool attempts to minimize internal iterations in the calculation sequence through the use of an explicit algorithm, as

Aero-engine designs are subject to a large number of constraints and these need to be considered during conceptual design. Constraints can be applied within the optimiser environment at the end of the calculation sequence i.e., after the last design module has been executed. During a numerical optimisation, the optimiser will select a new set of input design parameters for every iteration and the resulting combination of aircraft and engine will be assessed. Using user specified objective functions the optimiser will home in on the best engine designs, determining the acceptability/feasibility of each design through the constraints set by the user. Infeasible designs will be ruled out, while non-optimum design values will result in engine designs with non-optimum values for the objective function selected. The optimiser will therefore avoid regions in the design pool that result in infeasible

• Take-off HPC delivery temperature and other important performance parameters.

• FAR (Federal Aviation Regulations) take-off field length for all engines operating and

NEWAC with the objective of reducing emissions of oxides of nitrogen (NO*x*): • Lean Direct Injection (LDI) combustor for high and ultra-high OPR cores. • Partial Evaporation and Rapid Mixing (PERM) combustor for high OPR cores.

• Lean Premixed Pre-vaporized (LPP) combustor for medium OPR cores.

can refer to McDonald et al. (2008a;b;c).

**4. Design space exploration**

**4.1 Methodology, design feasibility and constraints**

described in detail by Kyprianidis (2010).

or non-optimum engine designs.

• Time to height.

Design constraints set by the user include among others:

balanced field length for one engine inoperative conditions.

Fig. 5. Conceptual design tool algorithm (Kyprianidis, 2010).


Where component design is concerned, for a conventional core the High Pressure Compressor (HPC) delivery temperature, and hence the engine OPR, is typically constrained by the mechanical properties of the HPC disc or HPC rear drive cone or High Pressure Turbine (HPT) disc material (Rolt & Baker, 2009). For an intercooled core, the OPR value is no longer constrained by a maximum allowable HPC delivery temperature. Nevertheless, the intercooling process increases the air density in the gas path and as a result the compressor blades tend to become smaller. Losses from tip clearances become increasingly important and a minimum compressor blade height limitation needs to be applied to maintain state of the art compressor efficiency. Core architecture selections for the conventional core set an upper limit to the HPC design pressure ratio that can achieved when driven by a single-stage HPT. A transonic single-stage HPT design can allow for relatively higher HPC pressure ratios at the expense of a lower polytropic efficiency. A two-stage HPT can offer high HPC pressure ratios at a high polytropic efficiency but a trade-off arises with respect to the need for more cooling air and increased engine length associated with the introduction of a second row of vanes and blades. With respect to the intercooled core, the minimum design pressure ratio for the Intermediate Pressure Compressor (IPC) can in some cases be limited by icing considerations during the descent flight phase. The maximum area variation that may be achieved by the variable area auxiliary nozzle is also constrained by mechanical (and aerodynamic) considerations.

Fig. 6. Typical flight cycle.

and Joint Aviation Authorities (2008).

models and are considered reasonable numbers.

the take-off and time to height criteria set.

Airbus A330-200 (Swiss International Air Lines, 2009). This does not necessarily mean that the model's business case is not a realistic one; nor that it wouldn't fit well with operational practices followed by other airlines. Furthermore, regional Air Traffic Management (ATM) practices can skew available block fuel data, while global ATM regulations may very well change significantly by 2020. It should be noted that fuel planning within the model respects the requirements defined for international flights by Federal Aviation Administration (n.d.)

Future Aero Engine Designs: An Evolving Vision 13

Where conceptual design is concerned, exchange rates are perhaps a better type of parameter for evaluating the accuracy of a rubberised wing model, rather than just simply comparing absolute values. Block fuel exchange rates produced with the rubberised wing baseline aircraft models are presented in Table 1 for the business case of the long and short range

During a block fuel optimization all engine aircraft combinations which do not fulfil the take-off and time to height criteria set will be discarded as infeasible. Due to the underlying physics, this will naturally lead to an optimal engine and aircraft combination for the defined objective function. All large engines will produce heavier aircraft with more drag and thus higher block fuel weight. Engines which are too small will not deliver enough thrust to satisfy

As discussed earlier, designing a combustor at very low air to fuel ratio levels is also limited by the need for adequate combustor liner film-cooling air as well as maintaining an acceptable temperature traverse quality (Lefebvre, 1999); this sets an upper bound on combustor outlet temperature. Furthermore, a maximum permissible mean metal temperature needs to be set to consider turbine blade material limitations. A lower bound on engine time between overhaul also needs to be set to limit the frequency of workshop visits. For short range applications the minimum engine time between overhaul was set to 18000 [hr] while for long range applications to 23000 [hr]. This reflects the fact that designs for short range applications are typically operated at high power conditions for a significantly larger part of their operational life. Significantly lower levels of maximum combustor outlet temperature and turbine blade mean metal temperature had to be selected, compared to what could be selected for engine designs for long range applications that are often operated at derated thrust levels and spend most of their life at cruise.

A rubberised aircraft wing model was used in these studies to capture 'snowball effects' with respect to maximum take-off weight variation, rather than using fixed engine thrust requirements. The aircraft drag polar and weight breakdown were predicted at component level from the aircraft geometry and high lift device settings for the take-off and approach phases. Fuel burned was calculated for the entire flight mission including reserves assuming ISA conditions, as illustrated in Fig. 6. Cruise is performed at the optimum altitude for specific range (fixed cruise Mach number) using a step-up cruise procedure as the aircraft gets lighter. A comprehensive take-off field length calculation is performed for all engines operating and one engine inoperative conditions up to 1500 [ft].

Two baseline aircraft models have been used herein; one model for long range applications and one for short range. The former model is largely based on public domain information available for the Airbus A330-200 while the latter model is based on the Airbus A320-200. The short range aircraft was designed to carry 150 [pax] for a distance of 3000 [nmi] and a typical business case of 500 [nmi]; for long range applications it was designed for 253 [pax], 6750 [nmi] and 3000 [nm], respectively. For the step-up cruise procedure, a minimum residual rate of climb of 300 [ft/min] was set as a constraint for flying at the cruise altitude for maximum specific range.

The maximum values for FAR take-off field length and time to height were set for a load factor of 1 and no cargo. The choice of both is based on customer operational requirements as the aircraft needs to be able to: (i) take-off from a large number of airports around the world and (ii) climb to the initial cruise altitude sufficiently fast to ease operations with local air traffic control (and hence reduce waiting time on the ground). A cumulative distribution of the world's major runway lengths, based on data from Jenkinson et al. (1999), is illustrated in Fig. 7. For short range applications fairly stringent constraints are typically set for the maximum take-off distance and time to height; in this study these were set to 2.0 [km] and 25 [min], respectively. For long range applications a maximum take-off distance of 2.5 [km] was set instead. Stringent constraints result in bigger engines but allow for greater flexibility for engine derating at a smaller block fuel cost.

The choice of load factor and cargo is considered sensible but it does not necessarily constitute a typical airline practice. Validating absolute block fuel predictions with public domain airline data is not a trivial task as different airlines will follow different operational practices. For example for the long range aircraft model, the business case prediction is 10% lower than the published annually-averaged value, given in [lt/(km\*pax)], by SwissAir for 2009 for the 10 Will-be-set-by-IN-TECH

As discussed earlier, designing a combustor at very low air to fuel ratio levels is also limited by the need for adequate combustor liner film-cooling air as well as maintaining an acceptable temperature traverse quality (Lefebvre, 1999); this sets an upper bound on combustor outlet temperature. Furthermore, a maximum permissible mean metal temperature needs to be set to consider turbine blade material limitations. A lower bound on engine time between overhaul also needs to be set to limit the frequency of workshop visits. For short range applications the minimum engine time between overhaul was set to 18000 [hr] while for long range applications to 23000 [hr]. This reflects the fact that designs for short range applications are typically operated at high power conditions for a significantly larger part of their operational life. Significantly lower levels of maximum combustor outlet temperature and turbine blade mean metal temperature had to be selected, compared to what could be selected for engine designs for long range applications that are often operated at derated thrust levels and spend

A rubberised aircraft wing model was used in these studies to capture 'snowball effects' with respect to maximum take-off weight variation, rather than using fixed engine thrust requirements. The aircraft drag polar and weight breakdown were predicted at component level from the aircraft geometry and high lift device settings for the take-off and approach phases. Fuel burned was calculated for the entire flight mission including reserves assuming ISA conditions, as illustrated in Fig. 6. Cruise is performed at the optimum altitude for specific range (fixed cruise Mach number) using a step-up cruise procedure as the aircraft gets lighter. A comprehensive take-off field length calculation is performed for all engines operating and

Two baseline aircraft models have been used herein; one model for long range applications and one for short range. The former model is largely based on public domain information available for the Airbus A330-200 while the latter model is based on the Airbus A320-200. The short range aircraft was designed to carry 150 [pax] for a distance of 3000 [nmi] and a typical business case of 500 [nmi]; for long range applications it was designed for 253 [pax], 6750 [nmi] and 3000 [nm], respectively. For the step-up cruise procedure, a minimum residual rate of climb of 300 [ft/min] was set as a constraint for flying at the cruise altitude for maximum

The maximum values for FAR take-off field length and time to height were set for a load factor of 1 and no cargo. The choice of both is based on customer operational requirements as the aircraft needs to be able to: (i) take-off from a large number of airports around the world and (ii) climb to the initial cruise altitude sufficiently fast to ease operations with local air traffic control (and hence reduce waiting time on the ground). A cumulative distribution of the world's major runway lengths, based on data from Jenkinson et al. (1999), is illustrated in Fig. 7. For short range applications fairly stringent constraints are typically set for the maximum take-off distance and time to height; in this study these were set to 2.0 [km] and 25 [min], respectively. For long range applications a maximum take-off distance of 2.5 [km] was set instead. Stringent constraints result in bigger engines but allow for greater flexibility for

The choice of load factor and cargo is considered sensible but it does not necessarily constitute a typical airline practice. Validating absolute block fuel predictions with public domain airline data is not a trivial task as different airlines will follow different operational practices. For example for the long range aircraft model, the business case prediction is 10% lower than the published annually-averaged value, given in [lt/(km\*pax)], by SwissAir for 2009 for the

most of their life at cruise.

specific range.

one engine inoperative conditions up to 1500 [ft].

engine derating at a smaller block fuel cost.

Fig. 6. Typical flight cycle.

Airbus A330-200 (Swiss International Air Lines, 2009). This does not necessarily mean that the model's business case is not a realistic one; nor that it wouldn't fit well with operational practices followed by other airlines. Furthermore, regional Air Traffic Management (ATM) practices can skew available block fuel data, while global ATM regulations may very well change significantly by 2020. It should be noted that fuel planning within the model respects the requirements defined for international flights by Federal Aviation Administration (n.d.) and Joint Aviation Authorities (2008).

Where conceptual design is concerned, exchange rates are perhaps a better type of parameter for evaluating the accuracy of a rubberised wing model, rather than just simply comparing absolute values. Block fuel exchange rates produced with the rubberised wing baseline aircraft models are presented in Table 1 for the business case of the long and short range models and are considered reasonable numbers.

During a block fuel optimization all engine aircraft combinations which do not fulfil the take-off and time to height criteria set will be discarded as infeasible. Due to the underlying physics, this will naturally lead to an optimal engine and aircraft combination for the defined objective function. All large engines will produce heavier aircraft with more drag and thus higher block fuel weight. Engines which are too small will not deliver enough thrust to satisfy the take-off and time to height criteria set.

Fig. 8. Visualization example of constrained design space exploration.

the late 30's is summarised in Fig. 9 based on data from Avellán (2008).

exchange rate, as well as interest and inflation rates.

the US\$ to Euro exchange rate was assumed to be 0.8222.

