Conceptual Design Study of a Hydrogen Powered Ultra Large Cargo Aircraft

R.A.J. Jansen University of Technology Technology of University Delft Delft

Towards a competitive and sustainable alternative of maritime transport

R.A.J. Jansen

to obtain the degree of Master of Science at the Delft University of Technology, to be defended publicly on Tuesday January 10, 2017 at 9:00 AM.

Student number: 4036093 Thesis registration: 109#17#MT#FPP Project duration: January 11, 2016 – January 10, 2017 Thesis committee: Dr. ir. G. La Rocca, TU Delft, supervisor Dr. A. Gangoli Rao, TU Delft Dr. ir. H. G. Visser, TU Delft

An electronic version of this thesis is available at http://repository.tudelft.nl/.

Acknowledgements

This report presents the research performed to complete the master track Flight Performance and Propulsion at the Technical University of Delft. I am really grateful to the people who supported me both during the master thesis as well as during the rest of my student life. First of all, I would like to thank my supervisor, Gianfranco La Rocca. He supported and motivated me during the entire graduation project and provided valuable feedback during all the status meeting we had. I would also like to thank the exam committee, Arvind Gangoli Rao and Dries Visser, for their ﬂexibility and time to assess my work. Moreover, I would like to thank Ali Elham for his advice throughout the project as well as during the green light meeting. Next to these people, I owe also thanks to the fellow students in room 2.44 for both their advice, as well as the enjoyable chats during the lunch and coffee breaks. Special thanks to Harm, Frederick, and Raoul who were always available to think about possible solutions when I was stuck. Last but not least, special thanks to my parents, sister and rest of the family for their inexhaustible support during my entire life. R.A.J. Jansen December 23, 2016

iii

Summary

The market of the intercontinental transportation of containerized goods is dominated by the maritime sec- tor due to its low transportation costs. However, the current cargo ships significantly contribute to the environmental pollution due to the emission of greenhouse gases. Therefore, a research has been initiated to investigate the feasibility of a hydrogen powered ultra large cargo aircraft as a competitive and sustainable alternative to maritime transport. To gain at least 6-8% market share, a freight rate of less than 250% of the freight rate of current container ships is the estimated performance target for this aircraft concept. Due to its ability to transport 100 lightweight standardized containers, significant transportation cost savings are expected compared to current air cargo transport. Furthermore, the emission of the greenhouse gases is significantly reduced by making use of hydrogen instead of kerosene as aviation fuel. The objective and scope of this research is to investigate the design and performance of this new aircraft concept by performing a conceptual design study, where the competitiveness of this aircraft concept with respect to maritime transport is evaluated by an operational and economical study performed in parallel to this research. Two unconventional aircraft configurations, the multifuselage configuration and the blended- wing-body configuration, are investigated for this aircraft concept, because of its efficient storage of the huge amount of containers and the large pressurized fuel tanks required to store hydrogen compared to a conventional aircraft configuration. To make a quantitative trade-off between the two concepts for this application, a conceptual design framework has been developed to support the synthesis of a conceptual design for both aircraft concepts. A multidisciplinary optimization approach has been applied to integrate the mutual interaction between multiple disciplines for the design and analysis of both aircraft concepts already in the conceptual design phase. The conceptual design framework makes use of semi-empirical and quasi-analytical methods, because the simple and fast statistical methods are unreliable due to the significant differences of both aircraft concepts with respect to existing aircraft concepts. Although the level of design accuracy is relatively low, it is possible to compare the two proposed aircraft concepts inside a relatively large design space by making use of this conceptual design framework. Designed for the same top level requirements, it was found that both aircraft concepts have a wing span larger than 200 meters. Due to these extreme large aircraft dimensions and excessive payload weight requirement, this aircraft concept is not able to operate on existing airports. Therefore, large investment costs are required to build new airports and hydrogen facilities for the aircraft operations of this new concept. In terms of transport efficiency, the blended-wing-body concept requires 8.2% less fuel weight compared to the multifuselage concept. This reduction is mainly caused by a higher aerodynamic efficiency and a lower operational empty weight of the blended-wing-body concept because of the absence of the heavy fuselages compared to the multifuselage concept. Compared to current large cargo aircraft, the multifuselage concept performs slightly better (1.4%) in terms of transport efficiency, whereas the blended-wing-body concept is 8.6% more efficient. However, the transport efficiency of cargo ships is still around 50 times higher with respect to a hydrogen powered ultra large cargo aircraft. On the other hand, the transportation time of this aircraft concept is around 20 times faster than maritime transport for intercontinental transport of containerized goods. Based on these results and a sensitivity study to investigate the influence of TLRs on the aircraft design and performance, a quantitative trade-off has been performed. Designed for the initial top level requirements set at the start of this research, it was found that the BWB concept is preferred over the multifuselage concept because of the smaller aircraft dimensions, the lower aircraft weight, and the higher transport efficiency. On the contrary, the multifuselage concept has the potential to become more transport efficient compared to the BWB concept for a higher cruise speed. Therefore, further research is required to reassess the top level requirements and analyze both aircraft concepts is more detail in order to complete the conceptual design study and select a baseline design to enter the preliminary design phase.

Contents

List of Figures ix List of Tables xi Nomenclature xiii 1 Introduction 1 1.1 Motivation of a ultra large cargo aircraft...... 1 1.2 Motivation of hydrogen as aviation fuel ...... 2 1.3 Research scope, objective and approach ...... 3 1.4 Top level requirements ...... 3 1.5 Report structure...... 4 2 Aircraft Concepts Definition 5 2.1 Identification of aircraft configurations ...... 5 2.1.1 Multifuselage configuration ...... 5 2.1.2 All-lifting configuration ...... 6 2.1.3 Discussion ...... 8 2.2 Identification of propulsion system ...... 8 2.2.1 Gas turbine engines ...... 9 2.2.2 Fuel cell ...... 9 2.2.3 Discussion ...... 10 2.3 Definition of ultra large cargo aircraft concepts ...... 10 2.3.1 Multifuselage concept ...... 10 2.3.2 Blended-wing-body concept ...... 12 3 Conceptual Design Methodology 13 3.1 Design approach ...... 13 3.2 Development of a MDO framework for conceptual design ...... 15 3.2.1 Design of experiments ...... 17 3.2.2 Numerical optimization ...... 17 3.2.3 Implementation ...... 18 3.3 Multidisciplinary analysis ...... 18 3.3.1 Class I weight estimation...... 19 3.3.2 Wing power loading diagram...... 21 3.3.3 Geometric modeller ...... 21 3.3.4 Analysis of operative empty weight ...... 26 3.3.5 Center of gravity ...... 28 3.3.6 Aerodynamic analysis ...... 30 3.3.7 Maximum lift coefficient ...... 32 3.3.8 Performance analysis ...... 33 3.3.9 Constraint analysis...... 34 3.4 Set-up of numerical optimization ...... 35 3.4.1 Formulation of optimization problem ...... 36 3.4.2 Optimization algorithm ...... 36 3.4.3 Implementation ...... 37 4 Aircraft Design and Performance 39 4.1 Multifuselage concept ...... 39 4.2 Blended-wing-body concept ...... 45 4.3 Design and performance comparison between aircraft concepts ...... 52 4.3.1 Aircraft dimensions ...... 53

vii viii Contents

4.3.2 Aircraft weight ...... 53 4.3.3 Aerodynamic performance...... 54 4.3.4 Stability and controllability ...... 54 4.3.5 Transport efﬁciency ...... 54 4.3.6 Conclusions ...... 55 4.4 Design and performance comparison to aircraft concept of DSE ...... 58 5 Sensitivity Analysis 59 5.1 Multifuselage concept ...... 59 5.2 Blended-wing-body concept ...... 60 5.3 Conclusions...... 61 6 Conceptual Design Trade-Off 65 6.1 Trade-off criteria and weights ...... 65 6.2 Quantitative trade-off table ...... 66 6.3 Discussion ...... 66 7 Conclusions 69 7.1 Potential aircraft concepts ...... 69 7.2 Conceptual design framework ...... 69 7.3 Aircraft design and performance ...... 70 7.4 Sensitivity analysis ...... 71 7.5 Conceptual design trade-off ...... 71 8 Recommendations 73 Bibliography 77 A User manual 81 B Database of large cargo aircraft 83 C Veriﬁcation of wing power loading diagram 85 D Final report of multifuselage concept 87 E Final report of blended-wing-body concept 95 List of Figures

1.1 Main characteristics of lightweight container ...... 2

2.1 Distribution along the span of lift, mass and bending moment for a twin-fuselage compared to a conventional aircraft ...... 6 2.2 Illustration of two lifting fuselage configurations...... 6 2.3 Illustration of different aircraft configurations of the all-lifting configuration...... 7 2.4 Comparison of lift and weight distribution of a BWB design and a conventional aircraft design . 7 2.5 Examples of contra-rotating propeller system used by current commercial aircraft...... 9 2.6 Illustration of the multifuselage and BWB concept considered for this study...... 11 2.7 Overview of the placement of the containers and the fuel tank into the fuselage for the multifuselage concept...... 11 2.8 Characteristics of wing planform of the BWB concept considered for this research...... 12

3.1 A comparison between the traditional design approach and the multidisciplinary optimization design approach, which is indicated by the dotted line ...... 14 3.2 Architecture of MDO framework for conceptual design...... 16 3.3 Architecture of MDA system...... 18 3.4 Mission profile of the aircraft...... 19 3.5 Regression line based on statistical data of existing large cargo aircraft and the Pelican Ultra . .. 20 3.6 Wing power loading diagram for the Lockheed Super Hercules C130J...... 21 3.7 Parametrization of the fuselage and the wing planform for the multifuselage concept...... 23 3.8 Parametrization of the BWB concept...... 25 3.9 Flow diagram of the iterative sizing process of the planform for the BWB concept...... 25 3.10 Illustration of the methodology to determine the container distribution along the wing span for the BWB concept...... 26 3.11 Cross-section of the wing to illustrate the placement of the containers and the fuel tank...... 26 3.12 Relationship between the tank diameter and the structural tank weight and insulation weight . . 27 3.13 Placement of the landing gear with respect to the center of gravity positions based on the design rules of the handbook of Raymer ...... 30 3.14 Clearance requirements for the calculation of the landing gear height...... 30 3.15 Drag break-down for a low subsonic aircraft...... 31 3.16 Streamlines over a BWB configuration determined by a full panel method ...... 31 3.17 Increase in chord extension as a function of the flap deflection and type of high-lift device . ... 34

4.1 Surrogate models of the objective function and constraint functions for multifuselage concept. 41 4.2 A three-dimensional overview of the multifuselage configuration...... 42 4.3 Wing power loading diagram of the multifuselage concept...... 42 4.4 Lift distribution along the wing span of the multifuselage concept...... 44 4.5 Airfoil lift coefficient as a function of the angle of attack for the multifuselage concept...... 44 4.6 Surrogate models on objective function and constraint functions for BWB concept...... 47 4.7 A three-dimensional overview of BWB concept...... 48 4.8 Wing power loading diagram of the BWB concept...... 49 4.9 Overview of the container distribution...... 49 4.10 Cut-outs of the aircraft to show the placement of the fuel tank and the containers, starting from the center line towards the wing tip...... 50 4.11 Lift distribution along the wing span of the BWB concept...... 52 4.12 Airfoil lift coefficient as a function of the angle of attack for the BWB concept...... 52 4.13 Illustration of geometric characteristics of both concepts...... 53

ix x List of Figures

5.1 Results of sensitivity study of multifuselage concept...... 62 5.2 Results of sensitivity study of BWB concept...... 63

7.1 Illustration of the geometric difference between the multifuselage, BWB concept and the Antonov An-225...... 70

A.1 Overview of the folder structure for the BWB concept...... 82

C.1 Wing power loading diagram for the Lockheed Super Hercules C130J...... 86

D.1 Aircraft geometry (all dimensions in meters) ...... 87 D.2 Wing Power Loading Diagram ...... 88 D.3 Mass Distribution ...... 89 D.4 Lift distribution along the wing span ...... 91 D.5 C Æ curve ...... 91 l ° D.6 Plot of the fuselage including container and fuel tank...... 92

E.1 Aircraft geometry (all dimensions in meters) ...... 95 E.2 Wing Power Loading Diagram ...... 96 E.3 Mass Distribution ...... 97 E.4 Lift distribution along the wing span ...... 98 E.5 C Æ curve ...... 99 l ° E.6 Overview of the placement of containers ...... 100 E.7 Cut-outs of the aircraft to show the placement of the fuel tank and the containers, starting from the centerline towards the wing tip...... 101 List of Tables

1.1 Main difference in properties of liquid hydrogen and kerosene as aviation fuel ...... 3

3.1 Description of the symbols of the XDSM for the MDO framework...... 15 3.2 Fuel fraction of the different mission segments ...... 19 3.3 Input parameters for Class I weight estimation provided by the handbook of Raymer ...... 20 3.4 Input vector of multifuselage configuration for the geometric model...... 22 3.5 Input vector for the geometric model of BWB configuration...... 24 3.6 Material properties of aluminum for the wing box weight estimation by EMWET...... 28 3.7 Center of gravity location of the main components ...... 29 3.8 Interference factor (Q) of the different aircraft components provided by Raymer ...... 32 3.9 Design vector of the multifuselage configuration including bounds...... 36 3.10 Design vector of the BWB configuration including bounds...... 36 3.11 Parameters selected for the construction of the surrogate models...... 37

4.1 Initial design vector of multifuselage concept for numerical optimization...... 40 4.2 Final design vector of multifuselage concept for numerical optimization...... 40 4.3 Weight breakdown of the multifuselage concept...... 43 4.4 Aerodynamic properties of the multifuselage concept...... 43 4.5 Stability and controllability of the multifuselage concept...... 44 4.6 Initial design vector of BWB concept for numerical optimization...... 45 4.7 Final design vector of multifuselage concept for numerical optimization...... 46 4.8 Weight breakdown of the BWB concept...... 51 4.9 Aerodynamic properties of the BWB concept ...... 51 4.10 Stability and controllability of the BWB concept...... 52 4.11 Comparison of different transportation modes in terms of PFEE...... 55 4.12 Total sailing time for a container ship between Singapore and Rotterdam ...... 55 4.13 Performance overview of BWB concept and multifuselage concept ...... 57 4.14 Comparison between aircraft concept of DSE project and BWB concept considered for this study 58

5.1 Parameters selected for the sensitivity study including the perturbations with respect to the baseline TLRs...... 59 5.2 Results of sensitivity study of multifuselage concept in percentages with respect to baseline design...... 60 5.3 Results of sensitivity study of BWB concepts in percentages with respect to baseline design. ... 60

6.1 Trade-off table to quantitatively compare the two aircraft concepts...... 67

7.1 Performance overview of BWB concept and multifuselage concept ...... 71

B.1 Database of OEW and MTOW for different aircraft (in kilograms)...... 83

C.1 Speciﬁcations of Lockheed Super Hercules C130J based on Jane’s all the world’s aircraft and the handbook of Raymer ...... 85

D.1 Speciﬁcation Mission Requirements ...... 88 D.2 Performance results ...... 88 D.3 Mass summary ...... 89 D.4 Masses of aircraft components ...... 89 D.5 Location of neutral point and CG ...... 90 D.6 Aerodynamic properties at cruise ...... 90

xi xii List of Tables

D.7 Characteristics of propulsion system...... 91 D.8 Main Wing ...... 92 D.9 Fuselage ...... 92 D.10 Horizontal stabilizer ...... 92 D.11 Vertical Stabilizer ...... 92 D.12 Final Overview ...... 93

E.1 Speciﬁcation Mission Requirements ...... 96 E.2 Performance results ...... 96 E.3 Mass summary ...... 97 E.4 Masses of aircraft components ...... 97 E.5 Location of neutral point and CG as a percentage of MAC ...... 98 E.6 Aerodynamic properties at cruise ...... 98 E.7 Characteristics of propulsion system...... 99 E.8 Main Wing ...... 99 E.9 Vertical Stabilizer ...... 99 E.10 Final Overview ...... 102 Nomenclature

Latin symbols

Symbol Description Unit A Area m2 AR Aspect ratio - b Wing span m c Chord length m CD Drag coefficient - C f Skin friction coefficient - CL Aircraft lift coefficient - Cl Airfoil lift coefficient - Cp Specific fuel consumption kg/J crate Climb rate m/s D Drag force N d Diameter m e Oswald efficiency factor - f Objective function - h Altitude m L Lift force N l Length m M Mach number - M ff Overall fuel fraction - n Number - Q Interference factor - R Range m Re Reynolds number - S Surface area m2 s Distance m t Thickness m T Thrust N V Cruise speed m/s W Weight kg w Weight factor - x Longitudinal distance m y Lateral distance m

Greek symbols

Symbol Description Unit Æ Angle of attack ° Ø Side slip angle ° ± Deﬂection ° ¥ Efﬁciency - § Sweep angle ° ∏ Taper ratio - Ω Air density kg/m3

xiii xiv 0. Nomenclature

Subscripts

Symbol Description ac Aerodynamic center avg Average cg Center of gravity cont Containers des Design fe Fixed equipment ft Fuel tank fus Fuselage HT Horizontal tail i Induced iwing Inner wing LE Leading edge lg Landing gear nac Nacelle OEI One engine inoperative owing Outer wing p Propulsive pp Power plant ref Reference req Required TO Take-off VT Vertical tail wet Wetted Introduction

Nowadays, different transportation modes are used for the world wide transportation of all types of commodities. Examples of transportation modes are the train, truck, ship or airplane. Regarding intercontinental distances, there are only two modes of transportation: maritime transport and air cargo transport. The main benefit of maritime transport is the low transportation cost, although the transportation process is slow. On the other hand, air cargo transport is faster and more reliable but more expensive compared to maritime transport [1]. Therefore, high value and time-sensitive goods are in general transported by cargo aircraft, where less time-sensitive and bulk goods are transported by maritime freighters. In terms of weight, cargo aircraft only transport 1% of the total goods, where maritime freighters transport 99% of the total goods [1]. A drawback of both transportation modes is the significant contribution to the global environmental pollution due to the emission of greenhouse gases. Because of the large market share of the maritime transport with respect to the intercontinental transportation of goods, the total environmental impact of the maritime transport is quite large compared to the total environmental impact of air transport. Therefore, a research is initiated to investigate the feasibility to introduce a new aircraft concept as a sustainable and competitive alternative to maritime transport for intercontinental transportation of containerized goods: a hydrogen powered ultra large cargo aircraft. This chapter starts with a brief motivation for this new aircraft concept, followed by a clear definition of the research scope, objective and approach. Subsequently, the main Top Level Requirements (TLRs) defined for this new aircraft concept are discussed. Finally, the outline of the report is described.

1.1. Motivation of a ultra large cargo aircraft

In order to increase the market share of air cargo transport with respect to intercontinental cargo transportation, significant cost reductions have to be accomplished. The largest cost reduction could be achieved by increasing the payload capacity of the cargo aircraft [2]. First of all, this should reduce the amount of fuel per kilogram of cargo resulting in a significant transportation cost reduction. Secondly, the transportation costs per kilogram of cargo are further reduced by distributing the fixed costs per flight, such as crew salary and landing fees, over a larger payload weight. Another cost saving could be accomplished by minimizing the ground handling time. Currently, specialized air cargo containers are used to transport air commodities. After a flight, the commodities have to be trans- ferred from these specialized containers into standardized Twenty-foot Equivalent Unit (TEU) containers which are used by ground transportation modes, such as trucks and trains. Therefore, it would be beneficial to make use of lightweight TEU containers to reduce the transfer time between the cargo aircraft and the ground transportation mode, resulting in lower transportation time and costs. An example of a lightweight TEU container including the specifications is given in Figure 1.1. Therefore, an aircraft with a very large payload capacity for the storage of lightweight TEU containers has the potential to reduce the transportation costs of air cargo transportation significantly. Because of the excessive payload capacity compared to conventional cargo aircraft, this aircraft concept is also called an ultra large

1 2 1. Introduction

Figure 1.1: Main characteristics of lightweight container [3]. cargo aircraft. Due to the costs savings discussed in this section, an ultra large cargo aircraft has the potential to become a cost competitive alternative to maritime transport.

1.2. Motivation of hydrogen as aviation fuel

Besides being a cost competitive alternative to maritime transport, the ultra large cargo aircraft has also to become a sustainable alternative. The Advisory Council for Aviation Research and Innovation in Europe (ACARE) has set high goals for the aviation industry to reduce the environmental impact in 2050 [4]. The main goal is to reduce the total amount of carbon dioxide (CO2) and oxides of nitrogen (NOx ) emissions of air transport by 75% and 90%, respectively. Furthermore, ACARE would like to decrease the dependency on fossil fuels as aviation fuel in the future. One promising alternative fuel with respect to the goals of ACARE is hydrogen, because it can be produced from renewable resources and has very low emissions of greenhouse gases [5–8].