Safety considerations aside, civil aero engine design has been driven primarily by economic considerations even from its fairly early days. A testament to this has been the advent of the world's first commercial jet-airliner, the de Havilland Comet, powered by 4 Rolls-Royce Avon turbojet engines. Although it burned nearly four times as much fuel compared to piston-driven engines, it's business case was very strong since it permitted significantly higher flight speeds resulting in reduced flight times (i.e. a better airline product) and increased aircraft annual utilization. Furthermore, the excellent power to weight ratio of the turbojet engine meant that it could be used to power aircrafts with significantly higher passenger capacities than what was feasible before. The evolution of aircraft transport efficiency since

Future Aero Engine Designs: An Evolving Vision 15

The aero engine designs proposed herein have been optimized for minimum block fuel for a given aircraft mission (business case), which implies minimum global warming impact if one considers CO2 emissions alone. The market competitiveness of these fuel optimal designs however is highly dependent on the development of jet fuel prices in the years to come until 2020. The volatility of jet fuel price over the last 10 years is illustrated in Fig. 10. A further economic consideration for European markets may also be the development of the Euro/US\$

For the economic calculations conducted in this study certain assumptions were made. The assumed jet fuel price was 172c\$/US gallon. It is worth noting that at the time of writing the average jet fuel price was 320 [c\$/US gallon] (International Air Transport Association, 2011; Platts, 2011). Interest and inflation rates were assumed to be 6% and 2%, respectively, while

It is worth noting that an increase in inflation rates from 2% to 3% can increase the net present cost by as much as 17%, over a period of 30 years. An increase in interest rates from 6% to 7% can increase Direct Operating Costs (DOC) by 2.5% and 4.5% for short and long range

**4.3 Economic considerations**

applications, respectively.

Fig. 7. Cumulative distribution of world's major runway lengths (based on data from Jenkinson et al. (1999)).


Table 1. Block fuel exchange rates using the baseline long range and short range rubberised wing aircraft models.

#### **4.2 Engine design optimality**

Whereas optimisation constraints can help ensure the feasibility of an engine design, they do little to help with it's optimality. The optimality of the engine design will depend on the careful selection of the figures of merit used during the optimisation process, such as minimum block fuel, maximum time between overhaul, minimum direct operating costs, minimum noise and LTO NO*<sup>x</sup>* emissions etc.

Determining the optimal aero-engine design is essentially the subject of a multi-objective optimisation, and therefore Pareto fronts need typically be constructed to visualize the region of optimal designs within the design space. A simplified example of utilizing the tool for design space exploration, with active constraints, is illustrated in Fig. 8. In principle, nacelle drag should also be added as a third dimension when plotting design space exploration results that consider varying levels of specific thrust, but this has been omitted here in order to simplify the plot. The aircraft exchange rates for the baseline design were used for plotting a constant block fuel line (ignoring nacelle drag effects and nonlinearities) and this iso-line therefore defines, in a simple manner, the boundaries of trading specific fuel consumption for weight. During a block fuel optimization, the optimizer continuously evaluates different engine designs as it searches for the optimal solution. Designs that fail to meet constraints set by the user are discarded and have been labeled as infeasible in the plot.

12 Will-be-set-by-IN-TECH

Fig. 7. Cumulative distribution of world's major runway lengths (based on data

Perturbation Long range Short range 1000 [kg] weight penalty 0.73% 1.26% +1% SFC 1.28% 1.09%

Table 1. Block fuel exchange rates using the baseline long range and short range rubberised

Whereas optimisation constraints can help ensure the feasibility of an engine design, they do little to help with it's optimality. The optimality of the engine design will depend on the careful selection of the figures of merit used during the optimisation process, such as minimum block fuel, maximum time between overhaul, minimum direct operating costs,

Determining the optimal aero-engine design is essentially the subject of a multi-objective optimisation, and therefore Pareto fronts need typically be constructed to visualize the region of optimal designs within the design space. A simplified example of utilizing the tool for design space exploration, with active constraints, is illustrated in Fig. 8. In principle, nacelle drag should also be added as a third dimension when plotting design space exploration results that consider varying levels of specific thrust, but this has been omitted here in order to simplify the plot. The aircraft exchange rates for the baseline design were used for plotting a constant block fuel line (ignoring nacelle drag effects and nonlinearities) and this iso-line therefore defines, in a simple manner, the boundaries of trading specific fuel consumption for weight. During a block fuel optimization, the optimizer continuously evaluates different engine designs as it searches for the optimal solution. Designs that fail to meet constraints set

by the user are discarded and have been labeled as infeasible in the plot.

Exchange rate

from Jenkinson et al. (1999)).

wing aircraft models.

**4.2 Engine design optimality**

minimum noise and LTO NO*<sup>x</sup>* emissions etc.

Fig. 8. Visualization example of constrained design space exploration.

#### **4.3 Economic considerations**

Safety considerations aside, civil aero engine design has been driven primarily by economic considerations even from its fairly early days. A testament to this has been the advent of the world's first commercial jet-airliner, the de Havilland Comet, powered by 4 Rolls-Royce Avon turbojet engines. Although it burned nearly four times as much fuel compared to piston-driven engines, it's business case was very strong since it permitted significantly higher flight speeds resulting in reduced flight times (i.e. a better airline product) and increased aircraft annual utilization. Furthermore, the excellent power to weight ratio of the turbojet engine meant that it could be used to power aircrafts with significantly higher passenger capacities than what was feasible before. The evolution of aircraft transport efficiency since the late 30's is summarised in Fig. 9 based on data from Avellán (2008).

The aero engine designs proposed herein have been optimized for minimum block fuel for a given aircraft mission (business case), which implies minimum global warming impact if one considers CO2 emissions alone. The market competitiveness of these fuel optimal designs however is highly dependent on the development of jet fuel prices in the years to come until 2020. The volatility of jet fuel price over the last 10 years is illustrated in Fig. 10. A further economic consideration for European markets may also be the development of the Euro/US\$ exchange rate, as well as interest and inflation rates.

For the economic calculations conducted in this study certain assumptions were made. The assumed jet fuel price was 172c\$/US gallon. It is worth noting that at the time of writing the average jet fuel price was 320 [c\$/US gallon] (International Air Transport Association, 2011; Platts, 2011). Interest and inflation rates were assumed to be 6% and 2%, respectively, while the US\$ to Euro exchange rate was assumed to be 0.8222.

It is worth noting that an increase in inflation rates from 2% to 3% can increase the net present cost by as much as 17%, over a period of 30 years. An increase in interest rates from 6% to 7% can increase Direct Operating Costs (DOC) by 2.5% and 4.5% for short and long range applications, respectively.

optimal designs. It would therefore be worthwhile to redirect further research investments towards developing fuel efficient aero engine designs, as has also been the case in the late 70's and through large part of the 80's. The introduction of carbon taxes could also have a similar

Future Aero Engine Designs: An Evolving Vision 17

A summary of three different design space exploration case studies using the tools and algorithm described is presented here. This work has looked at the potential block fuel

The thrust requirements for the first two concepts are for an engine designed to power the long range aircraft model while the latter concept is centered around powering the short range aircraft model. More details on these studies can be found in Kyprianidis et al. (2011) and

For the intercooled core assessment, a year 2020 Entry Into Service (EIS) turbofan engine with a conventional core was set up as the baseline. The intercooled core engine is an ultra high OPR design with also year 2020 EIS level of technology, and features a tubular heat-exchanger, while the fan for both engines has the same diameter and flow per unit of area. Business case block fuel benefits of approximately 3.2% are predicted for the intercooled engine, mainly due to the reduced engine weight and the core's higher thermal efficiency which results in a better SFC. **These intercooling benefits are highly dependent on achieving technology targets such as low intercooler weight and pressure losses**; the predicted lower dry weight, compared to the conventional core engine, can be attributed to various reasons. The intercooler weight penalty is largely compensated by the higher core specific output allowing a smaller core size and hence a higher BPR at a fixed thrust and fan diameter. The high OPR provides an additional sizing benefit, for components downstream of the HPC, by reducing further the corrected mass flow and hence flow areas. The intercooled core Low Pressure Turbine (LPT) was designed in this study with one less stage which reduced both engine weight and length, despite the high cycle OPR requiring a greater number of HPC stages. These observations are summarised in Table 2 with the added components weight group considering the intercooler and its installation standard; this group is not considered in the core weight group which also does not consider the core nozzle or the LPT and its casing. For the intercooled recuperated core assessment, a year 2000 EIS turbofan engine with a conventional core was set up as the baseline. The intercooled recuperated core configuration is an UHBR design with a year 2020 level of technology. Significant business case block fuel benefits of nearly 22% are predicted for the geared intercooled recuperated core engine due to its higher thermal and propulsive efficiency. The use of HPT cooling air bled from the recuperator exit (Boggia & Rud, 2005; Walsh & Fletcher, 1998) results in a 1.3% SFC improvement due to more energy being recuperated from the exhausts, at a fixed effectiveness level - and despite the considerable increase in cooling air requirements (+3.5% of core mass flow). The predicted dry weight for the intercooled recuperated configuration is higher compared to the conventional core engine. There is a weight benefit from the use of EIS

**5. Summary of design space exploration results**

• An intercooled core in a direct drive UHBR turbofan configuration.

• An open rotor propulsor in a geared pusher configuration.

• An intercooled recuperated core in a geared UHBR turbofan configuration.

benefits resulting from the introduction of:

effect.

Larsson et al. (2011).

Fig. 9. Evolution of aircraft transport efficiency (based on data from Avellán (2008)).

Fig. 10. Long term perspective of jet fuel price movements (based on data from International Air Transport Association (2011) and Platts (2011)).

The cost of fuel as a fraction of the total DOC was predicted to be 13% and 19% for short and long range applications, respectively. An increase in block fuel by 1% translates in an increase of 0.13% and 0.19% in DOC, respectively, and as can be observed it is directly dependent on the ratio of fuel cost over DOC. A doubling of the fuel price would change this ratio to roughly 23% and 32%, respectively, and would also result in 13% and 19% higher DOC levels, respectively.

Higher levels of DOC, as a result of a significant increase in fuel price, would most probably be absorbed by airlines through an increase in fares. This could make fuel efficient designs increasingly market competitive, as the DOC optimal designs would further approach the fuel 14 Will-be-set-by-IN-TECH

Fig. 9. Evolution of aircraft transport efficiency (based on data from Avellán (2008)).

Fig. 10. Long term perspective of jet fuel price movements (based on data from International

The cost of fuel as a fraction of the total DOC was predicted to be 13% and 19% for short and long range applications, respectively. An increase in block fuel by 1% translates in an increase of 0.13% and 0.19% in DOC, respectively, and as can be observed it is directly dependent on the ratio of fuel cost over DOC. A doubling of the fuel price would change this ratio to roughly 23% and 32%, respectively, and would also result in 13% and 19% higher DOC levels,

Higher levels of DOC, as a result of a significant increase in fuel price, would most probably be absorbed by airlines through an increase in fares. This could make fuel efficient designs increasingly market competitive, as the DOC optimal designs would further approach the fuel

Air Transport Association (2011) and Platts (2011)).

respectively.

optimal designs. It would therefore be worthwhile to redirect further research investments towards developing fuel efficient aero engine designs, as has also been the case in the late 70's and through large part of the 80's. The introduction of carbon taxes could also have a similar effect.

#### **5. Summary of design space exploration results**

A summary of three different design space exploration case studies using the tools and algorithm described is presented here. This work has looked at the potential block fuel benefits resulting from the introduction of:


The thrust requirements for the first two concepts are for an engine designed to power the long range aircraft model while the latter concept is centered around powering the short range aircraft model. More details on these studies can be found in Kyprianidis et al. (2011) and Larsson et al. (2011).

For the intercooled core assessment, a year 2020 Entry Into Service (EIS) turbofan engine with a conventional core was set up as the baseline. The intercooled core engine is an ultra high OPR design with also year 2020 EIS level of technology, and features a tubular heat-exchanger, while the fan for both engines has the same diameter and flow per unit of area. Business case block fuel benefits of approximately 3.2% are predicted for the intercooled engine, mainly due to the reduced engine weight and the core's higher thermal efficiency which results in a better SFC. **These intercooling benefits are highly dependent on achieving technology targets such as low intercooler weight and pressure losses**; the predicted lower dry weight, compared to the conventional core engine, can be attributed to various reasons. The intercooler weight penalty is largely compensated by the higher core specific output allowing a smaller core size and hence a higher BPR at a fixed thrust and fan diameter. The high OPR provides an additional sizing benefit, for components downstream of the HPC, by reducing further the corrected mass flow and hence flow areas. The intercooled core Low Pressure Turbine (LPT) was designed in this study with one less stage which reduced both engine weight and length, despite the high cycle OPR requiring a greater number of HPC stages. These observations are summarised in Table 2 with the added components weight group considering the intercooler and its installation standard; this group is not considered in the core weight group which also does not consider the core nozzle or the LPT and its casing. For the intercooled recuperated core assessment, a year 2000 EIS turbofan engine with a conventional core was set up as the baseline. The intercooled recuperated core configuration is an UHBR design with a year 2020 level of technology. Significant business case block fuel benefits of nearly 22% are predicted for the geared intercooled recuperated core engine due to its higher thermal and propulsive efficiency. The use of HPT cooling air bled from the recuperator exit (Boggia & Rud, 2005; Walsh & Fletcher, 1998) results in a 1.3% SFC improvement due to more energy being recuperated from the exhausts, at a fixed effectiveness level - and despite the considerable increase in cooling air requirements (+3.5% of core mass flow). The predicted dry weight for the intercooled recuperated configuration is higher compared to the conventional core engine. There is a weight benefit from the use of EIS

Geared turbofan Geared open rotor GTF SR GOR SR EIS 2020 EIS 2020

Engine installed weight Ref. +11% Nacelle weight Ref. -88% Fan/propeller weight Ref. +73% LPT weight Ref. +20% Core weight Ref. -31% Block fuel weight Ref. -15% Mid-cruise SFC Ref. -14% Thermal efficiency Ref. -0.013 Propulsive efficiency Ref. +0.16

Future Aero Engine Designs: An Evolving Vision 19

Table 4. Comparison of a geared open rotor engine with a geared turbofan engine at aircraft

block fuel benefits of nearly 15% are predicted for the geared open rotor engine primarily due to its higher propulsive efficiency. Although, the geared turbofan engine benefits from a better thrust to weight ratio it suffers from significantly higher nacelle drag losses, compared to the

A NO*<sup>x</sup>* emissions assessment of the presented engine configurations has been performed and is illustrated in Fig. 11. The same combustor concept has been considered for both designs i.e., conventional Rich-burn/Quick-quench/Lean-burn (RQL) combustion technology. The results obtained are compared against ICAO Annex 16 Volume II legislative limits (ICAO, 1993), as well as the Medium Term (MT) and Long Term (LT) technology goals set by CAEP (Newton et al., 2007). Balloons have been used to indicate the uncertainty in the NO*<sup>x</sup>* predictions due to the lower technology readiness level associated with the introduction of

Fig. 11. NO*<sup>x</sup>* emissions assessment for different future aero engine design concepts.

open rotor design. These observations are summarised in Table 3.

system level.


Table 2. Comparison of an intercooled engine with a conventional core turbofan engine at aircraft system level.


Table 3. Comparison of an intercooled recuperated engine with a conventional core turbofan engine at aircraft system level.

2020 light-weight materials in most major engine components, as well as from the high speed LPT - due to the reduced stage count. Also, the relatively low engine OPR and the use of an intercooler increases core specific output, resulting in a smaller core. The introduction however of the gearbox, intercooler and recuperator components inevitably results in a significant weight penalty. It should be noted that a lower level of specific thrust, and hence a larger fan diameter, has been assumed for the intercooled recuperated core engine; this results in both a heavier fan and a heavier nacelle. These observations are summarised in Table 3 with the added components weight group considering the intercooler and recuperator and their installation standard, as well as the gearbox.

For the geared open rotor assessment, a year 2020 EIS geared turbofan engine with a conventional core was set up as the baseline. The geared open rotor concept design also assumes year 2020 EIS level of technology, and features two counter-rotating propellers in a pusher configuration powered by a geared low pressure turbine. Significant business case 16 Will-be-set-by-IN-TECH

Engine dry weight Ref. -5.9% LPT weight Ref. -27.1% Core weight Ref. -32.5% Added components weight - 7.7%

Block fuel weight Ref. -3.2% Mid-cruise SFC Ref. -1.5% Thermal efficiency Ref. +0.007 Propulsive efficiency Ref. +0.000 Table 2. Comparison of an intercooled engine with a conventional core turbofan engine at

Thrust/weight Ref. -12% Engine dry weight Ref. +16.5% Nacelle weight Ref. +29.7% Fan weight Ref. +36.6% LPT weight Ref. -17.1% Added components weight - 25.4%

Block fuel weight Ref. -21.6% Mid-cruise SFC Ref. -18.3% Thermal efficiency Ref. +0.024 Propulsive efficiency Ref. +0.120

Table 3. Comparison of an intercooled recuperated engine with a conventional core turbofan

2020 light-weight materials in most major engine components, as well as from the high speed LPT - due to the reduced stage count. Also, the relatively low engine OPR and the use of an intercooler increases core specific output, resulting in a smaller core. The introduction however of the gearbox, intercooler and recuperator components inevitably results in a significant weight penalty. It should be noted that a lower level of specific thrust, and hence a larger fan diameter, has been assumed for the intercooled recuperated core engine; this results in both a heavier fan and a heavier nacelle. These observations are summarised in Table 3 with the added components weight group considering the intercooler and recuperator and

For the geared open rotor assessment, a year 2020 EIS geared turbofan engine with a conventional core was set up as the baseline. The geared open rotor concept design also assumes year 2020 EIS level of technology, and features two counter-rotating propellers in a pusher configuration powered by a geared low pressure turbine. Significant business case

(as % of engine dry weight)

(as % of engine dry weight)

their installation standard, as well as the gearbox.

engine at aircraft system level.

aircraft system level.

Conventional core Intercooled core DDTF LR DDIC LR EIS 2020 EIS 2020

Conventional core Intercooled recuperated core BASE LR IRA LR EIS 2000 EIS 2020


Table 4. Comparison of a geared open rotor engine with a geared turbofan engine at aircraft system level.

block fuel benefits of nearly 15% are predicted for the geared open rotor engine primarily due to its higher propulsive efficiency. Although, the geared turbofan engine benefits from a better thrust to weight ratio it suffers from significantly higher nacelle drag losses, compared to the open rotor design. These observations are summarised in Table 3.

A NO*<sup>x</sup>* emissions assessment of the presented engine configurations has been performed and is illustrated in Fig. 11. The same combustor concept has been considered for both designs i.e., conventional Rich-burn/Quick-quench/Lean-burn (RQL) combustion technology. The results obtained are compared against ICAO Annex 16 Volume II legislative limits (ICAO, 1993), as well as the Medium Term (MT) and Long Term (LT) technology goals set by CAEP (Newton et al., 2007). Balloons have been used to indicate the uncertainty in the NO*<sup>x</sup>* predictions due to the lower technology readiness level associated with the introduction of

Fig. 11. NO*<sup>x</sup>* emissions assessment for different future aero engine design concepts.

**9. References**

Advisory Council for Aeronautical Research in Europe (2001). European Aeronautics: A

Future Aero Engine Designs: An Evolving Vision 21

Avellán, R. (2008). *Towards Environmentally Friendly Aero Engines*, Licentiate thesis, Chalmers

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such combustor designs in the proposed future cycles. A sufficient margin against the ICAO CAEP/6 LTO cycle NO*<sup>x</sup>* certification limit may be achieved for all the configurations that have been assessed assuming year 2020 EIS.

## **6. Conclusions**

The research work presented started by reviewing the evolution of the aero engine industry's vision for the aero engine design of the future. Appropriate research questions were set that can influence how this vision may further involve in the years to come. Design constraints, material technology, customer requirements, noise and emissions legislation, technology risk and economic considerations and their effect on optimal concept selection were also discussed in detail.

With respect to addressing these questions, several novel engine cycles and technologies currently under research - were identified. It was shown that there is a great potential to reduce fuel consumption for the different concepts identified, and consequently decrease the CO2 emissions. Furthermore, this can be achieved with a sufficient margin from the ICAO NO*<sup>x</sup>* certification limits, and in line with the medium term and long term goals set by CAEP. It must be noted however that aero engine design is primarily driven by economic considerations. As fuel prices increase, the impact of fuel consumption on direct operating costs also increases. The question therefore rises:

#### **Can the potential reduction in fuel consumption and direct operating costs outweigh the technological risks involved in introducing novel concepts into the market?**

The answer is left to be given by the choices the aero engine industry makes in the years to come.

## **7. Acknowledgements**

The author is grateful to Richard Avellán (Volvo Aero) for providing the transport efficiency data used in Fig. 9. Stimulating discussions with A.M. Rolt (Rolls-Royce), J.A. Borradaile, S. Donnerhack (MTU Aero Engines), P. Pilidis, (Cranfield University), R. Singh, (Cranfield University), S.O.T. Ogaji, (Cranfield University), P. Giannakakis (Cranfield University), T. Grönstedt (Chalmers University), A. Lundbladh (Volvo Aero) and L. Larsson (Volvo Aero) on advanced concepts and aero engine design are gratefully acknowledged. Finally, the author would like to thank the reviewers of this work for their constructive suggestions to improve the overall quality and clarity of the article.

## **8. Nomenclature**


T4 Combustor outlet temperature

#### **9. References**

18 Will-be-set-by-IN-TECH

such combustor designs in the proposed future cycles. A sufficient margin against the ICAO CAEP/6 LTO cycle NO*<sup>x</sup>* certification limit may be achieved for all the configurations that have

The research work presented started by reviewing the evolution of the aero engine industry's vision for the aero engine design of the future. Appropriate research questions were set that can influence how this vision may further involve in the years to come. Design constraints, material technology, customer requirements, noise and emissions legislation, technology risk and economic considerations and their effect on optimal concept selection were also discussed

With respect to addressing these questions, several novel engine cycles and technologies currently under research - were identified. It was shown that there is a great potential to reduce fuel consumption for the different concepts identified, and consequently decrease the CO2 emissions. Furthermore, this can be achieved with a sufficient margin from the ICAO NO*<sup>x</sup>* certification limits, and in line with the medium term and long term goals set by CAEP. It must be noted however that aero engine design is primarily driven by economic considerations. As fuel prices increase, the impact of fuel consumption on direct operating

**Can the potential reduction in fuel consumption and direct operating costs outweigh the technological risks involved in introducing novel concepts into the**

The answer is left to be given by the choices the aero engine industry makes in the years to

The author is grateful to Richard Avellán (Volvo Aero) for providing the transport efficiency data used in Fig. 9. Stimulating discussions with A.M. Rolt (Rolls-Royce), J.A. Borradaile, S. Donnerhack (MTU Aero Engines), P. Pilidis, (Cranfield University), R. Singh, (Cranfield University), S.O.T. Ogaji, (Cranfield University), P. Giannakakis (Cranfield University), T. Grönstedt (Chalmers University), A. Lundbladh (Volvo Aero) and L. Larsson (Volvo Aero) on advanced concepts and aero engine design are gratefully acknowledged. Finally, the author would like to thank the reviewers of this work for their constructive suggestions to

been assessed assuming year 2020 EIS.

costs also increases. The question therefore rises:

improve the overall quality and clarity of the article.

OPR Engine overall pressure ratio SFC Engine specific fuel consumption T4 Combustor outlet temperature

**6. Conclusions**

in detail.

**market?**

**7. Acknowledgements**

**8. Nomenclature**

come.