Hydrogen is the lightest and most abundant element in the universe [8]. Hydrogen can be stored as a gas, a hybride, a liquid or slush. Slush hydrogen is a gelled liquid hydrogen with a higher density compared to liquid hydrogen [9]. The cost to produce slush is signiﬁcantly higher compared to liquid hydrogen, which makes it not attractive for aircraft application [9]. Next, the volume required for hydrogen stored as a gas or a hybride is too large for aircraft applications due to their low density [10]. Because of these reasons only liquid hydrogen is considered as possible aviation fuel in this report and hereafter hydrogen refers to liquid hydrogen.

In order to store hydrogen and to prevent evaporation, cryogenic conditions are required in the fuel tank, namely a temperature of 20.27 K at a pressure of one standard atmosphere (atm) [9, 10]. To ensure these conditions during the entire mission, the fuel tanks have to be pressurized and well-insulated, which increases the complexity of the structure of the fuel tank, but also the structural weight with respect to a conventional kerosene fuel tank. Because of the pressurization the shape of the fuel tanks are typically cylindrical or spherical [10]. The feasibility of cryogenic storage of hydrogen has already been proven by different space programs [9]. However, a challenge of the fuel tanks for aircraft applications is the longer life time. The cryogenic storage system has to withstand the loads during the entire life time of the aircraft, which is around 20-30 years. Next to that, the system has to be accessible for inspection, maintenance and replacement.

One of the main beneﬁts of hydrogen compared to kerosene is its very low environmental impact due to the absence of most greenhouse gases during combustion [10–12]. The only by-products of the combustion of hydrogen are water vapor (H2O), and oxides of nitrogen (NOx ). Compared to kerosene no carbon dioxide is produced and the emission of NOx is reduced by 1/3 to 1/2 [13]. However, more water vapour is produced by hydrogen powered aircraft compared to kerosene fuelled aircraft. Torenbeek [14] states that around 2.6 times the amount of water vapor is produced by hydrogen powered aircraft compared to kerosene fuelled aircraft for the same thrust level [14]. On the other hand, the impact of water vapor on the global warming effect becomes negligible small for cruise altitudes below 10,000 meters compared to other greenhouse gases [7].

Another beneﬁt of hydrogen compared to kerosene is the high heat of combustion, which is the amount of energy released after the complete combustion of one kilogram of the product. The heat of combustion is almost three times higher for hydrogen compared to kerosene [10]. On the other hand, the density of hydrogen is signiﬁcantly lower for hydrogen compared to kerosene. These main differences between the 1.3. Research scope, objective and approach 3

Table 1.1: Main difference in properties of liquid hydrogen and kerosene as aviation fuel [10].

Property Unit Hydrogen Kerosene Nominal composition - H2 CH1.93 Heat of combustion MJ/kg 120 42.8 Liquid density kg/m3 71 811 Boiling point at 1 atm K 20.27 440-539 two fuels are listed in Table 1.1. Brewer [10] states that a hydrogen powered aircraft requires around four times more fuel volume compared to a kerosene fuelled aircraft despite of the lower fuel weight of a hydrogen powered aircraft due to the high heat of combustion of hydrogen compared to kerosene.

1.3. Research scope, objective and approach

Based on the motivation described in the former sections, a project has been initiated to investigate the feasibility to introduce a hydrogen powered ultra-large cargo aircraft as competitive and sustainable alternative to maritime transport for intercontinental transport of containerized goods. In terms of cost competitiveness, a freight rate of less than 250% of the freight rate of container ships is the estimated target required to gain at least 6-8% market share of the intercontinental transport of containerized goods. In order to investigate the feasibility to achieve this estimated performance target, this project is divided into two thesis assignments:

1. An operational and economical study to investigate the competitiveness of this aircraft concept with respect to maritime transport.

2. A feasibility study to investigate the design and performance of this aircraft concept by performing a conceptual design study.

The research scope and objective of this thesis assignment is to investigate the design and performance of this aircraft concept by performing a conceptual design study. However, there are several interdependencies between the two assignments. For example, the operational and economical study requires speciﬁcations of the aircraft design and performance, such as the fuel consumption, cruise speed, the dimensions and the total weight of the aircraft, to investigate the competitiveness of this aircraft concept with respect to maritime transport in terms of freight rate and environmental impact. On the contrary, the conceptual design study requires a clear requirement for the range, which is determined by the operational and economical study by performing a market analysis to determine the optimal airport allocations for world wide transportation of containerized goods. Therefore, both studies are executed in parallel and intermediate results are exchanged.

To accomplish the research objective of this thesis assignment, the following research approach has been defined. The conceptual design study starts with the identification of promising aircraft concepts with respect to the predefined TLRs. The TLRs for this new aircraft concept are discussed in the next section. Therefore, the first step is to perform a literature study to identify and propose potential aircraft configurations and a propulsion system for this aircraft concept. In order to make a quantitative trade-off between the proposed aircraft concepts, a conceptual design framework has to be developed which is able to design and analyze the different aircraft concepts for the predefined TLRs. Next, a sensitivity analysis is performed to investigate the influence of TLRs (cruise speed, cruise altitude, and payload weight) on the design and performance of the different aircraft concepts. Based on the results of the conceptual design study and the sensitivity analysis, a quantitative trade-off is performed to compare the design and performance of the different aircraft concepts as a competitive and sustainable alternative to maritime transport.

1.4. Top level requirements

The primary aircraft design objective is to minimize the development and operational costs in order to maximize competitiveness with respect to maritime transport and air cargo transport, where a freight rate of less than 250% of the freight rate of container ships is the estimated target required to gain at least 6-8% market 4 1. Introduction share of the intercontinental transport of containerized goods. In order to achieve this performance target, the following TLRs were defined at the start of this project: • The maximum payload weight of the aircraft is 1,200 tonnes. • The aircraft requires a cargo bay capacity of 100 TEU containers. • No pressurization in the cabin is required to transport the commodities in the TEU containers. • Range of 6,000 kilometers at maximum payload weight. • The aircraft has to be powered by hydrogen. • The cruise altitude is 8,000 meters to minimize the environmental impact of the water vapor emission of hydrogen. A market research performed by the operational and economical study has shown that the hydrogen powered ultra large aircraft has to operate along the East-West main lane in order to gain a market share of at least 6- 8% [15]. This route covers the transportation of commodities between China, Europe and USA. A limitation is posed on the type of commodities transported by this new aircraft concept due to design requirement of no pressurization in the cabin. Therefore, explosives and living organisms, such as animals, cannot be transported by this aircraft concept. However, it was found from the market analysis of the operational and economical study that sufficient demand will still remain for this aircraft concept to gain a market share of at least 6-8% [15]. Subsequently, the optimal airport locations for the transportation of containerized goods between China, Europe and USA are determined with respect to range minimization by the operational and economical study. By taking into account a maximum of one intermediate stop required for operational efficiency, the following airport locations are defined: Shanghai region (China), Rotterdam (the Netherlands), Pennsylvania (USA), Novosibirsk (Russia), and Anchorage (USA). The latter two airports are only used for refuelling where the first three airport locations are used as full service airports (loading and unloading containers). It was found that the design range of 6,000 kilometers at maximum payload weight is not sufficient to operate between these airports. The design range at maximum payload is therefore adjusted to 7,400 kilometers where the influence of the wind on the range is also taken into account. A detailed description of the methodology for the airport allocations and the range determination can be found in [15]. The main design challenge of this new aircraft concept is the payload weight requirement. As a comparison, the Antonov An-225 has the world-record of greatest payload carried by an aircraft with a payload weight of 253,820 kilograms 1, where the hydrogen powered ultra-large cargo aircraft has a maximum payload weight requirement of 1,200,000 kilograms. Next to the payload weight requirement, the cargo bay capacity has to be large enough to store 100 TEU containers, where the length of each TEU container is more than 6 meters. Furthermore, sufficient volume has to be available to store the large pressurized cylindrical/spherical hydrogen fuel tanks into the aircraft. Because of these challenging design requirements, no restrictions are defined for the Maximum Take-Off Weight (MTOW) and the aircraft dimensions.

1.5. Report structure This ﬁrst chapter provided an introduction and motivation to the research of a hydrogen powered ultra-large cargo aircraft. The structure of the rest of the report is as follows. The identiﬁcation of the aircraft concepts investigated during this research are discussed in Chapter 2, where the methodology to generate a conceptual design for these aircraft concepts is presented in Chapter 3. Subsequently, the design and the performance of the aircraft concepts are described and compared to each other in Chapter 4. The results of the sensitivity analysis of the payload weight, cruise Mach number and cruise altitude on the design and the performance of the aircraft are discussed in Chapter 5. Next, the quantitative trade-off between the different aircraft concepts based on the results of the conceptual design study and the sensitivity analysis is described in Chapter 6. Finally, conclusions and recommendations for further research are given in Chapter 7 and Chapter 8,respectively.

1The Fédération Aéronautique Internationale, FAI Record ID #7129, http://www.fai.org/fai-record-ﬁle/?recordId=7129 (accessed 12 De- cember 2016) Aircraft Concepts Deﬁnition

This chapter describes the identification of potential suitable aircraft concepts that are able to potentially fulfill the TLRs described in Section 1.4. First, the potential aircraft configurations are described in Section 2.1, followed by a discussion about the propulsion system for this new aircraft concept in Section 2.2. Based on the identification of the potential aircraft configurations and the propulsion system for a hydrogen powered ultra large cargo aircraft, the aircraft concepts considered for this research are defined in Section 2.3.

2.1. Identiﬁcation of aircraft conﬁgurations

As can be concluded from the requirements described in Section 1.4, an aircraft configuration has to be chosen which is able to store large hydrogen fuel tanks and 100 standarized TEU containers efficiently. Regarding the conventional Tube and Wing (TAW) configuration, the payload is placed into one fuselage, whereas the fuel is stored into the wing for kerosene fuelled aircraft. The main benefit of storing fuel into the wing is the bending moment relief of the fuel weight on the wing, resulting in a lower structural wing weight. However, it is not possible to store hydrogen in a conventional wing because of the required large cylindrical/spherical pressurized fuel tanks [10]. As a consequence, both the fuel tanks and the cargo have to be stored into the fuselage. Besides the inefficient storage of the payload and fuel because of the large unused space in the wing, a high bending moment will occur at the root of the wing resulting in a heavy wing. Therefore, a literature study has been performed on unconventional aircraft configurations, which are able to store the fuel tanks and containers more efficiently. Two types of promising unconventional aircraft configurations were found for this aircraft concept: the multifuselage configuration and the all-lifting configuration.

2.1.1. Multifuselage conﬁguration

Instead of storing all the payload into one fuselage, an aircraft can also be designed with two fuselages distributed along the wing span. The main benefit of this configuration is a reduction in the bending moment by placing the fuselages outboard of the aircraft centerline [16, 17], which is illustrated in Figure 2.1. The reduction in bending moment has a positive effect on the performance of the aircraft. First of all, the bending moment relief results in a reduction of the structural wing weight. Next, a higher aspect ratio can be applied, which results in less induced drag and a higher lift-to-drag ratio (L/D). Another benefit of the multifuselage configuration is that the cargo can be loaded/unloaded simultaneously from the nose and the tail of the fuselage, resulting in a reduction of operational time at the ground.

With respect to a hydrogen powered ultra-large cargo aircraft, the fuel tanks also have to be placed into the multiple fuselages because the large cylindrical/spherical pressurized fuel tanks do not fit in the conventional wing of the multifuselage configuration. Therefore, a drawback of this configuration is that the volume in the wings is unused.

5 6 2. Aircraft Concepts Deﬁnition

Figure 2.1: Distribution along the span of lift, mass and bending moment for a twin-fuselage compared to a conventional aircraft [14].

2.1.2. All-lifting conﬁguration

Another potential unconventional aircraft configuration for a hydrogen powered ultra-large cargo aircraft is the all-lifting configuration. This configuration evolved from the inspiration of aircraft designers to have a design which has only components that contribute to lift [18]. This configuration consists of three categories: lifting fuselage configuration, lifting body configuration, and the flying wing configuration. An example for each category is illustrated in Figure 2.2 and Figure 2.3.

Compared to the conventional fuselage which does not significantly contribute to the lift production, the lifting fuselage configuration has a thick aerodynamically shaped body connected to a conventional wing and empennage, where all the payload is stored in the large airfoil shaped lifting body [18]. A well-known example is the Burnelli UB-14, illustrated in Figure 2.2(a) 1. Therefore, this aircraft configuration is also called the Bur- nelli configuration. The main benefit of this configuration compared to the conventional TAW configuration is the increased aerodynamic efficiency, because the fuselage also has a positive lift contribution [18]. Fur- thermore, the payload can be loaded and unloaded easily from a large flat nose. This configuration has been analyzed for the application of a hydrogen powered ultra large cargo aircraft during the 2015 Design Synthesis Exercise (DSE) [19]. An isometric view of this concept is shown in Figure 2.2(b). The main drawback of this aircraft design is that the payload and the fuel tanks are located into the main body, whilst the volume into the wings is unused. As a result, a large bending moment occurs on the wing due to the absence of a bending moment relief of the fuel weight or payload weight.

(a) Burnelli UB-14 (b) Concept of 2015 DSE project [19]

Figure 2.2: Illustration of two lifting fuselage conﬁgurations.

1All the World’s Helicopters and Rotorcraft, Burnelli UB-14, http://www.aviastar.org/air/usa/burnelli_ub-14.php (accessed 22-12-2016) 2.1. Identiﬁcation of aircraft conﬁgurations 7

The lifting body configuration, also called the Blended-Wing-Body (BWB) configuration, consists of thick aerodynamic shapes which are merged smoothly together [2, 18, 20–22]. An illustration of this concept is shown in Figure 2.3(a). Typically, this configuration has vertical surfaces for lateral stability and control, whereas trailing edge devices (also called elevons) instead of a horizontal tail are used for horizontal stability and control. Furthermore, reflexed airfoils are applied to minimize the pitching moment coefficient which is beneficial in terms of longitudinal stability [21]. The payload is distributed along the wing span, resulting in a lower bending moment and thereby a lower structural wing weight [21]. This effect is illustrated in Figure 2.4. The operative empty weight (OEW) of this configuration is even further reduced with respect to a conventional TAW configuration by the elimination of heavy joints between the fuselage and the wing and the absence of the horizontal tail. Moreover, the aerodynamic efficiency in terms of L/D is expected to be up to 20% higher compared to a conventional aircraft configuration [21]. A drawback of the BWB configuration is the the accessibility. The operational time to load and unload the containers is expected to be higher with respect to the multifuselage configuration because this configuration has no large cargo door located at the nose or the tail of the aircraft.

(a) BWB conﬁguration [22] (b) Span loader conﬁguration [14]

Figure 2.3: Illustration of different aircraft conﬁgurations of the all-lifting conﬁguration.

Figure 2.4: Comparison of lift and weight distribution of a BWB design and a conventional aircraft design [21].

The flying wing configuration is a tailless aircraft which only consists of one single wing which accommodates all fuel and payload [18]. Boeing and Lockheed analyzed this configuration during the period 1975-1985 with the aim to develop a very large dedicated cargo aircraft for civil and military applications [23–25]. This concept, also called the span loader configuration, has a low aspect ratio and a taper ratio of one. As a result, this configuration has a very thick wing section such that the cargo can be placed in the wings, where the cargo is loaded from the wing tips into the wing. A sketch of this aircraft configuration is given in Figure 2.3(b). Because the cargo is distributed along the entire wing span, the reduction in bending moment is even larger compared to the multifuselage configuration resulting in a potential lower structural weight [23]. The main drawback of this configuration that the lift distribution is not elliptical due to the large and thick tip section. As a result, the aerodynamic efficiency decreases due to a significant increase in induced drag. 8 2. Aircraft Concepts Definition

2.1.3. Discussion

For the concept of a hydrogen powered ultra large cargo aircraft, two potential aircraft configurations were identified: the multifuselage configuration and the all-lifting configuration. Compared to a conventional TAW configuration, the main benefit of the multifuselage configuration is the reduced required structural weight of the wing due to the bending moment relief of multiple fuselages along the wing span, whereas the all-lifting configuration has a potential higher aerodynamic efficiency compared to the conventional TAW configuration because it only consists of lifting surfaces.

With respect to a hydrogen powered ultra large cargo aircraft, the multifuselage configuration has an operational benefit by loading and unloading the containers simultaneously into two fuselages. However, the fuel tanks also have to be placed into the multiple fuselages because the large cylindrical/spherical pressurized fuel tanks do not fit in the conventional wing of the multifuselage configuration. Therefore, a drawback of this configuration is that the volume in the wings is unused.

Regarding the all-lifting configuration, three concepts were defined: the Burnelli configuration, the BWB configuration and the span loader configuration. With respect to a hydrogen powered ultra large cargo aircraft, the BWB configuration and the span loader configuration distributes the containers along the wing span, which is beneficial because of two reasons. First of all, a reduction of structural wing weight is achieved due to the bending moment relief of the containers. Next, the volume in the wing is used for the allocation of the containers. However, the storage of the large spherical/cylindrical pressurized fuel tanks have also to be kept in mind for the concept of a hydrogen powered ultra large cargo aircraft. Therefore, it would be beneficial to place the fuel tanks at the center line of the aircraft where a large amount of volume is available. Similar as the Burnelli configuration, the fuel tanks are then stored into a large airfoil shaped lifting body, whereas the containers are distributed over the rest of the wing like the span loader configuration. As a result, a BWB configuration, which combines the design characteristics of the Burnelli configuration and the span loader configuration, is a potential suitable concept for the application of the hydrogen powered ultra large cargo aircraft because of the efficient storage of the fuel tank and the containers along the wing span and the potential increased aerodynamic efficiency with respect to a conventional TAW configuration. A major drawback with respect to concept are the expected higher manufacturing and development costs compared to a conventional aircraft configurations because of the unconventional shape. Furthermore, the ease of accessibility to load and unload the containers is a design challenge because this concept does not have a large cargo door located at the nose or tail of the aircraft.

As can be concluded, a multifuselage configuration and a BWB configuration, which combines the design characteristics of the Burnelli configuration and the span loader configuration, are considered as potential aircraft configurations for the application of a hydrogen powered ultra large cargo aircraft. The multifuselage configuration has an operational and manufacturing benefit compared to the BWB concept, whereas the BWB configuration has a potential higher aerodynamic efficiency compared to the multifuselage configuration. In order to make a quantitative trade-off between the multifuselage concept and the BWB concept, a conceptual design study is performed for both aircraft configurations. The concepts considered for this study are explained in more detail in Section 2.3, but first the selection of the propulsion system for this aircraft concept is discussed in the next section.

2.2. Identiﬁcation of propulsion system

One of the main requirements is that the aircraft has to be powered by hydrogen. Nowadays, commercial aircraft are powered by kerosene instead of hydrogen. However, due to the run out of fossil fuels, several studies are performed to investigate alternative aviation fuels such as hydrogen. Already in 1976, NASA Langley performed a study on liquid hydrogen fueled subsonic passenger aircraft [26]. In 1998 the Sovjet-Union proved the technical feasibility by performing the first hydrogen powered aircraft flight with the Tupolev Tu-155 [27]. Regarding hydrogen fuel, there are two propulsion systems considered for this research: gas turbine engines and fuel cells. This section describes the two propulsion systems first, after which a selection is made for the concept of a hydrogen powered ultra large cargo aircraft. 2.2. Identification of propulsion system 9

2.2.1. Gas turbine engines For subsonic speed there are two types of gas turbine engines nowadays: the turbofan and the turboprop engine. Currently, most commercial passenger aircraft are powered by large turbofan engines. In the last half century, large improvements in the performance of the turbofan are made [14]. The maximum takeoff thrust of the largest turbofan engines has increased to 40 tons and a reduction in SFC of 30% has been achieved. Furthermore, the propulsive efficiency of turbofan engines has improved from 0.5 to 0.75 due to the increase in bypass ratio (BPR). While turbofans are nowadays mainly used for commercial passenger airplanes, this engine is not the most efficient engine in terms of propulsive efficiency [14]. A turboprop system with contra-rotating propellers namely has a propulsive efficiency between 0.8 and 0.9 [14]. The contra-rotating propellers of this propulsion system reduce the swirl in the slipstream of the rotor and thereby increases the propulsive efficiency by 7% compared to single-stage high speed propellers [14]. Two examples of this propulsion system are illustrated in Figure 2.5 2 3, where the NK-12 engine is the strongest turboprop engine nowadays with a power output of more than 11,000 kW [28]. The main drawback of this propulsion system is the significant reduction in propulsive efficiency for Mach numbers above 0.65, whereas turbofans can be used for Mach numbers above 0.8 [14]. Other concerns of the turboprop system are the impact of blade failure, cabin noise, and maintenance costs.