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**2** 

*Italy* 

**State-of-Art of Transonic Axial Compressors** 

Transonic axial flow compressors are today widely used in aircraft engines to obtain maximum pressure ratios per single-stage. High stage pressure ratios are important because they make it possible to reduce the engine weight and size and, therefore, investment and operational costs. Performance of transonic compressors has today reached a high level but engine manufacturers are oriented towards increasing it further. A small increment in efficiency, for instance, can result in huge savings in fuel costs and determine a key factor for product success. Another important target is the improvement of rotor stability towards

Important analytical and experimental researches in the field of transonic compressors were carried out since 1960's (e.g. Chen et al., 1991; Epstein, 1977; Freeman & Cumpsty, 1992; König et al., 1996; Miller et al., 1961; Wennerstrom & Puterbaugh, 1984). A considerable contribution for the new developments and designs was the progress made in optical measurement techniques and computational methods, leading to a deeper understanding of the loss mechanisms of supersonic relative flow in compressors (e.g. Calvert & Stapleton, 1994; Hah & Reid, 1992; Ning & Xu, 2001; Puterbaugh et al., 1997; Strazisar, 1985; Weyer & Dunker, 1978). Fig. 1 shows the low pressure and high pressure compressors of the EJ200 engine as examples for highly loaded, high performance transonic rotors of an aero engine. A closer look at the current trend in design parameters for axial flow transonic compressors shows that, especially in civil aircraft engines, the relative flow tip Mach number of the rotor is limited to maintain high efficiencies. A typical value for the rotor inlet relative flow at the tip is Mach ≈ 1.3. The continuous progress of aerodynamics has been focused to the increase in efficiency and pressure ratio and to the improvement in off-design behaviour at roughly the same level of the inlet relative Mach number. Today's high efficiency transonic axial flow compressors give a total pressure ratio in the order of 1.7-1.8, realized by combining high rotor speeds (tip speed in the order of 500 m/s) and high stage loadings (2Δh/u² in the order of 1.0). The rotor blade aspect ratio parameter showed a general trend towards lower values during past decades, with a current asymptotic value of 1.2 (Broichhausen & Ziegler,

The flow field that develops inside a transonic compressor rotor is extremely complex and presents many challenges to compressor designers, who have to deal with several and concurring flow features such as shock waves, intense secondary flows, shock/boundary layer interaction, etc., inducing energy losses and efficiency reduction (Calvert et al., 2003; Cumpsty, 1989; Denton & Xu, 1999; Law & Wadia, 1993; Sun et al., 2007). Interacting with secondary flows, shock waves concur in development of blockage (Suder, 1998), in corner

near stall conditions, resulting in a wider working range.

**1. Introduction** 

2005).

Roberto Biollo and Ernesto Benini

*University of Padova* 


## **State-of-Art of Transonic Axial Compressors**

Roberto Biollo and Ernesto Benini

*University of Padova Italy* 

#### **1. Introduction**

22 Will-be-set-by-IN-TECH

24 Advances in Gas Turbine Technology

Wilfert, G., Sieber, J., Rolt, A., Baker, N., Touyeras, A. & Colantuoni, S. (2007). New

Xu, L. & Grönstedt, T. (2010). Design and Analysis of an Intercooled Turbofan Engine, *ASME Journal of Engineering for Gas Turbines and Power* 132(11). doi:10.1115/1.4000857. Xu, L., Gustafsson, B. & Grönstedt, T. (2007). Mission Optimization of an Intercooled Turbofan

Young, P. (1979). The future shape of medium and long-range civil engines, *RAeS Aeronautical*

Zimbrick, R. & Colehour, J. (1988). An investigation of very high bypass ratio engines

for subsonic transports, *Proceedings of AIAA/SAE/ASME/ASEE 24th Joint Propulsion*

Engine, *ISABE 2007 Proceedings, ISABE-2007-1157*, Beijing, China.

*Conference, AIAA-88-2953*, Boston, Massachusetts, USA.

*ISABE-2007-1120*, Beijing, China.

*Journal* pp. 53–61.

Environmental Friendly Aero Engine Core Concepts, *ISABE 2007 Proceedings,*

Transonic axial flow compressors are today widely used in aircraft engines to obtain maximum pressure ratios per single-stage. High stage pressure ratios are important because they make it possible to reduce the engine weight and size and, therefore, investment and operational costs. Performance of transonic compressors has today reached a high level but engine manufacturers are oriented towards increasing it further. A small increment in efficiency, for instance, can result in huge savings in fuel costs and determine a key factor for product success. Another important target is the improvement of rotor stability towards near stall conditions, resulting in a wider working range.

Important analytical and experimental researches in the field of transonic compressors were carried out since 1960's (e.g. Chen et al., 1991; Epstein, 1977; Freeman & Cumpsty, 1992; König et al., 1996; Miller et al., 1961; Wennerstrom & Puterbaugh, 1984). A considerable contribution for the new developments and designs was the progress made in optical measurement techniques and computational methods, leading to a deeper understanding of the loss mechanisms of supersonic relative flow in compressors (e.g. Calvert & Stapleton, 1994; Hah & Reid, 1992; Ning & Xu, 2001; Puterbaugh et al., 1997; Strazisar, 1985; Weyer & Dunker, 1978). Fig. 1 shows the low pressure and high pressure compressors of the EJ200 engine as examples for highly loaded, high performance transonic rotors of an aero engine.

A closer look at the current trend in design parameters for axial flow transonic compressors shows that, especially in civil aircraft engines, the relative flow tip Mach number of the rotor is limited to maintain high efficiencies. A typical value for the rotor inlet relative flow at the tip is Mach ≈ 1.3. The continuous progress of aerodynamics has been focused to the increase in efficiency and pressure ratio and to the improvement in off-design behaviour at roughly the same level of the inlet relative Mach number. Today's high efficiency transonic axial flow compressors give a total pressure ratio in the order of 1.7-1.8, realized by combining high rotor speeds (tip speed in the order of 500 m/s) and high stage loadings (2Δh/u² in the order of 1.0). The rotor blade aspect ratio parameter showed a general trend towards lower values during past decades, with a current asymptotic value of 1.2 (Broichhausen & Ziegler, 2005).

The flow field that develops inside a transonic compressor rotor is extremely complex and presents many challenges to compressor designers, who have to deal with several and concurring flow features such as shock waves, intense secondary flows, shock/boundary layer interaction, etc., inducing energy losses and efficiency reduction (Calvert et al., 2003; Cumpsty, 1989; Denton & Xu, 1999; Law & Wadia, 1993; Sun et al., 2007). Interacting with secondary flows, shock waves concur in development of blockage (Suder, 1998), in corner

State-of-Art of Transonic Axial Compressors 27

interaction or shock-induced separation. The reduction of the pre-shock Mach number can be achieved by zero or even negative curvature in the front part of the blade suction side and by a resulting pre-compression shock system reducing the Mach number upstream of

Besides inducing energy losses, the presence of shock waves makes transonic compressors particularly sensitive to variations in blade section design. An investigation of cascade throat area, internal contraction, and trailing edge effective camber on compressor performance showed that small changes in meanline angles, and consequently in the airfoil shape and passage area ratios, significantly affect the performance of transonic blade rows

One of the most important airfoil design parameter affecting the aerodynamics of transonic bladings is the chordwise location of maximum thickness. An experimental and numerical evaluation of two versions of a low aspect ratio transonic rotor having the location of the tip blade section maximum thickness at 55% and 40% chord length respectively, showed that the more aft position of maximum thickness is preferred for the best high speed performance, keeping the edge and maximum thickness values the same (Wadia & Law, 1993). The better performance was associated with the lower shock front losses with the finer section that results when the location of the maximum thickness is moved aft. The existence of an optimum maximum thickness location at 55% to 60% chord length for such rotor was hypothesized. Similar results can be found in a recent work (Chen et al., 2007) describing an optimization methodology for the aerodynamic design of turbomachinery applied to a transonic compressor bladings and showing how the thermal loss coefficient decreases with increasing the maximum thickness location for all the

Not only the position of maximum thickness but also the airfoil thickness has been showed to have a significant impact on the aerodynamic behaviour of transonic compressor rotors, as observed in an investigation on surface roughness and airfoil thickness effects (Suder et al., 1995). In this work, a 0.025 mm thick smooth coating was applied to the pressure and suction surface of the rotor blades, increasing the leading edge thickness by 10% at the hub and 20% at the tip. The smooth coating surface finish was comparable to the bare metal blade surface finish; therefore the coating did not increase roughness over the blade, except at the leading edge where roughness increased due to particle impact damage. It resulted in a 4% loss in pressure ratio across the rotor at an operating point near design mass flow, with the largest degradation in pressure rise over the outer half of the blade span. When assessed at a constant pressure ratio, the adiabatic efficiency degradation at design speed was in the

The recent development of optimization tools coupled with accurate CFD codes has improved the turbomachinery design process significantly, making it faster and more efficient. The application to the blade section design, with a quasi three-dimensional and more recently with a fully three-dimensional approach, can lead to optimal blade geometries in terms of aerodynamic performance at both design and off-design operating conditions. Such a design process is particularly successful in the field of transonic

Fig. 2 shows the blade deformation obtained in a quasi 3-D numerical optimization process of a transonic compressor blade section along with the relative Mach number contours before and after the optimization (Burguburu et al., 2004). As shown, no modifications of the

compressors, where performance is highly sensitive to little changes in airfoil design.

the final strong passage shock.

(Wadia & Copenhaver, 1996).

sections from hub to tip.

order of 3-6 points.

stall separation (Hah & Loellbach, 1999; Weber et al., 2002), in upstream wakes destabilization (Estevadeordal et al., 2007; Prasad, 2003), and in many other negative flow phenomena. Particularly detrimental is the interaction with the tip clearance flow at the outer span of the rotor, where the compressor generally shows the higher entropy production (Bergner et al., 2005a; Chima, 1998; Copenhaver et al., 1996; Gerolymos & Vallet, 1999; Hofmann & Ballmann, 2002; Puterbaugh & Brendel, 1997; Suder & Celestina, 1996).

Fig. 1. Transonic LPC (left) and HPC (right) of the Eurofighter Typhoon engine EJ200 (Broichhausen & Ziegler, 2005)

As the compressor moves from peak to near-stall operating point, the blade loading increases and flow structures become stronger and unsteady. The tip leakage vortex can breakdown interacting with the passage shock wave, leading to not only a large blockage effect near the tip but also a self-sustained flow oscillation in the rotor passage. As a result, the blade torque, the low energy fluid flow due to the shock/tip leakage vortex interaction and the shock-induced flow separation on the blade suction surface fluctuate with time (Yamada et al., 2004).

Despite the presence of such flow unsteadiness, the compressor can still operate in a stable mode. Rotating stall arises when the loading is further increased, i.e. at a condition of lower mass flow rate. Two routes to rotating stall have been identified: long length-scale (modal) and short length-scale (spike) stall inception in axial compressors (Day, 1993). Modal stall inception is characterized by the relatively slow growth (over 10-40 rotor revolutions) of a small disturbance of long circumferential wavelength into a fully developed stall cell. Spike stall inception starts with the appearance of a large amplitude short length-scale (two to three rotor blade passages) disturbance at the rotor tip, the so-called spike, which grows into a fully developed rotating stall cell within few rotor revolutions.

The following paragraphs give a summary of the possible techniques for limiting the negative impacts of the above reported compressor flow features in aircraft gas turbine engines.

## **2. Blade profiles studies**

For relative inlet Mach numbers in the order of 1.3 and higher the most important design intent is to reduce the Mach number in front of the passage shock. This is of primary importance due to the strongly rising pressure losses with increasing pre-shock Mach number, and because of the increasing pressure losses due to the shock/boundary layer

stall separation (Hah & Loellbach, 1999; Weber et al., 2002), in upstream wakes destabilization (Estevadeordal et al., 2007; Prasad, 2003), and in many other negative flow phenomena. Particularly detrimental is the interaction with the tip clearance flow at the outer span of the rotor, where the compressor generally shows the higher entropy production (Bergner et al., 2005a; Chima, 1998; Copenhaver et al., 1996; Gerolymos & Vallet, 1999; Hofmann & Ballmann, 2002; Puterbaugh & Brendel, 1997; Suder & Celestina, 1996).

Fig. 1. Transonic LPC (left) and HPC (right) of the Eurofighter Typhoon engine EJ200

As the compressor moves from peak to near-stall operating point, the blade loading increases and flow structures become stronger and unsteady. The tip leakage vortex can breakdown interacting with the passage shock wave, leading to not only a large blockage effect near the tip but also a self-sustained flow oscillation in the rotor passage. As a result, the blade torque, the low energy fluid flow due to the shock/tip leakage vortex interaction and the shock-induced flow separation on the blade suction surface fluctuate with time

Despite the presence of such flow unsteadiness, the compressor can still operate in a stable mode. Rotating stall arises when the loading is further increased, i.e. at a condition of lower mass flow rate. Two routes to rotating stall have been identified: long length-scale (modal) and short length-scale (spike) stall inception in axial compressors (Day, 1993). Modal stall inception is characterized by the relatively slow growth (over 10-40 rotor revolutions) of a small disturbance of long circumferential wavelength into a fully developed stall cell. Spike stall inception starts with the appearance of a large amplitude short length-scale (two to three rotor blade passages) disturbance at the rotor tip, the so-called spike, which grows into

The following paragraphs give a summary of the possible techniques for limiting the negative impacts of the above reported compressor flow features in aircraft gas turbine

For relative inlet Mach numbers in the order of 1.3 and higher the most important design intent is to reduce the Mach number in front of the passage shock. This is of primary importance due to the strongly rising pressure losses with increasing pre-shock Mach number, and because of the increasing pressure losses due to the shock/boundary layer

a fully developed rotating stall cell within few rotor revolutions.

(Broichhausen & Ziegler, 2005)

(Yamada et al., 2004).

engines.

**2. Blade profiles studies** 

interaction or shock-induced separation. The reduction of the pre-shock Mach number can be achieved by zero or even negative curvature in the front part of the blade suction side and by a resulting pre-compression shock system reducing the Mach number upstream of the final strong passage shock.

Besides inducing energy losses, the presence of shock waves makes transonic compressors particularly sensitive to variations in blade section design. An investigation of cascade throat area, internal contraction, and trailing edge effective camber on compressor performance showed that small changes in meanline angles, and consequently in the airfoil shape and passage area ratios, significantly affect the performance of transonic blade rows (Wadia & Copenhaver, 1996).

One of the most important airfoil design parameter affecting the aerodynamics of transonic bladings is the chordwise location of maximum thickness. An experimental and numerical evaluation of two versions of a low aspect ratio transonic rotor having the location of the tip blade section maximum thickness at 55% and 40% chord length respectively, showed that the more aft position of maximum thickness is preferred for the best high speed performance, keeping the edge and maximum thickness values the same (Wadia & Law, 1993). The better performance was associated with the lower shock front losses with the finer section that results when the location of the maximum thickness is moved aft. The existence of an optimum maximum thickness location at 55% to 60% chord length for such rotor was hypothesized. Similar results can be found in a recent work (Chen et al., 2007) describing an optimization methodology for the aerodynamic design of turbomachinery applied to a transonic compressor bladings and showing how the thermal loss coefficient decreases with increasing the maximum thickness location for all the sections from hub to tip.

Not only the position of maximum thickness but also the airfoil thickness has been showed to have a significant impact on the aerodynamic behaviour of transonic compressor rotors, as observed in an investigation on surface roughness and airfoil thickness effects (Suder et al., 1995). In this work, a 0.025 mm thick smooth coating was applied to the pressure and suction surface of the rotor blades, increasing the leading edge thickness by 10% at the hub and 20% at the tip. The smooth coating surface finish was comparable to the bare metal blade surface finish; therefore the coating did not increase roughness over the blade, except at the leading edge where roughness increased due to particle impact damage. It resulted in a 4% loss in pressure ratio across the rotor at an operating point near design mass flow, with the largest degradation in pressure rise over the outer half of the blade span. When assessed at a constant pressure ratio, the adiabatic efficiency degradation at design speed was in the order of 3-6 points.

The recent development of optimization tools coupled with accurate CFD codes has improved the turbomachinery design process significantly, making it faster and more efficient. The application to the blade section design, with a quasi three-dimensional and more recently with a fully three-dimensional approach, can lead to optimal blade geometries in terms of aerodynamic performance at both design and off-design operating conditions. Such a design process is particularly successful in the field of transonic compressors, where performance is highly sensitive to little changes in airfoil design.

Fig. 2 shows the blade deformation obtained in a quasi 3-D numerical optimization process of a transonic compressor blade section along with the relative Mach number contours before and after the optimization (Burguburu et al., 2004). As shown, no modifications of the

State-of-Art of Transonic Axial Compressors 29

At 10% and 50% span (not shown here), the optimization gave a larger camber but lower thickness than the baseline design. The thinner airfoils contributed to reduce the weight of the new design. The calculated difference in the pressure distribution was rather small. At 90% span (see Fig. 3), the new design had a slightly smaller camber and thinner airfoil than the baseline. Nevertheless, the calculated pressure difference was rather large, indicating again that transonic flow is highly sensitive to the profile shape change. One noticeable impact was also in the shock position. The new design showed a more forward passage

Such optimized blade gave a decrease of 5.4% in weight and an improvement of 1.8% in the total pressure ratio. The lighter weight came from the thinner blade shape. The higher total pressure ratio was mainly attributed to a reduced separation zone after the shock at the outer span. In Fig. 3, the separation zones are characterized by streamlines going towards the separation lines, whereas reattachment lines look like flow is going away from the separation lines. Compared with the baseline design, downstream of the shock the new design gave a smaller separation zone, which was partially responsible for its higher total

Fig. 4 is again related to the redesign of rotor 67 using an optimization tool based on evolutionary algorithms (Oyama et al., 2004). Note the particular new design, an improbable design using manual techniques. The optimization gave rise to a double-hump blade shape,

In such new design, the flow acceleration near the leading edge at 33% span diminished because of the decrease of the incidence angle. In addition, at the 90% span, the shock on the suction side moved aft and was weakened considerably because of the aft movement of the maximum camber position. This new blade showed an overall adiabatic efficiency of 2%

higher than the baseline blade over the entire operating range for the design speed.

 Fig. 4. Comparison between the optimized and baseline design at 33% and 90% span

The preceding paragraph has shown that a certain maturity in transonic compressors has been reached regarding the general airfoil aerodesign. But the flow field in a compressor is not only influenced by the two-dimensional airfoil geometry. The three-dimensional shape of the blade is also of great importance, especially in transonic compressor rotors where an optimization of shock structure and its interference with secondary flows is required. Many experimental and numerical works can be found in the literature on the design and analysis of three-dimensional shaped transonic bladings (e.g. Copenhaver et al., 1996; Hah et al.,

shock than the baseline.

especially obvious on the pressure side.

pressure ratio.

(Oyama et al., 2004)

**3. Three-dimensional shaped bladings** 

inlet flow field occurred after optimization but the flow field structure in the duct is clearly different. The negative curvature of the blade upstream of the shock led to the reduction of the upstream relative Mach number from 1.4 to 1.2. With this curvature change, the velocity slowdown is better driven. Instead of creating a normal shock, the new shape created two low intensity shocks. The new blade gave an efficiency increment of 1.75 points at design condition, without changing the choking mass flow. A large part of the efficiency improvement at the design condition remained at off-design conditions.

Fig. 2. Blade deformation (left) and relative Mach number contours (right) before and after optimization (Burguburu et al., 2004)

Fig. 3 is related to a both aerodynamic and structural optimization of the well-known transonic compressor rotor 67 (Strazisar et al., 1989), where the aerodynamic objective aimed at maximizing the total pressure ratio whereas the structural objective was to minimize the blade weight, with the constraint that the new design had comparable mass flow rate as the baseline design (Lian & Liou, 2005). The optimization was carried out at the design operating point. Geometric modifications regarded the mean camber line (with the leading and trailing edge points fixed) and thickness distribution of four airfoil profiles (hub, 31% span, 62% span, and tip), linearly interpolated to obtained the new 3-D blade. The chord distribution along the span and the meridional contours of hub, casing, sweep, and lean were maintained.

Fig. 3. Blade section at 90% span (left) and streamlines close to the blade suction side (right) before and after the optimization (modified from Lian & Liou, 2005)

inlet flow field occurred after optimization but the flow field structure in the duct is clearly different. The negative curvature of the blade upstream of the shock led to the reduction of the upstream relative Mach number from 1.4 to 1.2. With this curvature change, the velocity slowdown is better driven. Instead of creating a normal shock, the new shape created two low intensity shocks. The new blade gave an efficiency increment of 1.75 points at design condition, without changing the choking mass flow. A large part of the efficiency

Fig. 2. Blade deformation (left) and relative Mach number contours (right) before and after

Fig. 3 is related to a both aerodynamic and structural optimization of the well-known transonic compressor rotor 67 (Strazisar et al., 1989), where the aerodynamic objective aimed at maximizing the total pressure ratio whereas the structural objective was to minimize the blade weight, with the constraint that the new design had comparable mass flow rate as the baseline design (Lian & Liou, 2005). The optimization was carried out at the design operating point. Geometric modifications regarded the mean camber line (with the leading and trailing edge points fixed) and thickness distribution of four airfoil profiles (hub, 31% span, 62% span, and tip), linearly interpolated to obtained the new 3-D blade. The chord distribution along the span and the meridional contours of hub, casing, sweep, and lean

 Fig. 3. Blade section at 90% span (left) and streamlines close to the blade suction side (right)

before and after the optimization (modified from Lian & Liou, 2005)

improvement at the design condition remained at off-design conditions.

optimization (Burguburu et al., 2004)

were maintained.

At 10% and 50% span (not shown here), the optimization gave a larger camber but lower thickness than the baseline design. The thinner airfoils contributed to reduce the weight of the new design. The calculated difference in the pressure distribution was rather small. At 90% span (see Fig. 3), the new design had a slightly smaller camber and thinner airfoil than the baseline. Nevertheless, the calculated pressure difference was rather large, indicating again that transonic flow is highly sensitive to the profile shape change. One noticeable impact was also in the shock position. The new design showed a more forward passage shock than the baseline.

Such optimized blade gave a decrease of 5.4% in weight and an improvement of 1.8% in the total pressure ratio. The lighter weight came from the thinner blade shape. The higher total pressure ratio was mainly attributed to a reduced separation zone after the shock at the outer span. In Fig. 3, the separation zones are characterized by streamlines going towards the separation lines, whereas reattachment lines look like flow is going away from the separation lines. Compared with the baseline design, downstream of the shock the new design gave a smaller separation zone, which was partially responsible for its higher total pressure ratio.

Fig. 4 is again related to the redesign of rotor 67 using an optimization tool based on evolutionary algorithms (Oyama et al., 2004). Note the particular new design, an improbable design using manual techniques. The optimization gave rise to a double-hump blade shape, especially obvious on the pressure side.

In such new design, the flow acceleration near the leading edge at 33% span diminished because of the decrease of the incidence angle. In addition, at the 90% span, the shock on the suction side moved aft and was weakened considerably because of the aft movement of the maximum camber position. This new blade showed an overall adiabatic efficiency of 2% higher than the baseline blade over the entire operating range for the design speed.