(a) NK-12 (b) D-27

Figure 2.5: Examples of contra-rotating propeller system used by current commercial aircraft.

Regarding the use of hydrogen for this propulsion system, Corchero and Montañés investigated the performance and the modifications required for conventional gas turbine engines to use hydrogen instead of kerosene [29]. It was found that no large hardware modifications on conventional gas turbine engines are required. Only the combustor and the fuel control system have to be redesigned, while the turbo-machinery remains the same [14, 29]. In terms of performance, a reduction of 64.7% in specific fuel consumption (SFC) is achieved by using hydrogen instead of kerosene [29]. According to Brewer [10], even more improvements in SFC can be achieved if hydrogen is used to cool components of the engine or if heat from the exhaust of the engine is used to vaporize the liquid hydrogen before it enters the engine. Jackson [30] investigated the improvement in SFC by pre-cooling the compressor or turbine air by the cold hydrogen or by pre-heating the hydrogen by the heat from the exhaust. He found an increase of SFC up to 4.4% for cruise condition compared to a conventional gas turbine engine powered by hydrogen.

2.2.2. Fuel cell Instead of using gas turbines to produce the desired thrust, fuel cells can also be used to generate power in order to drive propulsors. The promising beneﬁts of this technology are high reliability, low maintenance, low noise, and zero emissions (with the exception of water vapor) [31, 32]. Studies performed by NASA indicate that state of the art fuel cells can be used for electric propulsion of very light aircraft [31]. Currently, the

2Tumblr, Military and Aviation, http://enrique262.tumblr.com/post/136888197195/enrique262-kuznetsov-nk-12-soviet-turboprop (accessed 22-12-2016) 3Flickr, An-70 UR-NTK Progress D-27 propfans, https://www.ﬂickr.com/photos/lexich_76/3911940542 (accessed 22-12-2016) 10 2. Aircraft Concepts Deﬁnition largest hydrogen fuel cell powered aircraft is an unmanned aircraft powered by a 500 W fuel cell, built by Georgia Institute of Technology [33]. Sehra and Whitlow [32] state that a ten-fold increase in fuel cell power density is needed to enable electrically powered, large commercial passenger aircraft.

Fuel cells can also be used to generate power for the Auxiliary Power Unit (APU). Current turbine-powered APUs in commercial aircraft have an efficiency around 14% and emit 20% of the total airport ground based emissions [32]. These ground based emissions can be reduced by making use of fuel cells. Airbus investigated the benefits of APUs powered by hydrogen fuel cells with respect to the current turbine-powered APUs in commercial aircraft [34]. The estimated efficiency of a hydrogen fuel cell used for the APU is around 60%, which results in fuel weight savings for the APU up to 75% during ground operations. Drawbacks of the hydrogen fuel cells are the increased weight and the larger start-up time compared to turbine-powered APUs [34].

2.2.3. Discussion Despite of the potential environmental beneﬁts of the fuel cell, the low power density of current fuel cells makes this propulsion system impractical for the application of an ultra large cargo aircraft with a maximum payload weight of 1,200 tonnes. Therefore, a gas turbine engine is selected as propulsion system for this aircraft concept. The fuel cell technology could be applied as power generator for the APU to reduce the fuel consumption and emissions during ground operations. However, an analysis of this technology for this aircraft concept is outside the research scope.

Regarding the gas turbine engines, the main benefit of the turbofan engines with respect to the turboprop system is the possibility to fly efficient at high subsonic cruise speed, whereas the turboprop system has a higher propulsive efficiency for a lower cruise speed compared to the turbofan engine. No explicit requirements were defined for the cruise speed of the aircraft. The cruise speed of the aircraft has to be defined such that the aircraft meet the competitiveness target of a freight rate less than 250% of the freight rate of maritime transport. Therefore, it was decided to select turboprop engines with contra-rotating propellers as propulsion system for this aircraft because of the higher propulsive efficiency compared to a turbofan system. Because this propulsion system only operates at a high propulsive efficiency for low cruise speeds, the cruise Mach number is initially set to 0.5 for this aircraft concept. Despite of this low cruise Mach number compared to the cruise speed of current commercial aircraft, the aircraft still has a significant benefit in terms of transportation time with respect to maritime transport. The influence of the cruise speed on the aircraft design and performance is evaluated by the sensitivity analysis presented in Chapter 5. Similar to the largest cargo aircraft nowadays, the An-225, six engines are selected which are positioned at the leading edge of the wing.

2.3. Deﬁnition of ultra large cargo aircraft concepts

As can be concluded from the former sections, the multifuselage configuration and the BWB configuration are identified as potential suitable aircraft configurations for the application of a hydrogen powered ultra large cargo aircraft, where the turboprop engines with contra-rotating propellers are selected as most suitable propulsion system for this concept. This section discusses the main design choices of both aircraft concepts for the application of a hydrogen powered ultra large cargo aircraft, where an illustration of both aircraft concepts considered for this research is shown in Figure 2.6.

2.3.1. Multifuselage concept The multifuselage concept considered in this study has two fuselages positioned at 20% of the semi-span of the main wing measured from the center line. This position is chosen based on the conceptual design study on a twin-fuselage conﬁguration where De Jong and Slingerland found that the optimal location of the fuselage for minimal structural wing weight is at 20% of the semi-span measured from the center line [35]. The containers are distributed evenly over the two fuselages. The containers are placed in three rows next to each other, where the fuel tank is positioned above the containers. It is assumed that one cylindrical tank with a length equal to a row of containers is placed in each fuselage, where the radius of the fuel tank is determined based on the required fuel volume. Instead of multiple fuel tanks only one fuel tank in each fuselage is chosen 2.3. Deﬁnition of ultra large cargo aircraft concepts 11

(a) Multifuselage (b) Blended-Wing-Body

Figure 2.6: Illustration of the multifuselage and BWB concept considered for this study. to minimize the structural weight of the fuel tanks. An illustration of this lay-out is given in Figure 2.7. The position of the fuel tank is beneﬁcial for the minimization of the center of gravity (CoG) travel during the mission and the ease of access to load/unload the containers. Regarding the lay-out of the containers, three rows of containers are selected because this results in a similar slenderness ratio of the cabin with respect to large cargo aircraft [36].

The containers are loaded from the tail of the fuselage into the fuselage, where a rail system is used to move the containers to the front of the fuselage. By loading/unloading three containers in each fuselage simultaneously, the loading/unloading process is quite efficient. Regarding the nose of the fuselage, sufficient space is available for the cockpit. This section is pressurized to ensure good operational conditions for three crew members during the entire flight. On the contrary, the cabin is unpressurized where the containers and the fuel tank are located. Z-direction

Y-direction

(a) Cross-section front view (b) Side view

Figure 2.7: Overview of the placement of the containers and the fuel tank into the fuselage for the multifuselage concept.

The wing of this aircraft concept consists of two sections, divided by the position of the fuselage: the inner wing section and the outer wing section. The inner part is a straight untapered wing, which is beneﬁcial from a structural point of view [35], whereas the outer wing is shaped based on a deﬁned leading sweep and taper ratio. This is illustrated in Figure 2.6(a).

Regarding the empennage, a vertical tail is placed on the tail cone of each fuselage, where a horizontal tail is positioned between the two fuselages. The pilot is able to adjust the incidence angle of the horizontal stabilizer to trim the aircraft, whereas rudders on the vertical tails are used for lateral stability and control. Six turboprop engines with contra-rotating propellers are placed at the leading edge of the wing outboard of the fuselage to increase the bending moment relief of the engines on the wing. Because the engines are attached to the leading edge of the wing, a high wing configuration is chosen to provide sufficient engine clearance, whereas the landing gear system is stored into the fuselage to minimize the landing gear height. This is also 12 2. Aircraft Concepts Definition visualized in Figure 2.6(a).

2.3.2. Blended-wing-body concept The BWB concept considered for this study consists of one large tapered wing and two vertical tails. The wing is divided into three sections, where the inner wing section is defined as the first two sections and the outer wing by the third section. Trailing edge devices, also called elevons, are located at the inner wing section for longitudinal stability and control. The trailing edge sweep of the inner wing section is assumed to be zero to maximize the effectiveness of the elevons, whilst a reflexed airfoil is chosen to minimize the pitching moment. The different sections of the wing planform is shown in Figure 2.8(a).

(a) Top view (b) 3D sketch

Figure 2.8: Characteristics of wing planform of the BWB concept considered for this research.

As discussed in Section 2.1.3, the BWB concept considered for this study combines the design characteristics of the different all-lifting configurations. The first wing section is straight and untapered where the fuel tank is located in a large airfoil shaped body between the front and the rear spar. The fuel tank is placed next to the center line of the aircraft, because of the large required volume to store the hydrogen tanks. The second and third sections are swept and tapered wing sections, where the containers are distributed along the wing span towards the wing tip in order to decrease the bending moment on the wing. A hand drawing of the placement of the fuel tank and the containers in the wing of the BWB concept is illustrated in Figure 2.8. The containers are loaded from a panel at the bottom side of the wing located at the trailing edge of the wing, where a kneeling landing gear system is applied to minimize the vertical distance between the ground and the aircraft during loading and unloading. The containers are distributed along the wing span by a rail system. Because of simplicity for loading and unloading, only one level of containers are considered, which means that the containers are not stacked on top of each other. The containers are placed between the front and the rear spar. In order to prevent a too far aft position of the CoG which could lead to stability problems, the containers are placed to the front as far as possible. Similar as for the multifuselage concept, sufficient volume is available in the nose of the first section for the pressurized cockpit. Similar to the multifuselage concept, the engines are placed at the leading edge of the outer wing. Further- more, two vertical tails are placed between the first and the second section of the wing. The main benefit of two vertical tails is the reduction of the vertical tail height and the increase in redundancy of lateral stability and control with respect to one large vertical tail. Conceptual Design Methodology

In the former chapter two potential aircraft concepts are introduced for the application of a hydrogen powered ultra large cargo aircraft. A discussion on the benefits and the drawbacks of the aircraft configurations was given in Section 2.1.3. In order to make a quantitative trade-off between the different aircraft concepts, a conceptual design is developed and analyzed for both aircraft concepts. Subsequently, the aircraft design and performance are compared on several criteria, such as the aircraft dimensions, the aircraft weight (OEW), the aerodynamic performance and the transport efficiency. This chapter discusses the methodology of the conceptual design framework to design and analyze both aircraft concepts. First, the design approach is explained, followed by an overview of the architecture of the conceptual design framework. Next, a detailed description of the different design and analysis tools implemented in the conceptual design framework is given.

3.1. Design approach

The aircraft design process consists of three phases: the conceptual design phase, the preliminary design phase and the detailed design phase. The goal of the conceptual design phase is to analyze and compare different aircraft concepts to each other for given TLRs. Therefore, mainly low-fidelity analysis tools and empirical methods are used to quickly design and analyze the different aircraft concepts. Based on the first estimations of the design and performance of the different aircraft concepts, a quantitative trade-off between the different aircraft concepts is made. The best aircraft concept is then selected as baseline design for the preliminary design phase. Because of the simplified methods and tools used during the conceptual design phase, the level of design accuracy is very low at the end of the conceptual design phase. The aircraft design and performance is therefore analyzed in more depth during the preliminary design phase. A large amount of designers and engineers are used to analyze the baseline design by making use of higher fidelity methods. Because this requires a large amount of resources and time, only one aircraft concept is selected from the conceptual design phase and no large modifications are made to the aircraft concept. After the preliminary phase, the aircraft is designed to make it ready for production during the detailed design phase, which is the most time-consuming and expensive part of the aircraft design process. In the traditional aircraft design approach, the aircraft design and performance is quickly analyzed and optimized during the conceptual design phase through parametric variation of a few critical design parameters, for example the wing loading and thrust loading, such that the aircraft design and performance meet the TLRs defined by the customer [37]. The focus is on the analysis of the aerodynamic and propulsion discipline because these disciplines have the largest influence on the performance of the aircraft. A major drawback of this design approach is that the other disciplines, such as the structures and control, are analyzed during the preliminary and detailed design phase where the design freedom is significantly reduced. This is illustrated in Figure 3.1(a). The unconventional aircraft concepts proposed in the former chapter have a complex and integrated design,

13 14 3. Conceptual Design Methodology

(a) Traditional design approach (b) MDO design approach

Figure 3.1: A comparison between the traditional design approach and the multidisciplinary optimization design approach, which is indicated by the dotted line [38]. where the mutual interference of design parameters between the different disciplines is not always as intu- itive as for the conventional TAW conﬁguration. Therefore, a multidisciplinary optimization (MDO) approach has been used both to integrate more disciplines, as well as to take into account the mutual interactions between the different disciplines already in the conceptual design phase. By making use of this MDO approach different unconventional aircraft concepts can be designed and analyzed during the conceptual design phase. As a result, the design freedom and the design knowledge increases during the ﬁrst phase of the aircraft design process. This effect is shown in Figure 3.1(b).

Over the last years, researchers of the Flight Performance and Propulsion chair of Delft University of Technol- ogy have been working on a tool for the multidisciplinary aircraft design process, called the Initiator [39, 40]. The Initiator is a conceptual design tool implemented in the programming environment MATLAB [41]. This tool is able to generate both conventional configurations as well as unconventional configurations, such as a Prandtl Plane configuration, a canard configuration and a three-surface configuration [39]. However, in its current state of development this conceptual design tool is not able to suitable to support the design case considered for this research because of several limitations. The first limitation is that the tool is only able to generate passenger aircraft configurations and not dedicated cargo aircraft configurations. Next to that, the Initiator assumes that the aircraft is powered by kerosene, whereas alternative fuels are not considered by this tool. Another limitation of the Initiator is that it is currently not able to generate a converged design for a BWB configuration and a multifuselage configuration. Because of these limitations, it is decided to not use the Initiator as conceptual design tool for this research.

Therefore, a conceptual design framework is developed which is able to design and analyze both aircraft concepts. Usually handbook methods are used to quickly generate and analyze a conceptual design based on TLRs. These methods are, however, based on empirical data of existing aircraft and are therefore only valid for conventional aircraft configurations. The proposed aircraft concepts discussed in Chapter 2 have significant differences in design features with respect to the conventional TAW configuration, which make the handbook methods unreliable. As a result, more physics based methods instead of empirical methods are preferred and used where possible for the development and analysis of a conceptual design for these unconventional aircraft concepts. However, the level of design accuracy of this conceptual design framework is lower compared to the Initiator.

Because only a limited amount of data is available of these unconventional aircraft concepts, it is difficult to have a good starting point for the MDO framework. Therefore, a Design of Experiments (DOE) is performed to investigate the design space of both aircraft concepts and to define a good starting point for the MDO of the aircraft design. The design space is explored by generating and analyzing possible candidate designs, where the design space is defined by the amount of design parameters that are varied during the DOE, also called design variables. However, in order to have a reasonable uniform coverage of the design space, the number of candidate designs that have to be generated and analyzed increases exponentially with the number of design variables. This effect is also called the curse of dimensionality [42]. As mentioned earlier, quasi-analytical methods which rely more on physics instead of empirical data are used to design and analyze the unconven- 3.2. Development of a MDO framework for conceptual design 15 tional aircraft concepts. A drawback of these methods is the increased computational time compared to the empirical and statistical methods. Therefore, only a limited amount of candidate designs can be generated and analyzed in order to reduce the computational time of the DOE. As a result, only the main design parameters of the aircraft concepts are selected as design variables for the DOE, such as the wing loading and the aspect ratio, where other more detailed design parameters, such as the airfoil distribution along the wing span, are kept constant. Based on the information of the possible candidate designs evaluated during the DOE, a good starting point is selected for the MDO of the aircraft design and performance. The objective of this optimization is to design the aircraft concept for the minimum fuel weight required to fulfill the mission, while the aircraft concept meets predefined design requirements. The fuel weight is chosen as objective function to reduce both the environmental impact and the operational costs as much as possible [15]. Instead of using the computational expensive quasi-analytical methods, surrogate models are used to reduce the computational time of the optimization. Surrogate models are analytical models which mimic the input- output behavior of a complex system based on a set of experiments [42]. Therefore, surrogate models are built based on the design and performance of the possible candidate designs analyzed during the DOE. The set-up of the MDO framework for the synthesis and analysis of the conceptual design is discussed in more detail in the next section.

3.2. Development of a MDO framework for conceptual design The architecture of the MDO framework is visualized and explained by means of an Extended Design Struc- ture Matrix (XDSM) [43]. The XDSM is able to show both the data dependency and the process ﬂow into one diagram. Different shapes are used to distinguish different types of processes and input/output data. Generic processes are visualized by rectangles, whereas parallelograms are used for data input and output. The data dependency of each process is illustrated by the use of a thick gray line connecting the nodes. The order of execution of the different processes is indicated by a numbering scheme. Furthermore, a thin black line is used to emphasize the process connections. The XDSM developed to explain the MDO framework for the conceptual design study is illustrated in Figure 3.2, where a description of the symbols is given in Table 3.1. The framework can be divided into two parts: 1. Design of Experiments (DOE) and construction of surrogate models (step 1-8). 2. Numerical optimization based on surrogate models (step 9-16).

Table 3.1: Description of the symbols of the XDSM for the MDO framework.

Symbol Description c Constraint values c˜ Approximation of constraint values by surrogate models FW Fuel weight f Objective value f (0) Objective value for the initial design vector of the optimization f˜ Approximation of objective value by surrogate models f § Optimal objective value MTOW Maximum take-off weight x(0) Initial design vector for optimization x(s) Sample design vector of DOE x§ Optimal design vector 16 3. Conceptual Design Methodology Report f x ⇤ ⇤ Convergence ,16 0, Check Input 6: 16 : 16 6: 16 ! f f c ˜ 7: ,6 1, DOE ! 2: 4: MTOW,FW ,4 2, MDA ! 3: iue3.2: Figure 3: MTOW Analysis 3: 3: x i ( s t ) ,FW rhtcueo D rmwr o ocpuldesign. conceptual for framework MDO of Architecture t Constraint Analysis 5: 5: x i ( s ) 7: 7 feasible 7: 7: design a 7 Best FW a : 7: x c FW : i ( i ( s c s i ) ( ) s ) 8: Optimizer ,11 9, 1: 11 1:˜ : 11 x (0) i ! ,f f c ˜ 0 10: 0 (0) 8 7: Surrogate Objective 8 a 8: Model a , 8,10: FW 0: 10 : x FW i ( s i ( ) x s 0 ) 8 Surrogate a straints Models 8: 8: , 8 Con- 8,10: 0: 10 a x c : i ( i ( s c s ) ) x 0 3: 13 MTOW,FW 14 MDA ! 12, 13: 2: 12 MTOW Analysis 3: 13 13: x ⇤ t ,FW t Constraint Analysis 5: 15 15: x ⇤ 3.2. Development of a MDO framework for conceptual design 17

3.2.1. Design of experiments

The ﬁrst step is to explore the design space by generating and analyzing a number of possible candidate designs. Each design is deﬁned by a unique design vector, xi . As explained in the former section, this design vector includes the main design parameters to size and shape the aircraft concept. A multidisciplinary analysis (MDA) is performed to generate a conceptual design for each design vector. The MDA consists of a coupled system of multiple design and analysis modules, where the coordinator ’MDA’ in the XDSM ensures that the fuel weight and the MTOW are converged at the end of the analysis. Next, the feasibility of the conceptual design is evaluated by the constraint analysis module where the constraint values, ci , indicate if the design is feasible or infeasible.

After all possible candidate designs are evaluated, surrogate models are built based on the information of all candidate designs. These surrogate models are used for the numerical optimization of the aircraft design and performance. Because the objective of this optimization is to minimize the fuel weight of the aircraft design, a surrogate model is built of the fuel weight. Next, surrogate models are built of the constraint functions to take into account the design requirements deﬁned by the constraint analysis module during the numerical optimization.

As discussed in Section 3.1, physics based methods instead of empirical methods are used where possible to design and analyze both aircraft concepts. Because of the increased computational time of these methods with respect to empirical methods, only a limited amount of possible candidate designs, and thereby a limited amount of design variables, are selected for the DOE to reduce the total computational time of this process. The computational time required to evaluate one candidate design is on average in the order of 3 to 5 minutes using a computer with a 2.80 GHz Intel Core2Duo T9600 processor and 4GB installed RAM mem- ory. In order to analyze and compare the different aircraft concepts within a limited timeframe, six design variables and 200 possible candidate designs are chosen for the DOE. The main design parameters of the wing and the empennage are selected as design variables, because these components have the largest inﬂu- ence on the overall aircraft design and performance. The methods and tools used by the design and analysis modules of the MDA and the constraint analysis module are explained in more detail in Section 3.3. This section also deﬁnes which design variables are used to size and shape the wing and the empennage during the DOE.