Fig. 4. Comparison between the optimized and baseline design at 33% and 90% span (Oyama et al., 2004)

## **3. Three-dimensional shaped bladings**

The preceding paragraph has shown that a certain maturity in transonic compressors has been reached regarding the general airfoil aerodesign. But the flow field in a compressor is not only influenced by the two-dimensional airfoil geometry. The three-dimensional shape of the blade is also of great importance, especially in transonic compressor rotors where an optimization of shock structure and its interference with secondary flows is required. Many experimental and numerical works can be found in the literature on the design and analysis of three-dimensional shaped transonic bladings (e.g. Copenhaver et al., 1996; Hah et al.,

State-of-Art of Transonic Axial Compressors 31

calculated flow fields indicated that two mechanisms were primarily responsible for the differences in aerodynamic performance among these rotors. The first mechanism was a change in the radial shape of the passage shock near the casing by the endwall effect, and the second was the radial migration of low momentum fluid to the blade tip region. Similar results were obtained in a parallel investigation which identified the reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the tip, as the prime

contributor to the enhanced performance with forward sweep (Wadia et al., 1998).

Fig. 6. Blade axial curvature impact on shock, suction side boundary layer and blade wake

A recent numerical work gave another point of view on the impact of blade curvature in transonic compressor rotors, showing how the movement of blade sections in the axial direction can influence the internal flow field (Benini & Biollo, 2007; Biollo & Benini, 2008a). Such work showed that the axial blade curvature can help to influence the shock shape in the meridional plane, inducing the shock to assume the meridional curvature of the blade leading edge (Fig. 6). In addition, a considerable impact on the radial outward migration of

development (Biollo & Benini, 2008a)

2004; Puterbaugh et al., 1997). Fig. 5 shows two examples of non-conventional rotors (Rotor2 and Rotor3) derived from the baseline Rotor1 which is conventionally radially-stacked, all developed by TU Darmstadt and MTU Aero Engines. As far as their performance is concerned, Rotor2 gave no real improvement in efficiency and total pressure ratio with respect to the baseline configuration (Blaha et al., 2000; Kablitz et al., 2003a). Rotor3, instead, gave higher performance at design speed (1.5% peak efficiency increment) along with a significantly wider operating range (Passrucker et al., 2003). Information on the favourable impact of Rotor3 blade design on internal transonic flow field is available in the open literature (Bergner et al., 2005b; Kablitz et al., 2003b).

Fig. 5. Transonic compressor test rotors – TU Darmstadt and MTU Aero Engines (Broichhausen & Ziegler, 2005; Passrucker et al., 2003)

A numerical investigation on the aerodynamics of 3-D shaped blades in transonic compressor rotors showed the possibility to have better stall margin with forward sweep (upstream movement of blade sections along the local chord direction, especially at outer span region), maintaining a high efficiency over a wider range (Denton, 2002; Denton & Xu, 2002). This seems to be a general point of view, as confirmed by the following researches.

Numerical and experimental analyses carried out to evaluate the performance of a conventional unswept rotor, a forward swept rotor and an aft swept rotor showed that the forward swept rotor had a higher peak efficiency and a substantially larger stall margin than the baseline unswept rotor, and that the aft swept rotor had a similar peak efficiency with a significantly smaller stall margin (Hah et al., 1998). Detailed analyses of the measured and

2004; Puterbaugh et al., 1997). Fig. 5 shows two examples of non-conventional rotors (Rotor2 and Rotor3) derived from the baseline Rotor1 which is conventionally radially-stacked, all developed by TU Darmstadt and MTU Aero Engines. As far as their performance is concerned, Rotor2 gave no real improvement in efficiency and total pressure ratio with respect to the baseline configuration (Blaha et al., 2000; Kablitz et al., 2003a). Rotor3, instead, gave higher performance at design speed (1.5% peak efficiency increment) along with a significantly wider operating range (Passrucker et al., 2003). Information on the favourable impact of Rotor3 blade design on internal transonic flow field is available in the open

Fig. 5. Transonic compressor test rotors – TU Darmstadt and MTU Aero Engines

A numerical investigation on the aerodynamics of 3-D shaped blades in transonic compressor rotors showed the possibility to have better stall margin with forward sweep (upstream movement of blade sections along the local chord direction, especially at outer span region), maintaining a high efficiency over a wider range (Denton, 2002; Denton & Xu, 2002). This seems to be a general point of view, as confirmed by the following researches. Numerical and experimental analyses carried out to evaluate the performance of a conventional unswept rotor, a forward swept rotor and an aft swept rotor showed that the forward swept rotor had a higher peak efficiency and a substantially larger stall margin than the baseline unswept rotor, and that the aft swept rotor had a similar peak efficiency with a significantly smaller stall margin (Hah et al., 1998). Detailed analyses of the measured and

(Broichhausen & Ziegler, 2005; Passrucker et al., 2003)

literature (Bergner et al., 2005b; Kablitz et al., 2003b).

calculated flow fields indicated that two mechanisms were primarily responsible for the differences in aerodynamic performance among these rotors. The first mechanism was a change in the radial shape of the passage shock near the casing by the endwall effect, and the second was the radial migration of low momentum fluid to the blade tip region. Similar results were obtained in a parallel investigation which identified the reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the tip, as the prime contributor to the enhanced performance with forward sweep (Wadia et al., 1998).

Fig. 6. Blade axial curvature impact on shock, suction side boundary layer and blade wake development (Biollo & Benini, 2008a)

A recent numerical work gave another point of view on the impact of blade curvature in transonic compressor rotors, showing how the movement of blade sections in the axial direction can influence the internal flow field (Benini & Biollo, 2007; Biollo & Benini, 2008a). Such work showed that the axial blade curvature can help to influence the shock shape in the meridional plane, inducing the shock to assume the meridional curvature of the blade leading edge (Fig. 6). In addition, a considerable impact on the radial outward migration of

State-of-Art of Transonic Axial Compressors 33

casing treatments investigated in the 1970's. The interaction of the main flow with the flow circulating in these cavities seems to have a positive impact on rotor stability. However, early studies did not reach a detailed understanding of the phenomenon, since experimental investigations were too expensive and only few configurations could be tested. Only in the past fifteen years numerical simulations made it possible to investigate a larger number of casing treatment solutions and their effects on different compressors. Many researches were carried out on transonic compressor rotors and the potential of this kind of passive devices was revealed: a proper treatment can not only widen the stable working range of a transonic

Fig. 8. Various casing treatments investigated in the 1970's (Hathaway, 2007)

Recently, the influence of circumferential grooves on the tip flow field of an axial singlestage transonic compressor has been examined both experimentally and numerically (Fig. 9). The compressor stage provided a strongly increased stall margin (56.1%), with only small penalties in efficiency when the casing treatment was applied. Flow analyses showed that at near stall conditions with the smooth casing, the induced vortex originating from the tip clearance flow crossing the tip gap along 20-50% chord length, hits the front part of the adjacent blade, indicating the possibility of a spill forward of low momentum fluid into the next passage, a flow feature considered a trigger for the onset of rotating stall. With the casing treatment applied, the vortex trajectory remained instead aligned to the blade's

Disadvantages of casing treatments like these are the space they need and the weight increase of the compressor casing. So it is a goal to maintain the positive effects (increased surge pressure ratio in combination with high efficiency) while at the same time reducing the geometric volume of the device. On this regard, an experimental and numerical investigation on the first rotor of a two-stage compressor showed that grooves with a much

compressor rotor, but also improve its efficiency.

**4.1 Circumferential groove-type treatments** 

suction side.

fluid particles which takes place inside the blade suction side boundary layer after the interaction with the shock has been confirmed. The code predicted a reduction of the strength of such flow feature when the blade is curved downstream and an increment when the blade is curved upstream. Such flow phenomenon is harmful because obstructs the boundary layer development in the streamwise direction, leading to a thickening of blade wakes. A reduction of its strength helped to reduce the entropy generation and the aerodynamic losses associated with the blade wake development. The possibility to increase the peak efficiency of 0.8% at design speed using a proper downstream blade curvature has been showed for the high loaded transonic compressor rotor 37. Details on rotor 37 can be found in the open literature (Reid & Moore, 1978).

The same research group investigated the aerodynamic effects induced by several tangential blade curvatures on the same rotor. It was observed that, when the curvature is applied towards the direction of rotor rotation, the blade-to-blade shock tends to move more downstream, becoming more oblique to the incoming flow. This reduced the aerodynamic shock losses and entropy generation, showing in some cases a peak efficiency increment of over 1% at design speed (Benini & Biollo, 2008). Similar results were previously obtained using a numerical optimization algorithm (Ahn & Kim, 2002). Fig. 7 shows the predicted impact of the optimized design of rotor 37 on the blade-to-blade Mach number.

Fig. 7. Baseline (left) and optimized (right) Mach number distributions at 90% span (modified from Ahn & Kim, 2002)

Higher performance can be achieved using a proper combination of two orthogonal blade curvatures, i.e. the use of a blade curved both axially and tangentially, as well as swept and leaned at the same time. Peak efficiency increments from 1% to 1.5% were numerically observed using a blade prevalently curved towards the direction of rotor rotation and slightly backward inclined (Biollo & Benini, 2008b; Jang et al., 2006; Yi et al., 2006).

## **4. Casing treatments**

Hollow structures in the casing to improve the tip endwall flow field of axial flow compressors are commonly referred to as casing treatments. Fig. 8 shows some examples of

fluid particles which takes place inside the blade suction side boundary layer after the interaction with the shock has been confirmed. The code predicted a reduction of the strength of such flow feature when the blade is curved downstream and an increment when the blade is curved upstream. Such flow phenomenon is harmful because obstructs the boundary layer development in the streamwise direction, leading to a thickening of blade wakes. A reduction of its strength helped to reduce the entropy generation and the aerodynamic losses associated with the blade wake development. The possibility to increase the peak efficiency of 0.8% at design speed using a proper downstream blade curvature has been showed for the high loaded transonic compressor rotor 37. Details on rotor 37 can be

The same research group investigated the aerodynamic effects induced by several tangential blade curvatures on the same rotor. It was observed that, when the curvature is applied towards the direction of rotor rotation, the blade-to-blade shock tends to move more downstream, becoming more oblique to the incoming flow. This reduced the aerodynamic shock losses and entropy generation, showing in some cases a peak efficiency increment of over 1% at design speed (Benini & Biollo, 2008). Similar results were previously obtained using a numerical optimization algorithm (Ahn & Kim, 2002). Fig. 7 shows the predicted

impact of the optimized design of rotor 37 on the blade-to-blade Mach number.

Fig. 7. Baseline (left) and optimized (right) Mach number distributions at 90% span

slightly backward inclined (Biollo & Benini, 2008b; Jang et al., 2006; Yi et al., 2006).

Higher performance can be achieved using a proper combination of two orthogonal blade curvatures, i.e. the use of a blade curved both axially and tangentially, as well as swept and leaned at the same time. Peak efficiency increments from 1% to 1.5% were numerically observed using a blade prevalently curved towards the direction of rotor rotation and

Hollow structures in the casing to improve the tip endwall flow field of axial flow compressors are commonly referred to as casing treatments. Fig. 8 shows some examples of

found in the open literature (Reid & Moore, 1978).

(modified from Ahn & Kim, 2002)

**4. Casing treatments** 

casing treatments investigated in the 1970's. The interaction of the main flow with the flow circulating in these cavities seems to have a positive impact on rotor stability. However, early studies did not reach a detailed understanding of the phenomenon, since experimental investigations were too expensive and only few configurations could be tested. Only in the past fifteen years numerical simulations made it possible to investigate a larger number of casing treatment solutions and their effects on different compressors. Many researches were carried out on transonic compressor rotors and the potential of this kind of passive devices was revealed: a proper treatment can not only widen the stable working range of a transonic compressor rotor, but also improve its efficiency.

Fig. 8. Various casing treatments investigated in the 1970's (Hathaway, 2007)

#### **4.1 Circumferential groove-type treatments**

Recently, the influence of circumferential grooves on the tip flow field of an axial singlestage transonic compressor has been examined both experimentally and numerically (Fig. 9). The compressor stage provided a strongly increased stall margin (56.1%), with only small penalties in efficiency when the casing treatment was applied. Flow analyses showed that at near stall conditions with the smooth casing, the induced vortex originating from the tip clearance flow crossing the tip gap along 20-50% chord length, hits the front part of the adjacent blade, indicating the possibility of a spill forward of low momentum fluid into the next passage, a flow feature considered a trigger for the onset of rotating stall. With the casing treatment applied, the vortex trajectory remained instead aligned to the blade's suction side.