3.2.2. Numerical optimization

Based on the surrogate models constructed after the DOE, a numerical optimization is performed to optimize the aircraft design for the minimum fuel weight while it satisﬁes the predeﬁned constraint requirements. The feasible conceptual design corresponding to the lowest fuel weight found by the DOE is selected as starting point of the numerical optimization. More information about the set-up of the numerical optimization is given in Section 3.4.

Because surrogate models are an approximation of the input-output behavior of the complex MDA system, a check has to be performed on the accuracy of the fuel weight and constraints values given by the surrogate models. Therefore, a conceptual design is made by performing the MDA and the constraint analysis module based on the optimal design vector found by the numerical optimization, x§. The objective value found by the optimizer, f˜, and the fuel weight of the MDA, f , are compared by the convergence check module. If the difference is smaller than 1% and the design meet the requirements deﬁned by the constraint analysis module, a feasible conceptual design for the minimum fuel weight is found. Otherwise, the output of the MDA and the constraint analysis module are added to the data collected by the DOE. Based on this new information, the surrogate models are updated and the numerical optimization is performed again. This process iterates until the difference in fuel weight between the optimizer and the MDA is smaller than 1% and the design meets the predeﬁned constraint requirements.

After convergence is established, a ﬁnal report is generated. This report includes detailed information about the design and performance of the aircraft concept. The report is automatically developed in a LaTeX for- mat. 18 3. Conceptual Design Methodology

3.2.3. Implementation This framework is implemented in the programming environment MATLAB [41]. Compatible with this programming environment, the SUrrogate MOdelling (SUMO) toolbox is used for the generation of the different design vectors for the DOE and the construction of the surrogate models [44, 45]. The surrogate models are built by making use of Kriging, where the different design vectors, xi , are generated by means of Latin Hyper- cube Sampling (LHS) method. This method distributes random samples evenly over the design space such that the design space is explored as efﬁcient as possible [42]. The implementation of the numerical optimization is discussed in Section 3.4.3.

Regarding the MDA, aircraft analysis tools and custom written codes are used to generate and analyze the design and performance for both aircraft concepts. A user manual to execute this MDO framework is given in Appendix A. The methods used by the different disciplines of the MDA are discussed in detail in the next section.

3.3. Multidisciplinary analysis

As mentioned in Section 3.2, a MDA is performed for each design vector (x0) generated for the DOE. The output of the MDA system is a converged conceptual design for a given design vector and predeﬁned TLRs. Similar to the MDO framework the architecture of the MDA system is illustrated in a XDSM, as shown in Figure 3.3. As indicated by the thin black line, the analysis modules are executed in sequence. This architecture is also called a Gauss-Seidel MDA process [43].

The ﬁrst step is to make an estimation of the MTOW, fuel weight, and the OEW based on a Class I weight estimation. Next, a wing power loading diagram is generated to determine the design point of the aircraft. The design point is the combination of wing loading, WS, and power loading, WP, for which the aircraft is designed. After that, more reﬁned analysis modules are used to determine the fuel weight and MTOW more accurately based on the actual geometry of the aircraft. The ’MDA’ coordinator determines if convergence of the fuel weight and MTOW is established. Convergence of the fuel weight and MTOW is established if the difference in weight at the start of the iteration, MTOWt and FWt , and at the end of the iteration, MTOWt and FWt , is smaller than 5%. After a converged design is found, the feasibility of the aircraft is checked by the constraint analysis module. The methods and tools used by the different analysis modules are discussed next.

AR W/S x0

0: Class I 1:MTOW0,FW0 TLRs Weight Analysis

1, 9 2: 2:Ct ,Ct , ⌘ t 3:MTOWt,FWt 4:MTOWt,FWt 5:MTOWt,FWt 6:MTOWt,FWt MDA! D0 Lmax p

2: Wing W/S, W/P 3:W/P 10 : (W/S) Power max Loading

3: 4: 5: 6: 7: 10:Aircraft Aircraft Geometric Aircraft Aircraft Aircraft Aircraft 8:rprop Geometry Geometry Modeller Geometry Geometry Geometry Geometry Vfuel,Vcargo

4: OEW OEW 5: OEW 8: OEW Analysis

5: 6: CoG CoG CoG location Analysis location

6: 7: 10: L/D 8: L/D 9:CD0 Aero Stall xac,Cn CL,cruise,CD,cruise Dcruise Analysis Location Cn ,Cl

7: CLmax 9:CLmax CLmax Analysis

9: 8: MTOW, MTOW, Perfor- FW FW mance ⌘p ⌘p Analysis

10: Con- c straint Analysis

Figure 3.3: Architecture of MDA system. 3.3. Multidisciplinary analysis 19

3.3.1. Class I weight estimation The MDA starts with a first estimation of the fuel weight, MTOW and OEW based on a Class I weight estimation. The fuel fraction method is used to estimate the fuel weight [36]. This method divides the mission into different mission segments, where a fuel fraction is defined for each mission segment. The mission profile including the definition of the different mission segments is shown in Figure 3.4.

Figure 3.4: Mission proﬁle of the aircraft.

The fuel fractions of the cruise and loiter segment are determined by the Brequet equations (Equations 3.1- 3.2), whereas the fuel fractions of the other mission segments are taken from statistical data and can be found in Table 3.2 [46]. These statistical values are based on existing kerosene fuelled aircraft, which means that these values are conservative. In case of a hydrogen powered aircraft, the fuel fractions are expected to be higher due to the high heat of combustion.

¥p L Wcruise,begin R ln (3.1) = g C · D · W 0 · p µ ∂cruise µ cruise,end ∂

¥p L Wloiter,begin E ln (3.2) = V g C · D · W · 0 · p µ ∂loiter µ loiter,end ∂

Table 3.2: Fuel fraction of the different mission segments [46].

Mission segment Fuel fraction Start-up engine 0.990 Taxi 0.990 Take-off 0.995 Climb 0.980 Descent 0.990 Landing, taxi, shut-down 0.992

The lift-over-drag ratio (L/D) during cruise is estimated using Equation 3.3 [14], whereas the L/D during loiter is taken as 0.866 L/D [36]. The speciﬁc fuel consumption (SFC), C , is estimated based on the most powerful · p turboprop engine with contra-rotating propellers, NK-12 [28], and reduced by 64.7% to account for the use of hydrogen instead of kerosene [29]. An overview of the assumptions made for the estimation of the L/D and the Brequet equations is given in Table 3.3. Finally, the fuel weight can be expressed as a fraction of the MTOW based on the overall fuel fraction, as shown in Equation 3.4. The overall fuel fraction, M ff, is the product of all fuel fractions.

L º AR e · · (3.3) Swet D = v4 C f u Sref u · · t

W M MTOW (3.4) fuel = ff · 20 3. Conceptual Design Methodology

To make a first estimation of the OEW a regression line is made between the OEW and the MTOW based on a database of existing large cargo aircraft [28] and the Pelican Ultra [47]. The Pelican Ultra is a conceptual aircraft design analyzed by Boeing, which is able to transport standard 40 foot ISO containers up to a maximum payload weight of 2,800,000 lb [47]. The statistical data of the existing large cargo aircraft can be found in Appendix B. As shown in Figure 3.5, the coefficient of determination (R2) is almost equal to one meaning that proportion of variance is very small. Therefore, a good first estimation of the OEW can be made by making use of this regression line. Based on the coefficients a and b from the regression line, the OEW can be expressed as a fraction of the MTOW, as shown in Equation 3.5. Finally, the MTOW is calculated using Equation 3.6, after which the fuel weight and OEW can be determined by Equation 3.4 and Equation 3.5, respectively.

1200

Pelican Ultra

1000

800 y = 0.3529x + 25.972 R² = 0.99796 [kg] 3 -

x 10 600 MTOW

400

200

0 0 500 1000 1500 2000 2500 3000 OEW x 10-3 [kg]

Figure 3.5: Regression line based on statistical data of existing large cargo aircraft [28] and the Pelican Ultra [47].

OEW a MTOW b (3.5) = · +

MTOW W (1 M ) MTOW a MTOW b (3.6) = pay + ° ff · + · + OEW Fuel Weight | {z } | {z }

Table 3.3: Input parameters for Class I weight estimation provided by the handbook of Raymer [36].

Parameter Symbol Unit BWB Multifuselage Skin friction coefficient C f - 0.003 0.003 Oswald efficiency factor e - 0.8 0.8 Ratio wetted area and wing area Swet -3 6 Sref Propulsive efficiency ¥p - 0.8 0.8 2 Graviational constant g0 m/s 9.81 9.81 8 8 Specific fuel consumption C kg/J 2.16 10° 2.16 10° p · · 3.3. Multidisciplinary analysis 21

3.3.2. Wing power loading diagram Next, a wing power loading diagram is created to determine the power loading, W/P, based on a given wing loading, W/S. This diagram identiﬁes the region where combinations of W/P and W/S meet preliminary performance requirements. Based on the handbook of Roskam [46], the performance requirements of the take- off distance, landing distance, cruise speed and climb performance requirements imposed by the aviation authorities for the certiﬁcation of large transport aircraft (CS 25) are included in this diagram. The combination of W/P and W/S for which the aircraft is designed for is called the design point. Regarding the take-off and landing distance, it is assumed that the aircraft has to be able to take-off and land within 3,300 meters. This distance is equal to the runway distance of existing large airports.

Veriﬁcation of this module is performed by constructing a wing power loading diagram for the Lockheed Super Hercules C130J, a dedicated cargo aircraft powered by turboprop engines. Based on the reference data of Jane’s all the world’s aircraft [28], the same power loading was found for a wing loading of 4500 N/m2. The result is shown in Figure 3.6, where details of the veriﬁcation can be found in Appendix C.

0.5 Landing distance s=778 m

0.45 Take-o, sTO=930 m Cruise speed V= 177 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure 3.6: Wing power loading diagram for the Lockheed Super Hercules C130J.

3.3.3. Geometric modeller After the design point is taken, a geometric model of the aircraft is made. This module makes a parametric model of the proposed aircraft concepts described in Section 2.3. The geometric model consists of multiple components: a wing, a fuselage, an empennage and multiple engines. Because of the signiﬁcant differences in design features between the two proposed concepts, the geometric models of both aircraft concepts are discussed separately.

Multifuselage concept The multifuselage concept considered in this study consists of two fuselages, a main wing, two vertical tails, a horizontal tail and six engines. The input vector for the modelling of the different components is given in Table 3.4. There are two types of input parameters: fixed parameters and variables. The values of the variables are varied during the DOE to explore the design space, whereas the values of the fixed parameters are kept constant. The upper and lower limit of each variable are also indicated in Table 3.4. As can be seen in Table 3.4, the fuselage and the engines are sized based on fixed parameters, whereas the size and shape of the wing planform is defined by design variables. Next, it can be seen that the size of the horizontal and vertical tail is varied during the DOE by means of a design variable, where the shape of the empennage is fixed based on a predefined aspect ratio and taper ratio. The sweep of the empennage is derived from the aspect ratio and 22 3. Conceptual Design Methodology the taper ratio, while it is assumed that the sweep at the trailing edge is zero to maximize the effectiveness of control surfaces. During the research it was found that the size of the empennage has a larger influence on the stability and control of the aircraft compared to the shape of the empennage. Therefore, only the size of the empennage is varied during the DOE, whereas the shape of the empennage is fixed. The upper and lower bound of the design variables were first defined based on empirical data and refined during the research to ensure that feasible possible candidate designs are evaluated during the DOE. For example, a very low wing loading in combination with a large aspect ratio would result in structural problems with respect to the wing because of an unrealistic large wing span. On the other hand, the wing has to be long enough to provide sufficient space for the position of the engines.

Table 3.4: Input vector of multifuselage conﬁguration for the geometric model.

Parameter Symbol Type Value Unit Wing Wing loading W/S Variable 4500-5500 N/m2 Aspect ratio AR Variable 5-8 - Taper ratio ∏ Variable 0.3-0.5 - Sweep angle leading edge §LE Variable 0-15 °

Empennage Ratio vertical tail area over wing area SV /S Variable 0.03-0.06 - Ratio horizontal tail area over wing area SH /S Variable 0.1-0.3 - Aspect ratio horizontal tail ARh Fixed 4 - Aspect ratio vertical tail ARv Fixed 1.3 - Taper ratio horizontal tail ∏h Fixed 0.5 - Taper ratio vertical tail ∏v Fixed 0.3 -

Fuselage Number of containers ncont Fixed 100 - Slenderness ratio nose l/dnose Fixed 1.5 - Slenderness ratio tail l/dtail Fixed 2.0 -

Engines Disk loading DL Fixed 448 kW/m2

The geometric model of the two fuselages consists of the nose, the cabin and the tail cone, as shown in Figure 3.7(a). The cabin of the fuselage is sized based on the payload configuration and the fuel tank sizing, where the containers are distributed evenly over the two fuselages. As explained in Section 2.3.1, the containers are placed in three rows next to each other, where the cylindrical fuel tank is positioned above the containers. As a result, the length of the fuel tank is equal to the length of the container row, whereas the radius of the fuel tank is sized based on the required volume to store the fuel weight. Based on the amount of containers and the size of the fuel tank, the length and the outer diameter of the cabin (d fus) are determined. The nose and the tail cone of the fuselage are designed based on predefined slenderness ratios of 1.5 and 2.0, respectively. These values are provided by Kroo to ensure that flow seperation and shock waves are avoided [48]. The geometric model of the wing consists of two parts divided by the position of the fuselage. As mentioned in Section 2.3.1, the position of the fuselages with respect to the wing span are fixed at 20% of the semi-span based on the study of De Jong and Slingerland on a twin-fuselage configuration [35]. The inner part is a straight untapered wing, which is beneficial from a structural point of view [35], whereas the outer wing is shaped based on a defined leading sweep and taper ratio. The main parameters of the wing are determined based on Equations 3.7-3.9, where the parameters of these equations are illustrated in Figure 3.7(b)

MTOW g0 Sref · (3.7) = WS

b AR S (3.8) = · ref q 3.3. Multidisciplinary analysis 23

Sref cr (3.9) = 2b b (1 ∏) iwing + owing · +

(a) Fuselage

(b) Wing

Figure 3.7: Parametrization of the fuselage and the wing planform for the multifuselage concept.

In case of wing podded engines, an initial wing position with respect to the fuselage is that 25% of the mean aerodynamic chord (MAC) coincides at 50% of the fuselage length [49]. However, it was found that an unrealistic large horizontal tail is required to provide longitudinal stability in case of this wing position. Therefore, it was iteratively found that the wing has to be positioned with respect to the fuselage such that leading edge of the MAC coincides with 50% of the fuselage length. Regarding the airfoil characteristics, a NACA23015 airfoil is chosen at the root chord, which is a well-known airfoil for low-subsonic aircraft [49]. Similar to a conventional wing of existing aircraft, the thickness-to- chord ratio, t/c, decreases along the wing span [50]. The t/c at the tip is determined based on the empirical relationship given by Equation 3.10, where t/cavg is taken as 10% [50]. It is assumed that the t/c decreases linearly from the root towards the wing tip.

3 t/ctip t/croot t/cavg · + (3.10) = 4

The empennage of the multifuselage concept consists of two vertical tails and one horizontal tail. The horizontal tail ensures longitudinal stability and controllability, where the aircraft is trimmed by adjusting the incidence angle of the horizontal tail. The two vertical tails have a rudder located at 70% of the chord to provide lateral stability and controllability. The surface area of the horizontal and vertical tail are determined based on the ratio of the vertical tail area / horizontal tail area over the wing area defined by the input vector. Next, the horizontal and the vertical tail are shaped based on the fixed aspect ratio and taper ratio defined in 24 3. Conceptual Design Methodology

Table 3.4, which are provided by the handbook of Raymer [36], whilst it is assumed that the sweep at the trailing edge is zero to maximize the effectiveness of the control surfaces. Provided by the handbook of Raymer [36], a NACA0010 proﬁle is chosen as airfoil for both the horizontal and vertical tail. The vertical tails are positioned at the end of the fuselage, whereas the horizontal tail is positioned between the fuselages.

The engines are placed outboard with respect to the position of the fuselage at the leading edge of the wing. A detailed engine sizing would be outside the scope of this research. Therefore, analytical and statistical relationships are used for the preliminary sizing of the engine. The diameter and the length of the engine are sized by the design rules for a rubber turboprop engine provided in the handbook of Raymer [36], whereas the propeller radius is determined based on the required power per engine and the disk loading, as shown in Equation 3.11. The disk loading is taken as 448 kW/m2 based on the statistical data of the strongest turboprop engine, the NK-12 [28].

Preq,engine rprop (3.11) = s º DL ·

BWB concept

As described in Section 2.3.2, this aircraft concept consists of one wing, two vertical tails and six engines. The input vector for the modelling of these components is given in Table 3.5. As can be seen in Table 3.5, the modelling of the empennage only requires parameters of the vertical tail, because longitudinal stability and control is established by elevons instead of a horizontal tail. Similar to the multifuselage concept, the size of the vertical tail is varied during the DOE, whereas the shape of the vertical is based on a ﬁxed aspect ratio and taper ratio. The leading edge sweep of the vertical tail is derived from the aspect ratio and the taper ratio by making use of the assumption that the sweep at the trailing edge is zero to maximize the effectiveness of the rudder. Furthermore, it can be seen in Table 3.5 that the wing planform is mainly sized and shaped based on design variables. During the research the upper and lower bound of these design variables are reﬁned such that realistic wing and empennage planforms are analyzed during the DOE.

Table 3.5: Input vector for the geometric model of BWB conﬁguration.

Parameter Symbol Type Value Unit Wing Wing loading W/S Variable 2500-3100 N/m2 Aspect ratio AR Variable 3-6 - Taper ratio of section 2 ∏2 Variable 0.4-0.6 - Taper ratio of section 3 ∏3 Fixed 0.3 - Leading edge sweep of section 3 §3 Variable 0-20 ° Ratio span section 2 over span section 3 b fraction Variable 0.25-0.4 -

Empennage Ratio vertical tail area over wing area SV /S Variable 0.03-0.06 - Aspect ratio vertical tail ARv Fixed 1.3 - Taper ratio vertical tail ∏v Fixed 0.3 -

Engines Disk loading DL Fixed 448 kW/m2

In order to explain the parameters used to size and shape the wing planform, an illustration of the parame- terization is given in Figure 3.8. As mentioned in Section 2.3.2, the ﬁrst wing section is straight and untapered where the fuel tank is located. The second and third sections are swept and tapered wing sections where the containers are located. The sweep at the trailing edge of the ﬁrst two wing sections is assumed to be zero to maximize the effectiveness of the elevons. The sizing of the wing planform is an iterative process, because the fuel tank size depends on the chord length of section 1, c1, whereas c1 depends on the size of the fuel tank, b1. The iterative sizing process is illustrated in Figure 3.9. 3.3. Multidisciplinary analysis 25

Figure 3.8: Parametrization of the BWB concept.

Figure 3.9: Flow diagram of the iterative sizing process of the planform for the BWB concept.

First, the required planform area and the wing span are determined by Equations 3.7 and 3.8, whereas a first guess of c1 is made by Equation 3.12. Based on the length between the front and rear spar at c1, the diameter of the tank is determined to provide sufficient volume for the hydrogen storage. The wing span of section 1, b1, equals the diameter of the fuel tank. Next, the wing span of section 2 and section 3 are determined based on their ratio, b fraction, given by the input vector. The chord lengths of section 2, c2, and section 3, c3, are calculated by means of the taper ratio. A fixed taper ratio for the outer wing is chosen to limit the amount of design variables during the DOE. A taper ratio of 0.3 is chosen for the outer wing to ensure that the root chord is large enough to store the fuel tank, while an unrealistic pointy wing is avoided which could result in bad stall behavior. Because section 2 has a straight trailing edge, the sweep at the leading edge is derived from the wing span and the taper ratio of section 2. The leading edge sweep of section 3 is given by the input vector. Finally, the root chord of section 1, c1 is adjusted until the planform matches the required planform area.

Sref c1 (3.12) = 2b (1 ∏ ∏ ) · + 1 2 After the wing planform is defined, the container distribution along the wing span is determined by evaluating the available volume at the cross-section of the wing. This process starts at the cross-section next to to the fuel tank. Starting from the front spar position, the containers are placed behind each other until the back spar is reached or the thickness is not sufficient with respect to the height of the container. As mentioned in Section 2.3.2, only one floor of containers is considered for simplicity with respect to loading and unloading the containers. Then, the cross-section next to the former row of containers is evaluated for the placement of containers following the same procedure. This process continues until all containers are placed into the wing or too less space is available for the placement of the containers. The placement of the containers start at the front spar position to have the CoG position to the front as far as possible, which is beneficial in terms of stability. The process of the container distribution is visualized in Figure 3.10, where the placement of one row containers in side view direction is illustrated in Figure 3.11(b). As a consequence of this methodology, 26 3. Conceptual Design Methodology the rear spar should be located as much aft as possible to ensure that all containers are placed into the wing. However, a more forward position of the rear spar is beneficial for structural reasons. Based on this trade-off, the front and rear spar are located at 10% and 50% of the chord length, respectively.

(a) First row (b) Second row (c) Final distribution

Figure 3.10: Illustration of the methodology to determine the container distribution along the wing span for the BWB concept.

Regarding the airfoil characteristics, the reflexed airfoil HS522 is chosen because of the low moment coefficient which is beneficial for trimming capabilities [51]. A relatively high t/c ratio of 16% is taken at the root chord to provide sufficient volume for the placement of the tank, whereas the t/c ratio at c2 and c3 are set at 12% to provide sufficient volume for the placement of the containers. It is assumed that the t/c decreases linearly along the wing span. Based on the chord length at the center line of the aircraft and the t/c ratio of 16%, a check is performed to evaluate if the thickness at section 1 is large enough to store the fuel tank. The placement of the fuel tank including the shape of the reflexed airfoil is shown in Figure 3.11(a).

20 30 15 20 10 10 5 0 t [m] 0 t [m] -10 -5 -10 -20 -15 -30 0 20 40 60 80 0 20 40 c [m] c [m] (a) Fuel tank (b) Containers

Figure 3.11: Cross-section of the wing to illustrate the placement of the containers and the fuel tank.

Regarding the sizing of the vertical tail and the engines, the same design rules are used as deﬁned for the multifuselage concept. The vertical tails are positioned at the trailing edge between the ﬁrst and the second section, whereas the engines are placed at the leading edge of the outer wing section.

3.3.4. Analysis of operative empty weight A first rough estimation of the OEW is made by the Class I weight estimation module. This module determines the OEW more accurately by calculating the weight of the main components individually based on their actual dimensions. As shown in Equation 3.13, the OEW consists of the structural weight, the power plant weight (Wpp), the fixed equipment weight (Wfe) and the fuel tank weight (Wft). The fixed equipment weight includes the weight of all the aircraft systems. The structural weight is the sum of the wing weight (Wwing), fuselage weight (Wfus), horizontal tail weight (WHT), the vertical tail weight (WVT) and the landing gear weight (Wlg). 3.3. Multidisciplinary analysis 27

OEW W W W W W W W W (3.13) = wing + fus+ HT + VT + lg + pp + fe+ ft structural weight | {z } Except of the wing weight and the fuel tank weight, a Class II weight estimation method is used to determine the weights of the main components. This method consists of statistical equations derived from a large database of existing aircraft. Because the components of the aircraft, except of the wing and the fuel tank, do not differ signiﬁcantly from a conventional aircraft, the statistical equations of the Class II weight estimation given in the handbook of Raymer (Eq 15.26 - Eq. 15.45) are used [36]. The power plant weight of the turboprop engine is estimated based on the required power per engine [36]:

Preq,engine 0.803 Wpp 1.3 (0.96 ) (3.14) = · · 1000

Fuel tank weight The design of a hydrogen fuel tank differs from a conventional kerosene fuel tank, because a hydrogen fuel tank has to be pressurized and well-insulated to ensure cryogenic conditions for the storage of hydrogen. Therefore, statistical and empirical methods derived from existing aircraft are unreliable for the fuel tank weight estimation. The structural weight of the fuel tank and the required insulation weight are estimated based on the work of Böhm [52]. For a cylindrical tank surrounded by polyurethane foam, Böhm determined the hydrogen tank mass for different tank diameters at different altitudes. Furthermore, he determined the insulation mass as a function of the insulation thickness and the tank diameter. The results are shown in Figure 3.12. The relative structural mass to volume is extrapolated to the given cruise altitude, resulting in 7.4 kg/m3. The relative insulation mass to area is taken for an insulation thickness of 16 centimeters, which is provided by Brewer and Böhm to minimize evaporation of hydrogen [10, 52].

(a) Structural tank weight (b) Insulation weight

Figure 3.12: Relationship between the tank diameter and the structural tank weight and insulation weight [52].

Wing weight The wing weight mainly depends on the wing geometry, the wing structure and the loads acting on the wing. Regarding a conventional wing, engines are usually attached to the wing where the fuel is stored into the wing. Therefore, the main loads acting on a conventional wing of existing aircraft are the aerodynamic load, the fuel weight load, and the power plant load. This is different for the unconventional aircraft concepts considered in this study, where the payload weight is distributed along the wing span. As discussed in Section 2.1, this has a signiﬁcant effect on the bending moment of the wing and thereby on the structural weight of the wing. As a 28 3. Conceptual Design Methodology result, the statistical equations of the Class II weight estimation are unreliable for the wing weight estimation of both proposed aircraft concepts because the distribution of the payload weight along the wing span is not taken into account for the wing weight estimation.

In order to take into account all the applied loads on the wing weight estimation, the quasi-analytical method, called EMWET [53], is used. This method structurally sizes the wing box to ensure that the wing is able to resist the applied loads onto the wing. Therefore, the amount and the distribution of the material of the upper and lower panels of the wing box are determined analytically to withstand the bending moment, where the spar webs are structurally sized based on the shear loads. Subsequently, the wing box weight is computed based on the sizing of these components, where the weight of secondary components of the wing (ribs, ﬁxed leading and trailing edge, high-lift devices and control surfaces) are estimated based on semi-empirical relationships presented by Torenbeek [54]. These semi-empirical methods also include the weight penalties for joints, attachments and cut-outs.

In order to structurally size the spar webs and the equivalent upper and lower panels of the wing box, EMWET requires the critical loads acting on the wing, the material of the wing box and a geometric model of the wing. It is assumed that the wing is built of aluminum where the properties are listed in Table 3.6. Regarding the critical load conditions for the wing box weight estimation, a maximum positive load factor of 2.5 is taken at an operating Mach number of 0.5. Based on the parametric model of the wing built by the geometric modeller module, the commercial vortex lattice method (VLM) AVL [55] is used to determine the load distribution along the wing span for this critical condition. Next to the aerodynamic loads, the following bending moment relief loads including their position with respect to the semi-span are implemented into EMWET:

• Weight of the wing box

• Weight of the containers, including the container ﬂoor

• Weight of the fuel tank, including the fuel weight

• Weight of the fuselage in case of the multifuselage conﬁguration

• Weight of power plants

The quasi-analytical method EMWET is applied to both aircraft concepts for the wing weight estimation with the awareness that this method assumes a cantilever beam to determine the bending moment on the wing. This assumption is not completely appropriate for the multifuselage concept where the wing is attached to the fuselages, which are positioned outboard of the center line of the aircraft. As a result, this has an impact on the bending moment along the wing span and thereby on the wing box weight estimation. Recommendations for further research to determine the wing box weight more accurately are given in Chapter 8.

Table 3.6: Material properties of aluminum for the wing box weight estimation by EMWET.

Parameter Symbol Value Unit Material density Ω 2800 kg/m3 Maximum tensile stress upper panel æ 295 106 N/m2 tupper · Maximum tensile stress lower panel æ 295 106 N/m2 tlower · Maximum compressive stress upper panel æ 295 106 N/m2 clower · Maximum compressive stress lower panel æ 295 106 N/m2 clower · Young’s modulus E 7 1010 N/m2 ·

3.3.5. Center of gravity

Based on the weight and the position of the main components, the CoG of the aircraft is calculated. The position of the main components is given in Table 3.7 [56]. Due to the fuel consumption, the position of the CoG shifts during the mission. To evaluate this CoG shift, four conditions are deﬁned: OEW, MTOW, zero-fuel weight (ZFW) and cruise condition. The design weight during cruise condition is determined by Equation 3.15. The difference between the MTOW and the design weight deﬁnes the amount of fuel spent before cruise condition. 3.3. Multidisciplinary analysis 29

W MTOW (MTOW W ) (3.15) des = · ° fuel q

Table 3.7: Center of gravity location of the main components

Component Center of gravity location Fuel tank Halfway of the fuel tank, dependent on the conﬁguration Payload (containers) Halfway of each row, dependent of the conﬁguration Engines Halfway of the engine length Main wing 40% of MAC of main wing Vertical tail 40% of MAC of vertical wing Horizontal tail 40% of MAC of horizontal tail wing Landing gear Weighted average of the nose and main landing gear position Fuel systems Halfway of the fuel tank APU 90% of length aircraft Instruments 2 meters from the nose of the aircraft Hydraulics Rear spar position of MAC Electronics Center of wing box Avionics 2 meters from the nose of the aircraft Anti-ice system Leading edge position of MAC Container handling system Similar as center of gravity location of payload

Based on the different CoG locations during the mission due to the fuel consumption, the most forward and most backward CoG position are determined. These positions have an influence on the position of the landing gear system. The main landing gear system has to be positioned behind the most backward CoG position to prevent tip-over. Furthermore, the nose landing gear has to carry between 5% and 20% of the weight in order to ensure good steering capabilities [36]. The distance between the main landing gear and the nose landing gear, B, is calculated by Equation 3.16, whereas the position of the main landing gear and the nose landing gear with respect to the CoG position is determined by Equation 3.17. As an example, the parameters B, M f , and Ma are illustrated for the BWB concept in Figure 3.13. The position of the landing gear system has an influence on the backward CoG position, whereas the backward CoG position has an influence on the position of the landing gear system. This is an iterative process until the difference in landing gear position is smaller than 1%. In lateral direction, it is assumed that the main landing gear is located in the fuselages for the multifuselage concept, whereas the position of the main landing gear for the BWB concept is based on Equation 3.18, which is provided by the handbook of Raymer to prevent the aircraft from overturning the main wheels [36].

xcg,backward xcg,forward B ° (3.16) = 0.2 0.05 °

Ma 0.05 B = · (3.17) M 0.20 B f = ·

y x x tan(25°) (3.18) main,lg = lg,main ° lg,nose · ° ¢ The amount of wheels required to withstand the loads during ground operations is computed based on the loads acting on the nose and the main landing gear and the maximum allowable loading on current aircraft wheels [36]. It is assumed that six wheels are connected to one strut, which is similar to current large commercial aircraft such as the Airbus A380 and the Boeing 747. The landing gear height is determined based on clearance requirements provided by the handbook of Raymer [36]. The different requirements are illustrated in Figure 3.14. First of all, the landing gear has to be long enough to prevent that the tail touches the ground during take-off and landing. It is assumed that the maximum angle of attack during take-off and landing is 15 degrees. As a consequence, a maximum angle of attack of 15 degrees during take-off and landing is taken into account for the estimation of the maximum lift 30 3. Conceptual Design Methodology

Figure 3.13: Placement of the landing gear with respect to the center of gravity positions based on the design rules of the handbook of Raymer [36]. coefﬁcient. Next, a minimum clearance of 15 centimeters is required for the engines and the wing tip with respect to the ground for a bank angle of 5 degrees. Finally, a minimum distance of 50 centimeters is required between the aircraft and the ground.

(a) Side view

(b) Front view

Figure 3.14: Clearance requirements for the calculation of the landing gear height.

3.3.6. Aerodynamic analysis

A ﬁrst estimation of the aerodynamic performance of the aircraft, in terms of L/D, is made for the Class I weight estimation. This module analyzes the aerodynamic performance of the aircraft in more detail by taking into account the actual geometry of the aircraft. The aerodynamic analysis is analyzed in terms of L/D for a trimmed aircraft in cruise condition. For low subsonic cruise speed, the total drag of the aircraft mainly consists of the proﬁle drag of the different aircraft components, the induced drag and the trim drag. This drag break-down is illustrated in Figure 3.15.

The induced drag, trim drag and profile drag of the wing are determined by the quasi three-dimensional (3D) aerodynamic solver, called Q3D [57], whereas the component buildup method given in the handbook of Raymer [36] is used to calculate the profile drag of the other aircraft components. Based on the actual geometry of the wing and the flight parameters during cruise condition, the aerodynamic solver Q3D is able to determine the drag coefficients more accurately with respect to empirical methods, while the computational time is relatively low compared to higher fidelity methods [57]. Elham validated this aerodynamic solver with the results of higher fidelity computational dynamic fluid solvers for a range of different flight conditions for a conventional wing [58].

Regarding the aerodynamic analysis of the BWB concept, Q3D is not able to determine the proﬁle drag over the entire wing accurately because of the 3D effects over the centerbody of the aircraft. These 3D effects are 3.3. Multidisciplinary analysis 31

Figure 3.15: Drag break-down for a low subsonic aircraft. illustrated in Figure 3.16. The streamlines over the inner wing section are curved, which has a significant effect on the pressure distribution over the inner wing section [59]. The streamlines over the outer wing section are straight, which makes Q3D suitable to predict the profile drag of the outer wing. Similar to the research of the Silent Aircraft Initiative [60], the profile drag of the inner wing section is estimated based on the component buildup method given by Raymer [36]. More information about the aerodynamic solver Q3D and the component buildup method is given below.

Figure 3.16: Streamlines over a BWB conﬁguration determined by a full panel method [59].

Q3D Based on the geometry of the lifting surfaces including the control surfaces, the flight parameters and the CoG position during cruise condition, the aerodynamic solver Q3D is able to analyze the aerodynamic performance of the aircraft in trimmed cruise condition by making use of the commercial VLM tool AVL [55] and the two-dimensional (2D) airfoil analysis tool XFOIL [61]. The flight parameters include the cruise speed, 32 3. Conceptual Design Methodology cruise altitude and the required lift coefficient at cruise condition. This lift coefficient is calculated based on Equation 3.19, where the design weight, Wdes is determined by Equation 3.15.

Wdes CL (3.19) des = 0.5ΩV 2S

The induced drag of the aircraft is analyzed by AVL based on a Trefftz plane analysis [55]. The trim drag is included in this term, where AVL is able to compute the required flap deflection to trim the aircraft. Fur- thermore, AVL is able to determine the lift distribution over the wing, but also the stability derivatives of the aircraft. The stability derivatives are used by the constraint analysis module, described in Section 3.3.9, to analyze the stability and controllability of the aircraft. The profile drag of the wing is estimated based on the strip theory, where several 2D sections on the wing are analyzed by making use of XFOIL [57, 61]. More information about the procedure of this approach is given in [57].

Component buildup method This method determines the profile drag of each component as a function of the flat plate skin-friction drag coefficient, C f , the component form factor (FF), and the interference effect of the component on different components, Q. This relationship is defined in Equation 3.20 [36]. The form factor is an estimation of the pressure drag due to viscous separation. The flat plate skin-friction drag coefficient, C f , depends on the Reynolds number and on the ratio of laminar flow with respect to turbulent flow over the component, k, as shown in Equation 3.21. It is assumed that 20% of the flow over the aircraft components is laminar. In order to take into account the surface roughness of the components, a cut-off Reynolds number is defined. This Reynolds number depends on the characteristic length of the component and the skin-roughness parameter of smooth paint, provided by Raymer [36]. The lowest Reynolds number between the cut-off and the actual Reynolds number is taken for the calculation of C f . The form factor of the different components are estimated based on Equations 12.30 - 12.33 given in the handbook of Raymer [36], whereas the interference factors are stated in Table 3.8. Regarding the drag estimation of the inner wing section of the BWB concept, it is assumed that the flow is fully turbulent over the inner wing section because of the high Reynolds number. The interference factor is assumed to be zero, because of the smooth blended transition between the inner wing section and the outer wing section.

C fc FFc Qc Swetc CD0,c · · · (3.20) = Sref

1.328 0.455 C f k (1 k) 2.58 0.65 (3.21) = · pRe + ° · log Re 1 0.144M 2 10 + ° ¢ ° ¢ Table 3.8: Interference factor (Q) of the different aircraft components provided by Raymer [36].

Aircraft component Q Fuselage 1.0 Horizontal tail 1.05 Vertical tail 1.05 Nacelle 1.3

3.3.7. Maximum lift coefficient At the start of the design process, a first estimation is made for the maximum lift coefficient during landing,

CLmax,landing , and take-off, CLmax,TO. These parameters are required for the construction of the wing power loading, where CLmax,landing and CLmax,TO have an influence on the landing and take-off performance, respectively. This module determines the maximum lift coefficient by taking into account the wing geometry, the airfoil characteristics and the type of high-lift devices. The maximum lift coefficient is namely the 3.3. Multidisciplinary analysis 33

sum of the maximum lift coefﬁcient in clean condition, CLmax,clean , and the increase due to high-lift devices,

¢CLmax .

Clean wing

The maximum lift coefﬁcient for a clean wing is estimated based on the 2D maximum airfoil lift coefﬁcient,

Clmax , and a 3D conversion factor, CLmax /Clmax , as shown in Equation 3.22. The 2D airfoil analysis tool XFOIL [61] is used to determine Clmax. The airfoil geometry is taken where the maximum local lift coefﬁcient occurs in cruise condition. This position is determined based on the lift distribution derived from the aerodynamic analysis of Q3D. The ratio CLmax/Clmax is estimated based on the leading edge radius of the airfoil and the sweep angle by making use of the graphs in the handbook of Raymer [36].

CLmax CLmax,clean Clmax ( ) (3.22) = · Clmax

High lift devices

The maximum lift coefficient during landing and take-off is mainly increased by high-lift devices located at the trailing edge of the wing. These trailing edge devices increases the camber of the airfoil and thereby the maximum lift coefficient during landing and take-off. On the other hand, these devices also increases the pitching down moment by creating additional lift at the aft section of the airfoil. This nose-down moment is usually counteracted by the horizontal tail. However, the BWB configuration does not have a horizontal tail and therefore no high-lift devices are selected for this configuration. Furthermore, high-lift devices are not required for the BWB configuration because of its typical low wing loading compared to conventional TAW configurations [21]. Regarding the multifuselage concept, a single slotted Fowler flap is chosen as trailing edge device because of its simple working principle but significant effect on CLmax [36]. The increase of the maximum lift coefficient due to a high-lift device is determined by Equation 3.23, where the increase in maximum local lift coefficient, ¢Clmax , is estimated based on Equation 3.24 [49]. The ratio of the chord extension due to the high-lift device, c0/c, depends on the flap deflection, as shown in Figure 3.17. It is assumed that the flap deflection during take-off and landing is 15 and 40 degrees, respectively [49]. The area of the trailing edge device, S flapped/S, is taken as 0.4 to provide enough space for ailerons and spoilers.

S flapped ¢CL 0.9¢Cl cos(§hingeline) (3.23) max = max S

c0 ¢Cl 1.3 (3.24) max = c

3.3.8. Performance analysis

Finally, the performance of the propeller, in terms of the propulsive efficiency, is analyzed. The propulsive efficiency is determined based on the actuator disk theory and is defined by Equation 3.25 [62]. This method assumes that the flow passing through the propeller forms a well-defined streamtube, where the propeller is replaced by an actuator disk. Furthermore, it is assumed that the flow is incompressible and irrotational, while the pressures and velocities are evenly distributed over the disk area. Therefore, the propulsive efficiency is overestimated because this method does not include losses due to rotational kinetic energy in the slipstream. Moreover, the propeller blades are subjected to induced drag which also has a negative impact on the propulsive efficiency. In order to include these losses, the propulsive efficiency is multiplied by 0.9.

2 ¥p (3.25) = T 1 1 2 + + 0.5ΩAdiskV0 r 34 3. Conceptual Design Methodology

Figure 3.17: Increase in chord extension as a function of the ﬂap deﬂection and type of high-lift device [49].

Based on the more refined calculation of the lift-over-drag ratio and the propulsive efficiency, the fuel fraction of the cruise and loiter segment are recalculated. Subsequently, the fuel weight and the MTOW are determined based on the new OEW and fuel fractions by Equations 3.4 and 3.6. The design loop iterates until the difference between the fuel weight and MTOW defined at the beginning and end of design loop is smaller than 5%. After a converged design is found, the feasibility of the aircraft is checked by the constraint analysis module.

3.3.9. Constraint analysis Finally, the feasibility of the aircraft design is checked based on several design and performance requirements, or also called constraints. Constraints are defined to ensure preliminary performance requirements, volume requirements and stability and controllability requirements. The first requirement is that the wing loading of the aircraft does not exceed the maximum wing loading defined by the landing distance requirement in the wing power loading diagram in order to ensure that the aircraft is able to land within the predefined landing distance.

WS WS (3.26) ∑ max Next, in case of the BWB concept, enough volume has to be present to store both the fuel as well as the containers. This requirement only applies to the BWB concept, because the fuel tanks and containers are placed into the wing after the wing planform is already deﬁned, whereas the fuselage shape of the multifuselage concept is sized based on the container distribution and the fuel tank dimensions. Therefore, the available volume next to the center line of the BWB concept, Vft,available, has to be larger than than the volume required to place the fuel tank, Vft,required. Similarly, the available volume in the wing where containers can be placed, Vcargo,available, has to be larger than the volume required for the placement of the containers, Vcargo,required.