Disadvantages of casing treatments like these are the space they need and the weight increase of the compressor casing. So it is a goal to maintain the positive effects (increased surge pressure ratio in combination with high efficiency) while at the same time reducing the geometric volume of the device. On this regard, an experimental and numerical investigation on the first rotor of a two-stage compressor showed that grooves with a much

State-of-Art of Transonic Axial Compressors 35

induced a smaller inflow angle near the leading edge, i.e. a lower incidence, delaying

Another way to treat the casing with the aim to improve the rotor performance and stability is exemplified in Fig. 12. Here the casing wall is circumferentially treated with a discrete

A similar slot-type casing treatment consisting of four identical axial slots per blade passage and having an open area of 50% in the circumferential direction has been also proposed. (Wilke and Kau, 2004). In this case, the slots are parallel to the rotation axis and inclined by 45° against the meridional plane in the direction of rotor rotation. The slot shape is designed as a semi-circle. Two configurations have been numerically tested. In configuration 1, the position of the slots is centered above the rotor blade tip reaching from 7.5% to 92.5% chord length. In configuration 2, the slots are moved upstream so that only 25% chord length

Fig. 12. EJ200 LPC with axial slots casing treatment (Broichhausen & Ziegler, 2005)

rotating stall inception.

**4.2 Slot-type treatments** 

Fig. 11. Endwall casing contouring (Ito et al., 2007)

remains covered by the casing treatment.

number of axial rectangular slots over the blade tip section.

smaller depth than conventional designs are similarly effective in increasing the stall margin (Rabe & Hah, 2002). The same work also showed that two shallow grooves placed near the leading edge are better than five deep or shallow grooves all over the blade tip. Fewer shallower grooves clearly help to reduce the weight, fabrication costs, and loss generation associated with such a casing treatment.

Fig. 9. Cross section of a transonic compressor stage with circumferential grooves (modified from Müller et al., 2008)

Other possible groove-type casing treatment solutions are presented in Figs. 10 and 11. Fig. 10 shows a single extended casing circumferential groove all over the blade tip section. It has been numerically shown that such a casing treatment provides a means for fluid to exit the flow path where the blade loading is severe, migrate circumferentially, and re-enter the flow path at a location where the pressure is more moderate. This can lead to stability improvement since the flow relocation helps to relieve the locally severe blade loading. Using this device, the authors showed the possibility to improve both the efficiency and the stall margin.

Fig. 10. Single extended casing circumferential groove (Beheshti et al., 2004)

Fig. 11 is related to a numerical investigation of casing contouring effects on flow instability. While the "Type B1" solution gave no improvements in stall margin, the "Type C2" solution extended the calculated stable operating range from 94% to about 92.5% normalized mass flow. With respect to the baseline smooth casing configuration, the successful contouring

smaller depth than conventional designs are similarly effective in increasing the stall margin (Rabe & Hah, 2002). The same work also showed that two shallow grooves placed near the leading edge are better than five deep or shallow grooves all over the blade tip. Fewer shallower grooves clearly help to reduce the weight, fabrication costs, and loss generation

Fig. 9. Cross section of a transonic compressor stage with circumferential grooves (modified

Other possible groove-type casing treatment solutions are presented in Figs. 10 and 11. Fig. 10 shows a single extended casing circumferential groove all over the blade tip section. It has been numerically shown that such a casing treatment provides a means for fluid to exit the flow path where the blade loading is severe, migrate circumferentially, and re-enter the flow path at a location where the pressure is more moderate. This can lead to stability improvement since the flow relocation helps to relieve the locally severe blade loading. Using this device, the authors showed the possibility to improve both the efficiency and the

Fig. 10. Single extended casing circumferential groove (Beheshti et al., 2004)

Fig. 11 is related to a numerical investigation of casing contouring effects on flow instability. While the "Type B1" solution gave no improvements in stall margin, the "Type C2" solution extended the calculated stable operating range from 94% to about 92.5% normalized mass flow. With respect to the baseline smooth casing configuration, the successful contouring

associated with such a casing treatment.

from Müller et al., 2008)

stall margin.

induced a smaller inflow angle near the leading edge, i.e. a lower incidence, delaying rotating stall inception.

Fig. 11. Endwall casing contouring (Ito et al., 2007)

#### **4.2 Slot-type treatments**

Another way to treat the casing with the aim to improve the rotor performance and stability is exemplified in Fig. 12. Here the casing wall is circumferentially treated with a discrete number of axial rectangular slots over the blade tip section.

A similar slot-type casing treatment consisting of four identical axial slots per blade passage and having an open area of 50% in the circumferential direction has been also proposed. (Wilke and Kau, 2004). In this case, the slots are parallel to the rotation axis and inclined by 45° against the meridional plane in the direction of rotor rotation. The slot shape is designed as a semi-circle. Two configurations have been numerically tested. In configuration 1, the position of the slots is centered above the rotor blade tip reaching from 7.5% to 92.5% chord length. In configuration 2, the slots are moved upstream so that only 25% chord length remains covered by the casing treatment.

Fig. 12. EJ200 LPC with axial slots casing treatment (Broichhausen & Ziegler, 2005)

State-of-Art of Transonic Axial Compressors 37

incidence and blade loading at the tip, allowing increased loading at lower blade spans before the blade stalls. With tip injection present, the blade stalls when the loading at the tip reaches the level equal to that for which the blade stalls with no injection. The experiments showed that stability enhancement is related to the mass averaged axial velocity at the tip. For the tested rotor, experimental results demonstrated that at 70% speed the stalling flow coefficient can be reduced by 30% using an injected mass flow equivalent to 1% of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6% using an injected mass flow equivalent to 2% of the annulus flow. Tip injection has also been demonstrated as an effective tool for recovering a compressor from fully developed stall. In a self-induced (passive) solution, steady discrete tip injection using casing recirculation has been simulated in a single transonic fan rotor (Hathaway, 2002). The idea was to bleed pressurized fluid from downstream of the rotor and properly inject it upstream. Simulations were carried out assuming both a clear and distorted inlet flow. The distortion was circumferentially applied near the casing endwall fixing a lower inlet total pressure. The recirculation model gave 125% range extension without inlet distortion and 225% range

extension with inlet distortion, without significant impact on overall efficiency.

the multistage compressor.

**5.2 Bleeding** 

design mass flow.

Such a solution has been successively tested on a single-stage transonic compressor (Strazisar et al., 2004). Results clearly indicated that recirculation extends the stable operating range of that stage to lower mass flows than that can be achieved without recirculation. With the recirculation activated, the positive change in stalling flow coefficient was 6% at 70% of design speed and 2% at design speed, with a total injected flow of 0.9% of the annulus flow at both operating speeds. In the same work, the potential for using endwall recirculation to increase the stability of transonic highly loaded multistage compressors was demonstrated through results from a rig test of simulated recirculation driving both a steady injected flow and an unsteady injected flow. Unsteady injection increased stability more than steady injection and was capable of changing the unsteady near stall dynamics of

Fig. 14 shows the sketch of a transonic aspirated stage experimentally tested and numerically investigated to demonstrate the application of boundary layer aspiration for increasing the stage work (Schuler et al., 2005). The stage was designed to produce a pressure ratio of 1.6 at a tip speed of 750 ft/s resulting in a stage work coefficient of 0.88. The primary aspiration requirement for the stage was a bleed fraction 0.5% of the inlet mass flow on the rotor and stator suction surfaces. Additional aspiration totalling 2.8% was also used at shock impingement locations and other locations on the hub and casing walls. The stage achieved a peak pressure ratio of 1.58 and through flow efficiency of 90% at the design point. The rotor showed an extremely high efficiency of 97% for a transonic rotor, partially attributed to aspiration and partially to the elimination of the tip clearance flow due to the tip shroud. Aspiration was also effective in maintaining stage performance at off-design conditions. The experimental data showed unstalled stage performance at least 83% of the

The possibility of a very high pressure ratio per single-stage using aspiration has been demonstrated for the fan of Fig. 15. The fan stage has been designed to achieve a pressure ratio of 3.4 at 1500 ft/s. The low energy viscous flow was aspirated from diffusion-limiting locations on the blades and flow path surfaces. Experimental results gave a stage pressure

For both configurations, simulations showed a significant increment in flow stability compared to the solid wall, the stalling mass flow passed from 0.95% to 0.75% the design mass flow at the design speed. It was observed that the stabilizing effects are based on the positive impact of casing treatment on the tip clearance flow and its resulting vortex. Configuration 1 led to a massive destruction of the tip leakage vortex, whereas configuration 2 weakened the rolling-up of the tip clearance flow. Configuration 2, gave also a positive impact on the overall efficiency.

## **5. Air injection and bleeding**

One of the first flow control ideas to receive a considerable attention in gas turbine applications is the flow injection and bleeding concept. Air injection near the blade tips has proved to increase compressor stall margin, leading to higher engine operability. Fig. 13 shows a set of successful 12 non-intrusive prototype injectors recently installed in the casing of a transonic compressor test rotor. On the other hand, aspiration (or bleeding) can be used to delay blade separation, which limits the stage work and therefore increases the required number of stages. Doubling the work per stage using aspiration results in a dramatic reduction in the number of stages, though not necessarily an exactly proportional reduction in compressor weight due to the added complexity of blades and added plumbing. For fighter applications this technology looks very attractive due to its potential to improve the thrust/weight ratio of the engine.

Fig. 13. Discrete tip injection (Hathaway, 2007)

#### **5.1 Air injection**

Measurements and simulations for active discrete tip injection have been presented for a range of steady injection rates and distributions of injectors around the annulus of a tipcritical transonic compressor rotor operating in isolation (Suder et al., 2001). In such case, casing-mounted injectors were located at 200% tip axial chord upstream of the rotor. Each injector penetrated 6% span from the casing into the flow field. The simulations indicated that tip injection increases stability by unloading the rotor tip. Tip injection decreases

For both configurations, simulations showed a significant increment in flow stability compared to the solid wall, the stalling mass flow passed from 0.95% to 0.75% the design mass flow at the design speed. It was observed that the stabilizing effects are based on the positive impact of casing treatment on the tip clearance flow and its resulting vortex. Configuration 1 led to a massive destruction of the tip leakage vortex, whereas configuration 2 weakened the rolling-up of the tip clearance flow. Configuration 2, gave also

One of the first flow control ideas to receive a considerable attention in gas turbine applications is the flow injection and bleeding concept. Air injection near the blade tips has proved to increase compressor stall margin, leading to higher engine operability. Fig. 13 shows a set of successful 12 non-intrusive prototype injectors recently installed in the casing of a transonic compressor test rotor. On the other hand, aspiration (or bleeding) can be used to delay blade separation, which limits the stage work and therefore increases the required number of stages. Doubling the work per stage using aspiration results in a dramatic reduction in the number of stages, though not necessarily an exactly proportional reduction in compressor weight due to the added complexity of blades and added plumbing. For fighter applications this technology looks very attractive due to its potential to improve the

Measurements and simulations for active discrete tip injection have been presented for a range of steady injection rates and distributions of injectors around the annulus of a tipcritical transonic compressor rotor operating in isolation (Suder et al., 2001). In such case, casing-mounted injectors were located at 200% tip axial chord upstream of the rotor. Each injector penetrated 6% span from the casing into the flow field. The simulations indicated that tip injection increases stability by unloading the rotor tip. Tip injection decreases

a positive impact on the overall efficiency.

**5. Air injection and bleeding** 

thrust/weight ratio of the engine.

Fig. 13. Discrete tip injection (Hathaway, 2007)

**5.1 Air injection** 

incidence and blade loading at the tip, allowing increased loading at lower blade spans before the blade stalls. With tip injection present, the blade stalls when the loading at the tip reaches the level equal to that for which the blade stalls with no injection. The experiments showed that stability enhancement is related to the mass averaged axial velocity at the tip. For the tested rotor, experimental results demonstrated that at 70% speed the stalling flow coefficient can be reduced by 30% using an injected mass flow equivalent to 1% of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6% using an injected mass flow equivalent to 2% of the annulus flow. Tip injection has also been demonstrated as an effective tool for recovering a compressor from fully developed stall.

In a self-induced (passive) solution, steady discrete tip injection using casing recirculation has been simulated in a single transonic fan rotor (Hathaway, 2002). The idea was to bleed pressurized fluid from downstream of the rotor and properly inject it upstream. Simulations were carried out assuming both a clear and distorted inlet flow. The distortion was circumferentially applied near the casing endwall fixing a lower inlet total pressure. The recirculation model gave 125% range extension without inlet distortion and 225% range extension with inlet distortion, without significant impact on overall efficiency.

Such a solution has been successively tested on a single-stage transonic compressor (Strazisar et al., 2004). Results clearly indicated that recirculation extends the stable operating range of that stage to lower mass flows than that can be achieved without recirculation. With the recirculation activated, the positive change in stalling flow coefficient was 6% at 70% of design speed and 2% at design speed, with a total injected flow of 0.9% of the annulus flow at both operating speeds. In the same work, the potential for using endwall recirculation to increase the stability of transonic highly loaded multistage compressors was demonstrated through results from a rig test of simulated recirculation driving both a steady injected flow and an unsteady injected flow. Unsteady injection increased stability more than steady injection and was capable of changing the unsteady near stall dynamics of the multistage compressor.

#### **5.2 Bleeding**

Fig. 14 shows the sketch of a transonic aspirated stage experimentally tested and numerically investigated to demonstrate the application of boundary layer aspiration for increasing the stage work (Schuler et al., 2005). The stage was designed to produce a pressure ratio of 1.6 at a tip speed of 750 ft/s resulting in a stage work coefficient of 0.88. The primary aspiration requirement for the stage was a bleed fraction 0.5% of the inlet mass flow on the rotor and stator suction surfaces. Additional aspiration totalling 2.8% was also used at shock impingement locations and other locations on the hub and casing walls. The stage achieved a peak pressure ratio of 1.58 and through flow efficiency of 90% at the design point. The rotor showed an extremely high efficiency of 97% for a transonic rotor, partially attributed to aspiration and partially to the elimination of the tip clearance flow due to the tip shroud. Aspiration was also effective in maintaining stage performance at off-design conditions. The experimental data showed unstalled stage performance at least 83% of the design mass flow.

The possibility of a very high pressure ratio per single-stage using aspiration has been demonstrated for the fan of Fig. 15. The fan stage has been designed to achieve a pressure ratio of 3.4 at 1500 ft/s. The low energy viscous flow was aspirated from diffusion-limiting locations on the blades and flow path surfaces. Experimental results gave a stage pressure

State-of-Art of Transonic Axial Compressors 39

flow suction. It was also shown that the constant speed throttling characteristic of the new design can be tailored by using varying amounts of suction (up to 2%) to hold the passage shock stationary thereby making it possible to satisfy multiple operating conditions with

Transonic compressors are the state-of-art in the compression system of today's civil and military aero engines. Their capability to provide high pressure ratios maintaining high efficiencies made them preferable to both subsonic (lower pressure ratio) and supersonic

The researches of last decades have greatly contributed to their success. The progress made in optical measurement techniques and the development of computational methods led to a deep understanding of the loss mechanisms associated with their inherent supersonic relative flow, making it possible significant aerodynamic improvements and industrial

Nowadays there is still space for further radical improvements and new solutions will be developed in the next future, like highly effective stall/surge control systems and higher pressure ratio configurations, or even new concepts based on new advanced working

Ahn, C.-S. & Kim, K.-Y. (2002). Aerodynamic Design Optimization of an Axial Flow Compressor Rotor, *Proceedings of ASME Turbo Expo 2002*, GT-2002-30445 Beheshti, B. H.; Teixeira, J. A.; Ivey, P. C.; Ghorbanian, K. & Farhanieh, B. (2004). Parametric

Benini, E. & Biollo, R. (2008). Effect of Forward and Aft Lean on the Performance of a

Bergner, J.; Hennecke, D. K. & Hah, C. (2005a). Tip-Clearance Variations of an Axial High-

Bergner, J.; Kablitz, S.; Hennecke, D. K.; Passrucker, H. & Steinhardt, E. (2005b). Influence of

Biollo, R. & Benini, E. (2008a). Impact of Sweep and Lean on the Aerodynamic Behaviour of

Biollo, R. & Benini, E. (2008b). Aerodynamic Behaviour of a Novel Three-Dimensional

No. 1, (2008), pp. 13-26, ISSN 0334-0082

*Breathing Engines*, ISABE-2005-1096

*ASME Turbo Expo 2005*, GT2005-68835

*4th International Conference*, IGTC08\_P17

Study of Tip Clearance - Casing Treatment on Performance and Stability of a Transonic Axial Compressor, *Proceedings of ASME Turbo Expo 2004*, GT2004-53390 Benini, E. & Biollo, R. (2007). Aerodynamics of Swept and Leaned Transonic Compressor-

Rotors, *Applied Energy*, Vol. 84, No. 10, (October 2007), pp. 1012-1027, ISSN 0306-

Transonic Compressor Rotor, *International Journal of Turbo and Jet Engines*, Vol. 25,

Speed Single-Stage Transonic Compressor, *Proceedings of 17th Symposium on Air-*

Sweep on the 3D Shock Structure in an Axial Transonic Compressor, *Proceedings of* 

Transonic Compressor Rotors, *Proceedings of The Future of Gas Turbine Technology -* 

Shaped Transonic Compressor Rotor Blade, *Proceedings of ASME Turbo Expo 2008*,

minimal aerodynamic penalty.

(lower efficiency) solutions.

**6. Conclusions** 

diffusion.

principles.

**7. References** 

2619

GT2008-51397

ratio exceeding 3 at design speed, with an aspiration flow fraction of 3.5% of the stage inlet mass flow. CFD simulations showed that aspiration fixes the passage shock position, particularly in the tip region, maintaining good aerodynamic behaviour at off-design operating points.

Fig. 14. A tested suction configuration (Schuler et al., 2002)

Fig. 15. Test package (left) and rotor aspiration scheme (right) (Merchant et al., 2005)

In a previous work (Dang & Van Rooij, 2003), it was demonstrated the possibility to reduce the amount of aspiration without compromise its benefits. When used as an add-on to an existing design, aspiration can require excessive amounts of suction, whereas with a blade/ aspiration scheme coupled design a significantly lower amount of suction can be needed. A baseline supersonic rotor with 56 blades was used in that work as the starting point. The blade count in that rotor was then reduced to 42 blades, and it was found that 2% of flow suction was needed to pull the shock back into the blade passage for the required back pressure. The aspiration zone was located approximately behind the shock on the suction surface from hub to 95% span. With 42 blades, inspection of the flow field showed that the peak Mach number and loading were significantly higher than in the baseline blade case, resulting in increased shock losses, thickening of the blade suction surface boundary layer, and a large region of low momentum fluid in the tip clearance region.

A new pressure loading shape was developed to mitigate the situation. The new design was shown to have better efficiency potential and a cleaner casing endwall flow using only 0.3% flow suction. It was also shown that the constant speed throttling characteristic of the new design can be tailored by using varying amounts of suction (up to 2%) to hold the passage shock stationary thereby making it possible to satisfy multiple operating conditions with minimal aerodynamic penalty.

## **6. Conclusions**

38 Advances in Gas Turbine Technology

ratio exceeding 3 at design speed, with an aspiration flow fraction of 3.5% of the stage inlet mass flow. CFD simulations showed that aspiration fixes the passage shock position, particularly in the tip region, maintaining good aerodynamic behaviour at off-design

Fig. 14. A tested suction configuration (Schuler et al., 2002)

Fig. 15. Test package (left) and rotor aspiration scheme (right) (Merchant et al., 2005)

and a large region of low momentum fluid in the tip clearance region.

In a previous work (Dang & Van Rooij, 2003), it was demonstrated the possibility to reduce the amount of aspiration without compromise its benefits. When used as an add-on to an existing design, aspiration can require excessive amounts of suction, whereas with a blade/ aspiration scheme coupled design a significantly lower amount of suction can be needed. A baseline supersonic rotor with 56 blades was used in that work as the starting point. The blade count in that rotor was then reduced to 42 blades, and it was found that 2% of flow suction was needed to pull the shock back into the blade passage for the required back pressure. The aspiration zone was located approximately behind the shock on the suction surface from hub to 95% span. With 42 blades, inspection of the flow field showed that the peak Mach number and loading were significantly higher than in the baseline blade case, resulting in increased shock losses, thickening of the blade suction surface boundary layer,

A new pressure loading shape was developed to mitigate the situation. The new design was shown to have better efficiency potential and a cleaner casing endwall flow using only 0.3%

operating points.

Transonic compressors are the state-of-art in the compression system of today's civil and military aero engines. Their capability to provide high pressure ratios maintaining high efficiencies made them preferable to both subsonic (lower pressure ratio) and supersonic (lower efficiency) solutions.

The researches of last decades have greatly contributed to their success. The progress made in optical measurement techniques and the development of computational methods led to a deep understanding of the loss mechanisms associated with their inherent supersonic relative flow, making it possible significant aerodynamic improvements and industrial diffusion.

Nowadays there is still space for further radical improvements and new solutions will be developed in the next future, like highly effective stall/surge control systems and higher pressure ratio configurations, or even new concepts based on new advanced working principles.

## **7. References**


State-of-Art of Transonic Axial Compressors 41

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**3** 

Marek Dzida

*Poland* 

 *Gdansk University of Technology* 

**Possible Efficiency Increasing of Ship** 

**Gas Turbine and Steam Turbine** 

**Propulsion and Marine Power Plant with** 

**the System Combined of Marine Diesel Engine,** 

For years there has been, and still is, a tendency in the national economy to increase the efficiency of both the marine and inland propulsion systems. It is driven by economic motivations (rapid increase of fuel prices) and ecological aspects (the lower the fuel consumption, the lower the emission of noxious substances to the atmosphere). New design solutions are searched to increase the efficiency of the propulsion system via linking Diesel engines with other heat engines, such as gas and steam turbines. The combined systems implemented in marine propulsion systems in recent years are based mainly on gas and steam turbines (MAN, 2010). These systems can reach the efficiency exceeding 60% in inland applications. The first marine system of this type was applied on the passenger liner "Millenium". However, this is the only high-efficiency marine application of the combined propulsion system so far. Its disadvantage is that the system needs more expensive fuel, the marine Diesel oil, while the overwhelming majority of the merchant ships are driven by low-speed engines fed with relatively cheap heavy fuel oil. It seems that the above tendency

The compression-ignition engine (Diesel engine) is still most frequently used as the main engine in marine applications. It burns the cheapest heavy fuel oil and reveals the highest efficiency of all heat engines. The exhaust gas leaving the Diesel engine contains huge energy which can be utilised in another device (engine), thus increasing the efficiency of the

A possible solution here can be a system combined of a piston internal combustion engine and the gas and steam turbine circuit that utilises the heat contained in the exhaust gas from the Diesel engine. The leading engine in this system is the piston internal combustion engine. It seems that now, when fast container ships with transporting capacity of 8- 12 thousand TU are entering into service, the propulsion engines require very large power, exceeding 50-80 MW. On the other hand, increasing prices of fuel and restrictive ecological limits concerning the emission of NOx and CO2 to the atmosphere provoke the search for new solutions which will increase the efficiency of the propulsion and reduce the emission

The ship main engines will be large low-speed piston engines that burn heavy fuel oil. At present, the efficiency of these engines nears 45 – 50%. For such a large power output

entire system and reducing the emission of noxious substances to the atmosphere.

will continue in the world's merchant navy for the next couple of years.

**1. Introduction** 

of gases to the atmosphere.