Vft,required Vft,available ∑ (3.27) V V cargo,required ∑ cargo,available The feasibility of the aircraft also depends on the stability of the aircraft. The stability of the aircraft can be divided into longitudinal stability and lateral directional stability. In longitudinal direction the aircraft has to generate a counteracting moment if a small change in angle of attack occurs, which is called static longitudinal stability. This condition is met if the aerodynamic center of the aircraft is behind the CoG position. The 3.4. Set-up of numerical optimization 35

aerodynamic center is the point where the pitching moment coefﬁcient, Cm, is constant with respect to the angle of attack [36]. The position of the aerodynamic center is found by Q3D of the aerodynamic analysis module with the awareness that the effect of the fuselage and the propeller engines on the aerodynamic center position is not taken into account. The longitudinal stability is measured by the static margin (SM). The static margin is the distance between the aerodynamic center and the CoG as a percentage of the MAC, as shown in Equation 3.28. To ensure that longitudinal static stability is established, a minimum static margin, SM, of 5% is deﬁned.

xac xcg SM ° 5% (3.28) = MAC ∏

Next to the longitudinal stability, the stability in lateral direction is analyzed. Lateral directional stability is defined by two stability derivatives: the static directional stability derivative, CnØ , and the effective dihedral derivative, ClØ . To ensure lateral stability the aircraft requires a positive CnØ and a negative ClØ , which means that the aircraft generates both a restoring yawing moment as well as a restoring rolling moment at a small side slip angle. Both stability derivatives are taken from the aerodynamic analysis module. The effect of the fuselages on the static lateral stability is not taken into account by Q3D. Therefore, the destabilizing moment of the fuselage is estimated based on Equation 3.30, where the coefficients KN and KRL are determined from the graphs of the handbook of Roskam [63]. The influence of the wing-fuselage interaction and the propeller slipstream over the wing on the effective dihedral derivative is assumed to be negligble compared to the ClØ produced by the wing.

CnØ 0 > (3.29) C 0 lØ <

Sside lbody Cn 57.3 KN KRL · (3.30) Øfuselage = ° · · · S b ·

Finally, the controllability of the aircraft in case of one engine inoperative (OEI) is analyzed. According to Roskam [64], the maximum rudder deflection has to be smaller than 25 degrees to counteract the yawing moment due to the loss in thrust and the windmilling drag of the inoperative engine. This moment is calculated based on Equation 3.31, where it is assumed that the windmilling drag is equal to 25% of the thrust of one engine in take-off condition. The rudder deflection to counteract this moment is determined by Equa- tion 3.32. The most critical condition is taken where the most outer engine becomes inoperative. For this analysis the minimum control speed is taken, which is equal to 1.2 V . The effect of the rudder deflection · stall on the yawing moment, Cnrudder , is taken from the aerodynamic analysis module.

N y (T 0.25 T ) (3.31) OEI = eng TO + · TO

NOEI ±rudder 25° (3.32) = 0.5 Ω V S b C ∑ · · ms · · · nrudder

3.4. Set-up of numerical optimization

After all candidate designs of the DOE are designed and analyzed by making use of the MDA system described in the former section, surrogate models are built based on the information of the DOE for the numerical optimization. This section explains the set-up of the numerical optimization in more detail. First, the optimization problem is formulated, after which the optimization algorithm and the implementation of the numerical optimization are discussed in detail. 36 3. Conceptual Design Methodology

3.4.1. Formulation of optimization problem The formulation of the optimization problem is defined by the design vector, the objective function and the constraint functions. The design vector is optimized to either minimize or maximize the objective function, while the solution meet the requirements defined by the constraint functions. The objective of this numerical optimization is to minimize the required fuel weight to fulfill the mission requirements in order to reduce the operational costs and the environmental impact as much as possible. The constraint functions are defined based on the design and performance requirements of the constraint analysis module. In mathematical terms, the objective function and the constraint functions are given in Equations 3.33 and 3.34. Instead of using the computational expensive MDA system to find the optimal design vector for the minimum fuel weight, the numerical optimization makes use of surrogate models to reduce the computational time. The surrogate models are only able to predict the objective and constraint functions within the design space of the DOE, because these models are built based on the information of the DOE. In order to prevent the use of the computational expensive MDA system, this means that the design space of the numerical optimization has to be similar to the design space of the DOE. Therefore, the same design variables including the lower and upper bound with respect to the DOE are defined for the numerical optimization. The design vector of both aircraft concepts can be found in Table 3.9 and Table 3.10. As discussed in Section 3.2, only six design variables were selected for the DOE to limit the total computational time of the DOE because of the effect of the curse of dimensionality. Therefore, only the main parameters for the sizing of the wing and the empennage are chosen as design variables, because those parameters have the largest influence on the design and performance of the aircraft.

Table 3.9: Design vector of the multifuselage conﬁguration including bounds.

Parameter Symbol Unit Lower bound Upper bound Aspect ratio AR - 5 8 Wing loading W/S N/m2 4500 5500 Taper ratio ∏ - 0.3 0.5 Leading edge sweep angle §LE °0 15 Ratio horizontal tail area over wing area SH /S - 0.1 0.3 Ratio vertical tail area over wing area SV /S - 0.03 0.06

Table 3.10: Design vector of the BWB conﬁguration including bounds.

Parameter Symbol Unit Lower bound Upper bound Aspect ratio AR - 3 6 Wing loading W/S N/m2 2500 3100 Taper ratio section 2 ∏2 - 0.4 0.6 Leading edge sweep angle section 3 §3 °0 20 Ratio span section 2 over span section 3 b fraction - 0.25 0.4 Ratio vertical tail area over wing area SV /S - 0.03 0.06

3.4.2. Optimization algorithm The numerical optimization is performed by making use of a gradient-based optimization algorithm. The sequential quadratic programming (SQP) algorithm is selected for this optimization procedure. This algorithm is able to minimize a single-objective function, while taking into account both linear and non-linear constraints. By evaluating the gradient of the objective function and the constraint functions for each design variable, the optimizer iteratively searches towards an optimum by taking step sizes in the direction of the gradient. An optimum is found if a change of the design vector, dx, has an insignificant effect on the objective value. The main benefit of this optimization technique is that it finds relatively fast an optimum. However, the main drawback is that the optimizer cannot guarantee a global optimum, because it can get stuck into a local optimum. As a result, the starting point of the optimization has a large influence on the final result. Therefore, the feasible conceptual design corresponding to the lowest fuel weight found during the DOE is selected as starting point for the numerical optimization. 3.4. Set-up of numerical optimization 37

3.4.3. Implementation As mentioned in Section 3.2.3, the conceptual design framework is implemented in the programming environment MATLAB. In order to make use of the gradient-based optimizer SQP,the MATLAB function fmincon is used for this research. The design vector is normalized in order to make sure that the gradient is determined for the same relatively change of each design variable. Normalization is done by dividing each entry of the design vector by the reference value of the initial design vector, which corresponds to the feasible conceptual design with the lowest fuel weight of all candidate designs analyzed during the DOE. In order to make use of the MATLAB function fmincon, the constraints have to be deﬁned as equality constraints (c 0) or inequality = constraints (c 0). Therefore, the following objective function and constraint functions are deﬁned for this ∑ optimization problem:

Wfuel f (3.33) = Wfuel,start

5 SM c(1) ° 0 = 5 ∑ c(2) C 0 = lØ ∑ c(3) C 0 = ° nØ ∑ ± 25 c(4) rudder ° 0 = 25 ∑ (3.34) WS WSmax c(5) ° 0 = WSmax ∑ Vft,available Vft,required c(6)§ ° 0 = Vft,required ∑ Vcargo,available Vcargo,required c(7)§ ° 0 = Vcargo,required ∑

In order to evaluate the objective function and the constraint functions deﬁned by Equation 3.33 and Equa- tion 3.34, surrogate models are built for the parameters listed in Table 3.11. As mentioned in Section 3.2, the SUMO toolbox is used to built the surrogate models based on the characteristics of the design and performance of the candidate designs analyzed during the DOE [44, 45].

Table 3.11: Parameters selected for the construction of the surrogate models.

Parameter Symbol Unit

Fuel weight FW kg

Static margin SM % MAC

Weather cock stability ClØ -

Static lateral stability CnØ -

Rudder deﬂection ±rudder °

WS WS Wing loading constraint ° max - WSmax V V Fuel volume constraint ft,available° ft,required - Vft,required

V V Cargo volume constraint cargo,available° cargo,required - Vcargo,required

In order to prevent that the optimization algorithm searches endless towards an optimum, termination criteria are deﬁned. An optimum design is found if the change of the design vector and the objective value is smaller than 0.1% and 1% compared to the previous design point, respectively. Furthermore, the violation of 38 3. Conceptual Design Methodology all the inequality constraint values have to be smaller than 0.01 to ensure that the ﬁnal design meet all the design and performance requirements. The results of the numerical optimization for both aircraft concepts are discussed in the next chapter, where an insight into the surrogate models and the design space of both aircraft concepts is given. Furthermore, the design and performance characteristics of the conceptual design found by the numerical optimization are discussed in the next chapter. Aircraft Design and Performance

By making use of the conceptual design framework described in the former chapter, both aircraft concepts are designed and analyzed based on the TLRs stated in Section 1.4. At the end of the conceptual design framework a report is generated including all the characteristics of the aircraft design and performance. This chapter discusses the main results of the design and performance of both concepts, where the ﬁnal reports of the multifuselage concept and the BWB concept can be found in Appendix D and Appendix E, respectively. First, the main characteristics of the design and the performance of both aircraft concepts are discussed separately. Subsequently, the aircraft concepts are compared to each other in terms aircraft dimensions, aircraft weight (OEW), aerodynamic performance, stability and controllability, and transport efﬁciency. Finally, the main differences between the aircraft concept considered for this study and the aircraft concept analyzed by the 2015 Design Synthesis Exercise are discussed.

4.1. Multifuselage concept

As discussed in the former chapter, the design space is explored by means of a DOE. Subsequently, a numerical optimization is performed within this design space by making use of surrogate models. Therefore, ﬁrst an insight into the design space of the multifuselage concept is given. Next, the design and performance of this aircraft concept are discussed in terms of aircraft dimensions, weight analysis, aerodynamic analysis and stability analysis.

Design space As discussed in Section 3.2, six design variables are selected for the DOE. Therefore, it is not possible to visual- ize the design space into one plot. However, in order to get insight into the design space the influence of each design variable on the surrogate models is investigated. Therefore, each design variable is varied between the lower and upper bound, while keeping the rest of the design variables constant. As baseline configuration, the design vector at the start of the numerical optimization is selected, which is the best feasible conceptual aircraft design with respect to the fuel weight of all candidate designs analyzed during the DOE. This design vector is shown in Table 4.1, where the influence of the design variables on the surrogate models of the objective function and the constraint functions are given in Figure 4.1. A violation of the constraint functions is illustrated in red, which indicates the infeasible region of the design space.

As can be seen in Figure 4.1(a), the fuel weight is mainly inﬂuenced by the aspect ratio, the horizontal tail area (SH /S), and the taper ratio. A lower taper ratio results in a lower wing weight because of the reduced tip loading. Furthermore, an optimum can be found for the aspect ratio with respect to the fuel weight. The aspect ratio is a trade-off between the L/D and the wing weight. An increase of the aspect ratio results in a higher L/D due to the reduction in induced drag, but also in a heavier wing because of the increased bending moment caused by the enlarged wing span.

Regarding the constraint functions, it can be seen in Figures 4.1(b)-4.1(f) that the design space is constrained

39 40 4. Aircraft Design and Performance

Table 4.1: Initial design vector of multifuselage concept for numerical optimization.

Parameter Symbol Value Lower bound Upper bound Unit Wing loading W/S 4718 4500 5500 N/m2 Aspect ratio AR 7.2 5 8 - Taper ratio ∏ 0.32 0.3 0.5 - Sweep angle leading edge §LE 0.05 0 15 ° Ratio vertical tail area over wing area SV /S 0.057 0.03 0.06 - Ratio horizontal tail area over wing area SH /S 0.1 0.1 0.3 -

Table 4.2: Final design vector of multifuselage concept for numerical optimization.

Parameter Symbol Value Lower bound Upper bound Unit Wing loading W/S 4566 4500 5500 N/m2 Aspect ratio AR 6.8 5 8 - Taper ratio ∏ 0.3 0.3 0.5 - Sweep angle leading edge §LE 0.0 0 15 ° Ratio vertical tail area over wing area SV /S 0.05 0.03 0.06 - Ratio horizontal tail area over wing area SH /S 0.1 0.1 0.3 -

by the CnØ constraint and the maximum wing loading constraint during the optimization. The latter constraint is violated for a too large sweep angle. If the sweep angle increases the maximum lift coefﬁcient decreases, resulting in a lower maximum wing loading. However, the fuel weight is minimized for a minimum sweep angle meaning that this constraint is not active during the numerical optimization.

Regarding longitudinal stability, the static margin is always larger than 5% and mainly inﬂuenced by the horizontal tail area and the wing loading. The static margin becomes larger for an increase in horizontal tail area because more lift aft of the CoG is produced. Next, a higher wing loading results in a smaller and lighter wing. As a result, the CoG of the aircraft shifts forward and the static margin increases.

Regarding lateral stability and controllability, the static lateral directional stability derivative CnØ and the rudder deflection in case of OEI are mainly influenced by the size of the vertical tail. Because the sizing of the vertical tail is based on the ratio of vertical tail area over the wing surface area, CnØ increases for a higher ratio of Sv /S and a smaller wing loading (larger S and thereby larger Sv ), whereas the rudder deflection decreases for a higher ratio of Sv /S and a smaller wing loading. The rudder deflection is also influenced by the aspect ratio because of the influence of the wing span on the rudder deflection, as shown in Equation 3.32. As can be seen in Figures 4.1(e)-4.1(d), the rudder deflection is always smaller than 25 degrees, while the CnØ becomes negative for a low SV /S or too large W/S. The effective dihedral derivative, ClØ , is always negative and mainly influenced by the sweep angle.

As can be concluded from the evaluation of these surrogate models, an optimum can be found for the aspect ratio with respect to the fuel weight, whereas the horizontal tail area, the wing sweep angle and the taper ratio of the wing are minimized during the numerical optimization to design the aircraft concept for the minimum fuel weight. The taper ratio is at the lower bound to minimize the wing weight and thereby the fuel weight. For conventional aircraft, the taper ratio however has to be large enough to provide enough volume to store the fuel into the wing. Because the fuel is stored into the fuselage for this concept, there is no volume constraint for the wing. Therefore, a minimum lower bound of 0.3 was defined for the taper ratio to prevent an unrealistic pointy wing which has bad stall performance. Because of the low subsonic cruise speed, the leading edge sweep angle has no beneficial impact on the aerodynamic performance and is therefore minimized to reduce the wing weight. Furthermore, the sweep angle and the horizontal tail area are minimized because the longitudinal stability constraint is inactive. Because controllability is not taken into account, the horizontal tail area is minimized as well within the design space defined by the lower and upper bound. Finally, it can be concluded that the requirement of static lateral stability CnØ is active during the optimization, which is mainly influenced by the wing loading and the vertical tail area. These conclusions match with the results of the final design vector found by the numerical optimization, as shown in Table 4.2. 4.1. Multifuselage concept 41

#105 5.8 -4 AR AR W/S -5 W/S 5.7 taper taper sweep -6 sweep Sh/S Sh/S Sv/S Sv/S -7 5.6 -8

5.5 -9

-10 Fuel Weight [kg]

5.4 Static Margin [%] -11

-12 5.3 -13

5.2 -14 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

(a) Fuel weight (b) Static margin

0.02 0.01 AR 0.01 W/S taper 0.005 0 sweep Sh/S Sv/S -0.01 0

-0.02 -0.005 [-] [-] -0.03 l,beta n,beta C

C -0.01 -0.04

-0.05 -0.015 AR W/S -0.06 taper -0.02 sweep -0.07 Sh/S Sv/S -0.08 -0.025 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

26 0.2 AR 0.15 W/S 24 taper sweep 0.1 Sh/S 22 Sv/S 0.05 AR 20 W/S 0 taper sweep 18 Sh/S -0.05 Sv/S -0.1

16 WSmax constraint [-] Rudder deflection [deg] -0.15 14 -0.2

12 -0.25 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

(e) Rudder deﬂection in case of OEI (f) Maximum wing loading constraint

Figure 4.1: Surrogate models of the objective function and constraint functions for multifuselage concept.

Aircraft design By making use of the MDA system, a conceptual design is generated and analyzed based on the ﬁnal design vector given in Table 4.2. A geometric overview of this conceptual design is illustrated in Figure 4.2. Regarding the aircraft dimensions, the aircraft has a wing span of 206 meters, whereas the fuselage length and equivalent diameter are 150 meters and 13.6 meters, respectively. With a vertical tail of 20.3 meters tall positioned on top of the fuselage the aircraft height is 33.9 meters. The propulsion system consists of six contra-rotating turboprop engines with a propeller radius of 7.5 meters. The design point in the wing power loading diagram is given in Figure 4.3. The feasible design space is indi- 42 4. Aircraft Design and Performance

(a) Top View (b) Front View

Figure 4.2: A three-dimensional overview of the multifuselage configuration. cated in light blue, whereas the design point is indicated by a red point. As can be seen in this figure, the wing loading is not constrained by the landing distance requirement. Similarly, the power loading is not limited by the take-off distance requirement. This means that the aircraft is able to land and take-off within 3,300 meters. The power loading is defined by the climb gradient requirement during landing in case of all engines operative (AEO).

0.5 Landing distance s=3300 m

0.45 Take-o, sTO=3300 m Cruise speed V= 154.031 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure 4.3: Wing power loading diagram of the multifuselage concept. 4.1. Multifuselage concept 43

Weight analysis

The weight breakdown of the multifuselage concept is given in Table 4.8. The MTOW of this aircraft concept is around 3,000 tonnes, which mainly consists of the OEW and the payload weight. The OEW and payload weight are 43% and 40% of the MTOW respectively, whereas the fuel fraction with respect to the MTOW is signiﬁcantly lower (17%). Due to the high energy content of hydrogen, only a small weight fraction with respect to the MTOW is found for the fuel weight. A detailed weight breakdown of the OEW can be found in Appendix D.

Table 4.3: Weight breakdown of the multifuselage concept.

Parameter Symbol Value Unit Maximum take-off weight MTOW 3,004,081 kg Operational empty weight OEW 1,277,505 kg Payload weight PW 1,200,000 kg Fuel weight FW 526,321 kg

Aerodynamic analysis

The aerodynamic performance of the multifuselage concept is given in Table 4.4. As can be seen in this table, the aircraft flies at an angle of attack of 7.4 degrees during cruise condition with a L/D ratio of 22.4. This relatively high angle of attack for cruise condition is mainly caused due to the low cruise speed and the high payload weight requirement. As a result, the total drag of the aircraft is mainly influenced by the induced drag due to the high lift coefficient. The lift distribution in cruise condition is shown in Figure 4.4. As can be seen in this figure, the lift distribution is elliptical along the wing span, whereas the maximum lift coefficient occurs between 75% and 80% of the wing span. With respect to this position, the 2D lift curve slope in clean condition is shown in Figure 4.5. As can be seen in this figure, stall occurs at an angle of attack higher than 16 degrees, whereas the take-off and landing performance of the aircraft are analyzed for an angle of attack of 15 degrees. As can be seen in Table 4.4, the increase in maximum lift coefficient due to the high-lift devices for take-off and landing condition is 0.44 and 0.57, respectively.

Table 4.4: Aerodynamic properties of the multifuselage concept.

Parameter Symbol Value Unit Angle of attack Æ 7.4 ° Lift coefﬁcient (cruise) CL,cruise 0.67 - Drag coefﬁcient (cruise) CD,cruise 0.0297 - Lift-over-drag ratio L/Dcruise 22.4 -

Induced drag coefﬁcient CDi 0.0212 -

Proﬁle drag coefﬁcient CD0 0.00843 -

Maximum lift coefﬁcient (clean) CLmax,clean 1.58 -

Maximum lift coefﬁcient (take-off) CLmax,take-off 2.02 -

Maximum lift coefﬁcient (landing) CLmax,landing 2.15 -

Stability analysis

The CoG position moves during the mission due to the fuel consumption. The CoG position for the four conditions (MTOW, cruise, ZFW and OEW) are stated in Table 4.5. As can be seen in this table, the CoG shifts backwards during the mission. The most forward CoG position is for a fully loaded aircraft (MTOW), whereas the most backward position is for an empty aircraft (OEW). Furthermore, it can be seen in Table 4.5 that the neutral point is behind the CoG position for the different conditions except of the OEW. This means that the aircraft has to be loaded with a minimum amount of containers to meet the longitudinal stability requirement. The CoG position at ZFW is taken as CoG position for the longitudinal stability requirement, resulting in a static margin of 10.1%. 44 4. Aircraft Design and Performance

30 0.8

0.75 25 0.7

20 0.65 l l

C 0.6 c*C 15 0.55

0.5 10 0.45

5 0.4 0 20 40 60 80 100 120 0 20 40 60 80 100 120 Spanwise position Spanwise position (a) c C (b) C · l l

Figure 4.4: Lift distribution along the wing span of the multifuselage concept..

1.5 l

C 1

0.5

0 0 5 10 15 20 Angle of Attack

Figure 4.5: Airfoil lift coefﬁcient as a function of the angle of attack for the multifuselage concept.

Next to the static margin, the lateral directional stability and controllability parameters are defined in Table 4.5. As can be seen in this table, the multifuselage concept produces an restoring rolling moment in case of a small side slip angle, because it has a negative value for ClØ . With respect to the yawing moment in case of a small side slip angle, the multifuselage concept is neutrally stable. This means that the aircraft does not restore to its original position, but it also does not become unstable. Finally, it can be seen that the maximum rudder deflection in case of OEI is 14.7 degrees, which is less than 25 degrees defined by the constraint analysis module.

Table 4.5: Stability and controllability of the multifuselage concept.

Parameter Symbol Value Unit

Center of gravity at MTOW xCGMTOW 14.2 % MAC

Center of gravity at cruise xCGcruise 17.3 % MAC

Center of gravity at ZFW xCGZFW 20.7 % MAC

Center of gravity at OEW xCGOEW 56.3 % MAC Neutral point xnp 30.9 % MAC Static margin SM 10.1 % MAC

Static lateral directional stability derivative CnØ 0.0 -

Weather cock stability derivative ClØ -0.035 - Rudder deﬂection ±rudder 14.7 ° 4.2. Blended-wing-body concept 45

4.2. Blended-wing-body concept Similar to the multifuselage concept, ﬁrst an insight into the design space of the BWB concept is given. Sub- sequently, the aircraft design and performance in terms of aircraft dimensions, weight analysis, aerodynamic analysis and stability analysis are discussed in detail.

Design space Similar to the multifuselage concept, the influence of the design variables on the surrogate models of the objective function and the constraint functions is evaluated in order to give insight into the design space of the BWB concept. The initial design vector at the start of the numerical optimization is selected as baseline configuration, which is given in Table 4.6. This design vector corresponds to the feasible conceptual design with the lowest fuel weight of all candidate designs analyzed during the DOE. The influence of the design variables on the surrogate models of the objective function and the constraint functions are visualized in Figure 4.6. A violation of the constraint functions is illustrated in red, which indicates the infeasible region of the design space.

Table 4.6: Initial design vector of BWB concept for numerical optimization.

Parameter Symbol Value Lower bound Upper bound Unit Wing loading W/S 2865 2500 3100 N/m2 Aspect ratio AR 4.4 3 6 - Taper ratio section 2 ∏2 0.41 0.4 0.6 - Sweep angle of section 3 §LE 10.0 0 20 ° Ratio span section 2 over span section 2-3 b fraction 0.33 0.25 0.4 - Ratio vertical tail area over wing area SV /S 0.05 0.03 0.06 -

As can be seen in Figure 4.6(a), the fuel weight is mainly influenced by the wing loading, b fraction, the sweep angle of the outer wing, and the aspect ratio. Similar to the multifuselage concept, the aspect ratio has an optimum with respect to the fuel weight because of the same reasoning. Furthermore, an optimum can be found for the wing loading. An increase in wing loading results in a smaller and lighter wing, whereas the L/D ratio decreases due to an increase in induced drag. This can be explained by making use of Equation 4.1.A reduction in surface area results in a higher lift coefficient and the induced drag increases quadratically with the lift coefficient. Due to the increase in induced drag, the L/D decreases for a higher wing loading.

2 CL CD (4.1) i = ºARe

Moreover, the fuel weight is minimized for a minimum sweep angle and a maximum b fraction. An increase of b fraction corresponds to an increase in the wing span of the inner wing section. Because of the zero sweep at the trailing edge of the inner wing section, the leading edge sweep decreases at the inner wing section for an increase of the inner wing span and a fixed taper ratio. As a result, the structural wing weight decreases due to the a reduction in sweep angle. Usually, the leading edge sweep of the wing is determined to minimize the wave drag in transonic conditions. However, because of the low subsonic cruise speed for this aircraft concept, compressibility effects are not an issue. As a result, a reduction in sweep angle (both inner and outer wing) results in a lower fuel weight. As can be seen in Figures 4.6(b)-4.6(h), the feasible design space of the numerical optimization is limited by the cargo and fuel volume constraint, the longitudinal stability constraint, and the maximum rudder deflection in case of OEI. The volume constraints of the containers and the fuel tank are mainly influenced by the wing loading and the taper ratio, respectively. The wing surface has to be large enough to provide sufficient volume for the placement of the containers, whereas an increase in taper ratio results in a smaller chord length at the center line of the aircraft, c1. As a result, less volume is available for the fuel tank and the constraint becomes positive. Another consequence of a smaller c1 is that more containers are placed outboard, which leads to an aft movement of the CoG. Therefore, the static margin becomes smaller than 5% for a too large taper ratio. The static margin is also influenced by the sweep angle. If the sweep of the outer wing increases, the aerodynamic center moves aft resulting in a larger static margin. 46 4. Aircraft Design and Performance

In terms of lateral stability and controllability, the design space of the numerical optimization is limited by the constraint to ensure a maximum rudder deﬂection of 25 degrees in case of OEI. This constraint is mainly inﬂuenced by the size of the vertical tail, namely by SV /S and W/S. The static directional stability derivative,

CnØ , is always positive and mainly influenced by the SV /S, taper ratio, and aspect ratio. If the taper ratio and the aspect ratio increases, the root chord length (c1) decreases. As a result, the distance between the rudder and the CoG decreases and the rudder effectiveness reduces. As can be seen in Figure 4.6(e), the effective dihedral derivative, ClØ , is mainly influenced by by the outer sweep angle. However, ClØ is always negative and therefore inactive during the optimization. Finally, the maximum wing loading constraint is inactive during the optimization because the wing loading is already limited by the cargo volume requirement. In conclusion, the design space of this aircraft concept is limited by several constraints during the optimization. First of all, the wing loading cannot be too large to ensure enough volume into the wing for the placement of the fuel tank. Next, the taper ratio of the inner wing section is constrained to provide sufficient volume for the placement of the tank and to ensure longitudinal stability. Moreover, a minimum sweep angle of the outer wing section is required to provide longitudinal stability. Finally, the vertical tail area has to be large enough to provide sufficient lateral controllability in case of OEI. The final design vector found by the numerical optimization is given in Table 4.7.

Table 4.7: Final design vector of multifuselage concept for numerical optimization.

Parameter Symbol Value Lower bound Upper bound Unit Wing loading W/S 2928 2500 3100 N/m2 Aspect ratio AR 5.1 3 6 - Taper ratio section 2 ∏2 0.4 0.4 0.6 - Sweep angle of section 3 §LE 9.6 0 20 ° Ratio span section 2 over span section 2-3 b fraction 0.38 0.25 0.4 - Ratio vertical tail area over wing area SV /S 0.04 0.03 0.06 - 4.2. Blended-wing-body concept 47

#105 7.5 0.2 AR W/S taper 7 sweep 0 b fraction Sv/S 6.5 -0.2

6 -0.4

Fuel Weight [kg] 5.5 -0.6 AR W/S Cargo volume constraint [-] taper 5 -0.8 sweep b fraction Sv/S 4.5 -1 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

(a) Fuel weight (b) Cargo volume

0.3 12 AR W/S 10 0.25 taper sweep b 8 0.2 fraction Sv/S 6 0.15 4

0.1 2

0 0.05 Static Margin [%] -2 AR 0 Fuel volume constraint [-] W/S -4 taper sweep -0.05 b -6 fraction Sv/S -0.1 -8 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

0.02 0.06 AR 0.01 W/S taper 0.05 sweep 0 b fraction -0.01 Sv/S 0.04

-0.02 0.03 [-] [-] -0.03 l,beta n,beta C

C 0.02 -0.04

-0.05 0.01 AR W/S -0.06 taper 0 sweep b -0.07 fraction Sv/S -0.08 -0.01 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

(e) Effective dihedral derivative (f) Static lateral stability derivative

32 0.05 AR 30 W/S 0 taper sweep b 28 fraction -0.05 Sv/S 26 -0.1 AR 24 -0.15 W/S taper sweep b 22 -0.2 fraction Sv/S 20 -0.25 WSmax constraint [-] Rudder deflection [deg] 18 -0.3

16 -0.35

14 -0.4 0 0.2 0.4 0.6 0.8 1 0 0.2 0.4 0.6 0.8 1 Lower bound to upper bound Lower bound to upper bound

(g) Rudder deﬂection in case of OEI (h) Maximum wing loading constraint

Figure 4.6: Surrogate models on objective function and constraint functions for BWB concept. 48 4. Aircraft Design and Performance

Aircraft design Based on the final design vector given in Table 4.7, a conceptual design is generated and analyzed by making use of the MDA system. A geometric overview of the BWB concept is illustrated in Figure 4.7. This aircraft concept has a wing span of 218 meters and a length of 86.5 meters. The maximum thickness of the aircraft is 13.8 meters at the center line of the aircraft where the fuel tank is located. Two vertical tails with a span of 22.5 meters are positioned at the trailing edge of the airfoil. Similar to the multifuselage concept, the propulsion system also consists of six contra-rotating turboprop engines where the propeller radius is 7.1 meters. The design point of this aircraft concept in the wing power loading diagram is illustrated in Figure 4.8. Simi- lar to the multifuselage concept, the wing loading and the power loading do not exceed the landing distance and take-off distance requirement in the wing power loading diagram. As a result, this concept is also able to land and take-off within 3,300 meters. Similar to the multifuselage concept, the power loading is defined by the climb gradient requirement during landing in case of AEO, where the wing loading of the BWB concept is significantly lower compared to the multifuselage concept. As discussed in the former section, a low wing loading is required to provide sufficient volume for the placement of the containers into the wing. An overview of the container distribution in the aircraft is given in Figure 4.9. The placement of the fuel tank and the containers in side view orientation is given in Figure 4.10. As can be seen in Figure 4.10, all the containers and the fuel tank are placed between the front and rear spar. The thickness and the chord length gradually reduces along the wing span. As a result, the number of containers in each row decreases along the wing span.

(a) Top View (b) Front View

Figure 4.7: A three-dimensional overview of BWB concept. 4.2. Blended-wing-body concept 49

0.5 Landing distance s=3300 m

0.45 Take-o, sTO=3300 m Cruise speed V= 154.031 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure 4.8: Wing power loading diagram of the BWB concept.

-10

-20

-30

-40

Y-axis [m] -50

-60

-70

-80

-90 0 20 40 60 80 100 120 X-axis [m]

Figure 4.9: Overview of the container distribution. 50 4. Aircraft Design and Performance

30 30 30 30 20 20 20 20 10 10 10 10

t [m] 0 t [m] 0 t [m] 0 t [m] 0

-10 -10 -10 -10 -20 -20 -20 -20 -30 0 20 40 60 80 0 20 40 60 80 0 20 40 60 0 20 40 60 c [m] c [m] c [m] c [m] (a) Fuel Tank (b) Row 1 (c) Row 2 (d) Row 3

30 20 20 20 20

10 10 10 10

0 0 t [m] 0 t [m] 0 t [m] t [m]

-10 -10 -10 -10

-20 -20 -20 -20 0 20 40 60 0 20 40 60 0 20 40 60 0 20 40 60 c [m] c [m] c [m] c [m] (e) Row 4 (f) Row 5 (g) Row 6 (h) Row 7

20 20 20 15 15 15 10 10 10 10 5 5 5 0 0 0 0 t [m] t [m] t [m] t [m] -5 -5 -5 -10 -10 -10 -10 -15 -15 -20 -15 10 20 30 40 50 0 20 40 0 10 20 30 40 0 10 20 30 40 c [m] c [m] c [m] c [m] (i) Row 8 (j) Row 9 (k) Row 10 (l) Row 11

15 15 10 10 10 10 5 5 5 5

0 0 0 0 t [m] t [m] t [m] t [m]

-5 -5 -5 -5 -10 -10 -10 -10 -15 0 10 20 30 40 0 10 20 30 0 10 20 30 0 10 20 30 c [m] c [m] c [m] c [m] (m) Row 12 (n) Row 13 (o) Row 14 (p) Row 15

10 10 10 10

5 5 5 5

0 0 0 0 t [m] t [m] t [m] t [m]

-5 -5 -5 -5

-10 -10 -10 -10 0 10 20 30 0 10 20 30 0 10 20 30 0 10 20 c [m] c [m] c [m] c [m] (q) Row 16 (r) Row 17 (s) Row 18 (t) Row 19

0 t [m]

-5

-10 0 10 20 c [m] (u) Row 20

Figure 4.10: Cut-outs of the aircraft to show the placement of the fuel tank and the containers, starting from the center line towards the wing tip.

Weight analysis

The weight breakdown of the aircraft is given in Table 4.8. The BWB concept has a MTOW of 2,871,834 kilograms, which is slightly lower compared to the multifuselage concept. Similar to the multifuselage concept, the MTOW mainly consists of the OEW (41%) and the payload weight (42%). For this aircraft concept, the payload weight is even slightly larger than the OEW. The fuel weight fraction with respect to the MTOW is similar 4.2. Blended-wing-body concept 51 for this aircraft concept compared to the multifuselage concept, namely 17%. A detailed weight breakdown of the OEW for this aircraft concept can be found in Appendix E.

Table 4.8: Weight breakdown of the BWB concept.

Parameter Symbol Value Unit Maximum take-off weight MTOW 2,871,834 kg Operational empty weight OEW 1,185,164 kg Payload weight PW 1,200,000 kg Fuel weight FW 486,415 kg

Aerodynamic analysis The aerodynamic properties of the BWB concept are given in Table 4.9. As can be seen in this table, the aircraft flies at a high angle of attack of 7.0 degrees with a L/D ratio of 24.1. Similar to the multifuselage concept, the total drag is mainly influenced by the induced drag. The lift distribution along the wing span is shown in Figure 4.11(a). At the inner wing section the lift distribution is almost constant, whereas an increase in c C is · l visible at the intersection between the inner wing and the outer wing. This increase is caused by the change in sweep angle, which is significantly lower for the outer wing compared to the inner wing as shown in Figure 4.7(a). Finally, the lift distribution decreases along the wing span of the outer wing section. As can be seen in Figure 4.11(b), the stall position is located very far outboard. This is dangerous because a loss of lift and increase of drag at the tip of the wing result in a sudden rolling moment of the aircraft. Furthermore, the ailerons are usually located close to the wing tip which become ineffective if stall occurs near the wing tip. Therefore, the elevons have to be used to counteract the rolling moment if stall occurs at the wing tips. The 2D lift curve slope in clean condition at the stall position is shown in Figure 4.12. Similar to the multifuselage concept, stall starts at an angle of attack of 16 degrees, whereas the maximum lift coefficient for take-off and landing performance is estimated for an angle of attack of 15 degrees. Because of the absence of high-lift devices, the maximum lift coefficient in clean condition is equal to the maximum lift coefficient for take-off and landing condition.

Table 4.9: Aerodynamic properties of the BWB concept

Parameter Symbol Value Unit Angle of attack Æ 7.0 ° Lift coefﬁcient (cruise) CL,cruise 0.43 - Drag coefﬁcient (cruise) CD,cruise 0.0178 - Lift-over-drag ratio L/Dcruise 24.1 -

Induced drag coefﬁcient CDi 0.01191 -

Proﬁle drag coefﬁcient CD0 (Total) 0.00588 -

Maximum lift coefﬁcient (clean) CLmax,clean 1.60 -

Maximum lift coefﬁcient (take-off) CLmax,take-off 1.60 -

Maximum lift coefﬁcient (landing) CLmax,landing 1.60 -

Stability analysis The shift of the CoG position during the mission as a precentage of the MAC is given in Table 4.10. Similar to the multifuselage concept, the most forward and aft CoG position are at MTOW and OEW, respectively. As can be seen in Table 4.10, the neutral point is located aft with respect to all CoG positions. However, the minimum static margin of 5% to ensure longitudinal stability is valid for all conditions except of the OEW condition. Therefore, this concept also requires a minimum amount of containers or a ballast in order to meet the longitudinal stability requirement. Besides the longitudinal stability, the BWB concept is also stable in lateral direction. As can be seen in Table

4.10, the aircraft has a positive CnØ and a negative ClØ . With respect to controllability, a rudder deﬂection of 25 degrees is required in case of OEI. As a result, the BWB concept meet all the stability and controllability requirements deﬁned by the constraint analysis module. 52 4. Aircraft Design and Performance

25 1

0.9

20 0.8

0.7 l l

15 C 0.6 c*C 0.5

10 0.4

0.3

5 0.2 0 20 40 60 80 100 120 0 20 40 60 80 100 120 Spanwise position Spanwise position (a) c C (b) C · l l

Figure 4.11: Lift distribution along the wing span of the BWB concept.

1.8

1.6

1.4

1.2

l 1 C

0.8

0.6

0.4

0.2

0 0 2 4 6 8 10 12 14 16 18 20 Angle of Attack

Figure 4.12: Airfoil lift coefﬁcient as a function of the angle of attack for the BWB concept.

Table 4.10: Stability and controllability of the BWB concept.

Parameter Symbol Value Unit

Center of gravity at MTOW xCGMTOW 21.5 % MAC

Center of gravity at cruise xCGcruise 24.4 % MAC

Center of gravity at ZFW xCGZFW 27.5 % MAC

Center of gravity at OEW xCGOEW 31.0 % MAC Neutral point xnp 32.5 % MAC Static margin SM 5.0 % MAC

Static lateral directional stability derivative CnØ 0.028 -

Weather cock stability derivative ClØ -0.054 - Rudder deﬂection ±rudder 25.0 °

4.3. Design and performance comparison between aircraft concepts

In order to compare the aircraft concepts with each other, an overview of the main characteristics with respect to the aircraft design and performance of both aircraft concepts is given in Table 4.13. The difference of each parameter indicated in this table is defined as the difference of the BWB concept with respect to the multifuselage concept. The aircraft concepts are compared on the following criteria: the aircraft dimensions, the aircraft weight including the fuel weight required to fulfill the mission, the aerodynamic performance during cruise condition, the stability & controllability, and the transport efficiency. With respect to the transport efficiency, a metric is defined to compare the transport efficiency of both aircraft concepts also to current 4.3. Design and performance comparison between aircraft concepts 53 large cargo aircraft and cargo ships. Finally, conclusions are drawn based on the comparison of the aircraft concepts on these criteria.

4.3.1. Aircraft dimensions As can be seen in Table 4.13, the wing span of the BWB concept is slightly larger than the wing span of the multifuselage concept, whilst the length of the multifuselage concept is signiﬁcantly larger than the length of the BWB concept. This difference in aircraft dimensions is illustrated in Figure 4.13. As a result, the apron area required for the parking of the multifuselage concept is larger than for the BWB concept. The vertical height of the vertical tail is comparable of both concepts.

Regarding the propulsion system, both aircraft concepts have six engines with a propeller radius of approximately 7 meters and a propulsive efficiency of 0.74. As can be seen in Figure 4.13(d), the BWB concept requires a tall landing gear system to ensure sufficient clearance for these large engines, whilst the landing gear system of the multifuselage concept is significantly lower because of the high wing configuration.

(a) Front View - Multifuselage (b) Side View - Multifuselage

Figure 4.13: Illustration of geometric characteristics of both concepts.

4.3.2. Aircraft weight In terms of fuel weight required to fulfill the mission, the BWB concept requires 8.2% less fuel weight compared to the multifuselage concept. This reduction in fuel weight is caused by two reasons. Firstly, the OEW of the BWB concept is 7.8% lower compared to the multifuselage concept. As can be seen in Table 4.13, this reduction is mainly caused by a significant weight saving due to the absence of the fuselages. The wing weight of the BWB concept is higher, because of the large wing surface required to store the fuel tank and the containers. Due to the distributed loading of the containers along the wing span of the BWB concept, the difference 54 4. Aircraft Design and Performance in wing weight is small between the wing weight of the BWB concept and the multifuselage concept. On the other hand, the landing gear weight is significantly larger for the BWB concept due to the tall landing gear system.

The second reason for the reduced fuel weight of the BWB concept compared to the multifuselage concept is the better aerodynamic performance of the BWB concept in cruise condition, where the L/D ratio is 7.1% higher for the BWB concept. The aerodynamic performance of both aircraft concepts is discussed in more detail in the next paragraph.

4.3.3. Aerodynamic performance

As stated in the former paragraph, the aerodynamic performance of the BWB concept is higher in cruise condition compared to the multifuselage concept. In order to compare the main contributions to the L/D ratio of both concepts, the aerodynamic coefficients are multiplied by the wing surface area. As can be seen in Table 4.13, the BWB concept flies at a lower scaled lift coefficient and a lower angle of attack during cruise. In terms of drag, the scaled zero-lift drag is slightly larger for the BWB concept (3.5%) due to the large wing surface area, whereas the induced drag is 21% lower compared to the multifuselage concept. Due to this significant reduction in induced drag, the BWB concept has a higher L/D ratio during cruise condition with respect to the multifuselage concept.

4.3.4. Stability and controllability

In terms of longitudinal and lateral directional stability, both aircraft concepts are stable. The multifuselage concept has a higher static margin compared to the BWB concept meaning that the multifuselage concept is more stable in longitudinal direction, although both concepts require a minimum amount of containers or a ballast to ensure longitudinal stability along the entire mission. In terms of lateral directional stability, the multifuselage concept is neutrally stable in terms of CnØ , whereas the BWB concept generates a restoring yawing moment in case of a small side slip angle.

Regarding lateral controllability, both aircraft concepts are able to counteract the yawing moment generated in case of OEI with a rudder deflection smaller than 25 degrees. The multifuselage concept only requires a rudder deflection of 14.7 degrees, whilst the BWB concept requires a rudder deflection of 25 degrees.

4.3.5. Transport efﬁciency

Finally, the transport efficiency of both aircraft concepts are quantified and compared to each other by the Payload Fuel Energy Efficiency (PFEE) metric. This metric is able to compare the transport efficiency of different transportation modes designed for different mission requirements in terms of range and payload weight. By evaluating the amount of energy required to transport a certain amount of payload weight over the great circle distance, this metric is also able to compare transportation modes powered by different fuels. In mathematical terms, this metric is defined by Equation 4.2. The energy consumed during a mission, E fuel, is the product of the fuel weight and the heat of combustion. The specific heat of combustion of hydrogen and kerosene are taken as 120 MJ/kg and 42.8 MJ/kg, respectively [10]. The transport efficiency in terms of PFEE of the two proposed aircraft concepts are compared to each other and the Boeing 747-8F in Table 4.11. As can be seen in this table, the multifuselage concept performs slightly better (1.4%) to this reference aircraft in terms of PFEE, whereas the BWB concept is 8.6% more transport efficient compared to the Boeing 747- 8F.

Wpay R PFEE · (4.2) = E fuel

This metric can also be used to compare the performance of this new aircraft concept to maritime transport. Based on the data provided by Notteboom and Vernimmen [65], the transport efﬁciency of a cargo ship used to transport containers along the Europe-Far East trade line is determined. The trip including the sailing time is given in Table 4.12. The average vessel size for this route is around 8000 TEU, where the longest 4.3. Design and performance comparison between aircraft concepts 55

Table 4.11: Comparison of different transportation modes in terms of PFEE.

Transportation mode R Wpay Wfuel PFEE 1 [km] [kg] [kg] [kg km MJ° ] · · Blended-wing-body concept 7,400 1,200,000 486,415 152 Multifuselage concept 7,400 1,200,000 526,321 141 Boeing 747-8F [28] 8,130 132,630 181,535 139 Container ship [65] 10,544 96,000,000 3,262,500 7249 distance is between Singapore and Rotterdam with a great circle distance of 6552 miles 1. The commercial speed of this container ship is around 24 knots resulting in a sailing time of 14.5 days and an average fuel consumption of 225 tonnes kilogram of kerosene per day [65]. Based on this data, the PFEE of a typical cargo ship for intercontinental transport of containerized goods along the Europe-Far East trade line is 7249 1 kg km MJ° . As a result, the transport efficiency is around 47 to 51 times larger compared to the BWB · · concept and the multifuselage concept, respectively. However, the transportation time is not included in this transport efficiency metric, where the hydrogen powered ultra large cargo aircraft is significantly faster than the cargo ship. As can be seen in Table 4.12, the transportation time of a cargo ship from Shanghai to Rotterdam is approximately 21 days [65], where the transportation time of the hydrogen powered ultra large cargo aircraft is approximately one day for the same route [15].

Table 4.12: Total sailing time for a container ship between Singapore and Rotterdam [65]

Trip Distance (nm) Sailing time (days) Shanghai - Dalian 576 1.00 Dalian - Qingdao 280 0.49 Qingdao - Ningbo 512 0.89 Ningbo - Singapore 2143 3.72 Singapore - Rotterdam 8353 14.50 Total 11,864 20.6

4.3.6. Conclusions Regarding the aircraft dimensions, both aircraft concepts have a wing span of larger than 200 meters. Current regulations for large commercial airports state a maximum wing span of 80 meters and a MTOW of approximately 600,000 kilograms to satisfy pavement loading requirements [2]. Both the wing span as well as the MTOW of the hydrogen powered ultra large cargo aircraft excessively exceed these current airport regulations. Therefore, new dedicated airports have to be build for both aircraft concepts. These airports require a runway which is able to handle the heavy aircraft weight for landing and take-off. Furthermore, a sufﬁcient apron area is required for loading/unloading the containers and to fuel the aircraft. With respect to hydrogen as aviation fuel, new facilities close to the airport are required for the production, storage and transportation of hydrogen. Nowadays, hydrogen is already produced on a large worldwide scale where the United States has a production capacity of more than 50 million tonnes of hydrogen [66]. The most popular technique to produce hydrogen in large quantities is steam reforming of natural gas. A drawback of this production technique is the emission of green house gases (CO2) and the dependency on fossil fuels. Hydrogen can also produced by splitting water into hydrogen and oxygen by means of electrolysis. The beneﬁt of the latter method is that it does not rely on fossil fuels, because the energy required for electrolysis can be delivered by renewable energy sources, such as solar energy, wind power, hydro power or biomass [8]. Furthermore, no green house gases are emitted during the production of hydrogen. In order to become a sustainable alternative to maritime transport, facilities have to be build which are able to transform the renewable energy sources into electricity for the electrolysis process. As a result, large investment costs are required for the development of new airports including the facilities to produce, store and transport hydrogen as fuel for the ultra large cargo aircraft. Hordijk [15] investigates the economical feasibility to use onshore wind energy for the electrolysis process to produce hydrogen as aviation fuel for this aircraft concept.

1Swartz, Karl L., Great Circle Mapper, http://www.gcmap.com (accessed 14-12-2016) 56 4. Aircraft Design and Performance

Designed for the same TLRs, it can be concluded that the BWB concept has a higher transport efficiency in terms of PFEE with respect to the multifuselage concept. This means that less fuel weight is required to fulfill the mission requirements stated in Section 1.4. As a result, the operational costs of the BWB concept are expected to be lower compared to the multifuselage concept. Furthermore, it can be concluded that the OEW of the BWB concept is lower compared to the multifuselage concept because of the absence of the heavy fuselages. On the other hand, the wing span of the BWB concept is larger compared to the multifuselage concept. These conclusions are drawn based on the characteristics of the conceptual designs, which are designed for the TLRs defined in Section 1.4. The sensitivity of the cruise speed, cruise altitude, and the payload weight on the aircraft design and performance for both aircraft concepts is evaluated in the next chapter. 4.3. Design and performance comparison between aircraft concepts 57

Table 4.13: Performance overview of BWB concept and multifuselage concept

Parameter Symbol Unit BWB Multifuselage Difference Wing loading W/S N/m2 2928 4566 -54% 2 Maximum wing loading W /Smax N/m 4404 5912 -34% Power loading W/P - 0.066 0.062 6.1% Aspect ratio AR - 5.1 6.8 -33% Wing span b m 218 206 5.5% Span horizontal tail bHT m - 50.1 - Span vertical tail bVT m 22.5 20.3 9.8% Length L m 86.5 150 -73% Diameter fuselage fus m 13.8 13.6 1.4% Æ 2 Planform area Sref m 9256 6280 32% 2 Horizontal tail area SHT m - 628 - 2 Vertical tail area SVT m 389 316 21% Maximum take-off weight MTOW kg 2,871,834 3,004,081 -4.6% Operational empty weight OEW kg 1,185,164 1,277,505 -7.8% Payload weight PW kg 1,200,000 1,200,000 - Fuel weight FW kg 486,415 526,321 -8.2% Wing weight Wwing kg 816,729 731,623 10% Fuselage weight Wfus kg - 221,075 - Horizontal tail weight WHT kg - 40,139 - Vertical tail weight WVT kg 23,564 30,031 -27% Landing gear weight Wlg kg 156,722 51,332 67% Power plant weight Wpp kg 69,083 74,902 -8.4 Fuel tank weight Wft kg 65,334 81,210 -24% Fixed equipment Wfe kg 53,727 47,192 -12% Lift-over-drag ratio L/Dcruise - 24.1 22.4 7.1% Angle of attack Æ ° 7.0 7.4 -5.7% Cruise lift coefficient CL,cruise - 0.43 0.67 -56% Cruise drag coefficient CD,cruise - 0.0178 0.0297 -67% Cruise lift coefficient scaled C S m2 3980 4208 -5.7% L,cruise · ref Cruise drag coefficient scaled C S m2 164.8 186.5 -13% D,cruise · ref Induced drag coefficient scaled C S m2 110.1 133.3 -21% Di · ref Zero-lift drag coefficient scaled C S m2 54.6 52.7 3.5% D0 · ref Maximum lift-coefficient (clean) CLmax,clean - 1.60 1.58 1.3%

Maximum lift-coefﬁcient (TO) CLmax,TO - 1.60 2.02 -26%

Maximum lift-coefﬁcient (Landing) CLmax,landing - 1.60 2.15 -34% Propeller radius rprop m 7.1 7.5 -5.6% Propeller efﬁciency ¥p - 0.74 0.74 0.0%

Center of gravity at MTOW xCGMTOW % MAC 21.5 14.2 -

Center of gravity at Cruise xCGCruise % MAC 24.4 17.3 - Center of gravity at ZFW xCGZFW % MAC 27.5 20.7 - Center of gravity at OEW xCGOEW % MAC 31.0 56.3 - Neutral point xnp % MAC 32.5 30.9 - Static margin SM % 5 10 100% Effective dihedral derivative C - -0.054 -0.035 35% lØ

Static lateral stability derivative CnØ - 0.028 0.0 - Rudder deﬂection (OEI) ±rudder ° 25.0 14.7 41%

Cargo volume requirement Vcargoa /Vcargoreq - 1.0 - -

Fuel volume requirement Vfuela /Vfuelreq - 1.0 - - Payload Fuel Energy Efﬁciency PFEE kg km MJ 1 152 141 7.2% · · ° 58 4. Aircraft Design and Performance

4.4. Design and performance comparison to aircraft concept of DSE As mentioned in Section 2.1.2, ten students investigated also the design and performance of a hydrogen powered ultra large cargo aircraft for the same TLRs stated in Section 1.4 during the 2015 DSE [19]. The aircraft concept proposed by this project is a Burnelli configuration where all the payload and fuel is stored in a large lifting airfoil shaped body. The aircraft has a wing span of 200 meters and a length of 100 meters. Because of the limited time frame of 10 weeks for this preliminary research, mostly empirical data was used to design and analyze this aircraft concept. Very promising performance estimations were made for this aircraft concept in order to meet the performance target of a freight rate less than 250% of the freight rate for maritime transport. This section discusses the comparison of the main outcome of the aircraft concept of the DSE project with the BWB concept considered for this study. An overview of the main differences between the aircraft concept of the DSE project and the BWB concept considered for this study are given in Table 4.14. Regarding the aerodynamic performance estimation of the DSE project, a L/D ratio of 26 during cruise condition was estimated for this aircraft concept, where the wing is composed of different thick cambered NACA 4 series airfoils with a wing span of 200 meters. The aerodynamic analysis is estimated by making use of XFLR5, which is an analysis tool for airfoils, wings and planes [67]. However, several drag contributions were not taken into account, which resulted in an overestimation of the L/D ratio of this aircraft concept. First of all, the drag contribution of the engine nacelles and the empennage are not included in the aerodynamic performance estimation. Furthermore, the trim drag associated to trim the aircraft in cruise condition is not taken into account by XFLR5. As a result, the L/D ratio of 26 is overestimated by the DSE project due to the underestimation of the drag. As a comparison, the L/D of the BWB concept considered for this research is 24.1. Furthermore, the estimation of the fuel weight required to fulfill the mission was also very promising. The estimation of the fuel weight for the aircraft concept of the DSE project was 141,000 kilograms with a MTOW of 2,092,000 kilograms for this aircraft concept. The significant weight difference of the fuel weight and the MTOW between the DSE project and this research is caused by two reasons. First of all, the aircraft concept considered for the DSE project was designed for the initial range of 6,000 kilometers at maximum payload weight (100 TEU containers), whilst the aircraft concepts for this study are designed for a range of 7,400 kilometers as discussed in Section 1.4. The second reason is the difference of the weight analysis methods. Very crude estimations and approximations were made to determine the weight of the aircraft components during the DSE project, where semi-analytical and semi-empirical methods were considered for this study. Except of the wing, the weight of the main components of the aircraft concept of the DSE project were estimated based on weight fractions with respect to the MTOW derived from statistical data of existing aircraft [68]. With respect to the wing weight, a simplified structural analysis has been performed to estimate the structural weight of the wing box required to withstand the loads acting on the wing. However, the wing box geometry have been simplified to a rectangular-shaped box, where the sharp kink of the wing and the airfoil shape are not taken into account. Furthermore, a simple correction factor was applied to estimate the total secondary wing weight. In contrast to these empirical methods, the conceptual design framework developed for this research makes use of semi-analytical and semi-empericial methods where the actual geometry of the aircraft components are taken into account to estimate the weight of these components.

Table 4.14: Comparison between aircraft concept of DSE project [19] and BWB concept considered for this study

Parameter Symbol Unit BWB DSE Difference Wing span b m 218 200 8.3% Length L m 86.5 100 -16% Maximum take-off weight MTOW kg 2,871,834 2,092,000 27% . Operational empty weight OEW kg 1,185,164 701,000 41% Payload weight PW kg 1,200,000 1,200,000 - Fuel weight FW kg 486,415 141,000 71% Lift-over-drag ratio L/Dcruise - 24.1 26.0 -7.9% Sensitivity Analysis

A sensitivity study is performed to investigate the effect of the following TLRs on the design and performance of both aircraft concepts: the payload weight, the cruise Mach number and the cruise altitude. Regarding the design and performance of the aircraft, the sensitivity of each TLR is evaluated on:

• Transport efﬁciency by means of the PFEE metric and the fuel weight required to fulﬁll the mission

• Aircraft weight by means of the OEW

• Aerodynamic performance by means of L/D

• Aircraft dimensions by means of the wing span

In order to evaluate the sensitivity of the payload weight, the cruise Mach number and the cruise altitude on these parameters, both aircraft concepts are designed and analyzed for an increase and decrease of 10% for each TLR with respect to the initial TLRs. The initial TLRs are a payload weight of 100 containers (1,200 tonnes), a cruise Mach number of 0.5 at a cruise altitude of 8,000 meters. An overview of the parameters selected for this sensitivity analysis including the perturbations is given in Table 5.1. Subsequently, the relative changes in aircraft design and performance with respect to the results presented in the former chapter are evaluated. The results of the sensitivity analysis for both aircraft concepts are given in Tables 5.2 & 5.3 and illustrated in Figures 5.1 & 5.2. A decrease of a TLR is visualized by a dark color, whereas an increase of a TLR is visualized by a light color.

Table 5.1: Parameters selected for the sensitivity study including the perturbations with respect to the baseline TLRs.

Parameter Symbol Unit -10% Baseline +10% Cruise Mach number Mcruise - 0.45 0.5 0.55 Amount of containers ncont - 90 100 110 Cruise altitude hcruise m 7200 8000 8800

5.1. Multifuselage concept

As can be seen in Figure 5.1, the transport efficiency of the multifuselage concept in terms of PFEE is mainly influenced by the cruise speed and the cruise altitude, where an increase in cruise speed and a decrease of cruise altitude result in a higher PFEE. The dynamic pressure increases for a higher cruise Mach number and a lower cruise altitude (higher density), which results in a lower lift coefficient. This is beneficial in terms of aerodynamic efficiency because the induced drag decreases quadratically with respect to a decrease of the lift coefficient. Due to the relatively high cruise lift coefficient (0.67), the induced drag is the largest contribution to the total drag (around 70%). As a result, the L/D increases for an increase in cruise speed and a decrease of cruise altitude. This effect is clearly visible in Figure 5.1. Moreover, it can be seen in Table 5.2 that the wing span slightly reduces for a higher cruise Mach number and lower cruise altitude, which results in a lower

59 60 5. Sensitivity Analysis wing weight and thereby a lower OEW. Based on these aspects, the fuel weight reduces significantly for a higher cruise speed and a lower cruise altitude, resulting in a higher PFEE. In terms of PFEE, a higher cruise Mach number is more effective than a lower cruise altitude. The PFEE increases 6.7% for a 10% increase in cruise Mach number, whereas the PFEE increases 4.2% for a 10% decrease in cruise altitude. As can be seen in Figure 5.1, the payload weight has an insignificant effect on the transport efficiency in terms of PFEE compared to the cruise speed and the cruise altitude. A reduction of number of containers results in a lower payload weight, and thereby a lower OEW and fuel weight. The OEW and fuel weight change proportionally to the payload weight resulting in a constant PFEE. On the other hand, the payload weight has a significant effect on the wing span for the multifuselage concept. The wing span is reduced by 5% if the aircraft is designed for 90 containers instead of 100 containers.

Table 5.2: Results of sensitivity study of multifuselage concept in percentages with respect to baseline design.

M [-] ncont [-] hcruise [m] 0.45 0.55 90 110 7200 8800 PFEE -10.9 6.7 -0.3 0.1 4.2 -5.9 FW 12.2 -6.3 -9.8 9.9 -4.0 6.3 OEW 4.5 -1.6 -9.7 11.4 -0.3 1.1 b 0.8 -1.4 -5.0 3.8 -0.5 -0.4 L/D -11.1 8.0 0.0 1.4 5.4 -7.1

5.2. Blended-wing-body concept Similar to the multifuselage concept, the transport efficiency of the BWB concept in terms of PFEE is mainly influenced by the cruise speed and the cruise altitude. A decrease of the cruise altitude results in an increase of aerodynamic efficiency, where the wing span remains constant. As discussed in the former section, the aerodynamic efficiency increases due to a reduction of the induced drag for a lower cruise altitude. As a result, the OEW remains constant due to the constant wing span where the fuel weight decreases due to the increase in aerodynamic efficiency. As a consequence, the transport efficiency in terms of PFEE increases for a lower cruise altitude. Regarding an increase in cruise speed, it can be seen in Figure 5.2 that the wing span decreases resulting in a reduction of the OEW. Despite of the decreased wing span, the aerodynamic remains almost constant for a higher cruise Mach number due to the increased dynamic pressure. As a result of the reduced OEW and the constant L/D for a higher cruise speed, the fuel weight decreases resulting in an increased PFEE. In Table 5.3, it can be seen that an increase of 10% of the cruise speed or a decrease of 10% of the cruise altitude results both in an increase of approximately 3.0% with respect to PFEE. The sensitivity of the amount of the containers on the transport efficiency in terms of PFEE is significantly lower compared to the effect of the cruise altitude and the cruise speed. The BWB concept becomes 0.6% more efficient for a payload weight of 90 containers, whereas the PFEE decreases 1.6% for a payload weight of 110 containers. As a result, the BWB concept becomes more efficient for a lower cruise altitude, a higher cruise Mach number and a lower payload weight.

Table 5.3: Results of sensitivity study of BWB concepts in percentages with respect to baseline design.

M [-] ncont [-] hcruise [m] 0.45 0.55 90 110 7200 8800 PFEE -11.8 2.9 0.6 -1.6 3.1 -5.8 FW 13.4 -2.9 -10.5 11.9 -2.9 6.3 OEW 9.8 -2.9 -8.3 9.7 -1.2 -2.0 b 2.1 -4.0 -1.3 1.4 -0.9 -4.2 L/D -9.5 0.4 2.9 -3.3 2.5 -10.0 5.3. Conclusions 61

5.3. Conclusions Based on the results of the sensitivity analysis the following conclusions can be made. First of all, it can be concluded that the transport efficiency of both concepts are mainly influenced by the cruise speed and the cruise altitude, where the sensitivity of the payload weight on the transport efficiency is small. Furthermore, it can be seen that the sensitivity of the cruise altitude on the transport efficiency is almost the same for both concepts, whereas an increase in cruise Mach number is around two times more effective for the multifuselage concept compared to the BWB concept. From the sensitivity analysis, it can be concluded that it is more efficient for this aircraft concept to fly faster and at a lower cruise altitude with respect to the initial TLRs defined at the start of this research. First of all, an increase in cruise speed results in a higher transport efficiency in terms of PFEE. Because less fuel weight is required to fulfill the mission, the operational costs decreases. Furthermore, the transportation time and thereby the crew salary also decreases with increasing cruise speed. Therefore, the competitiveness with respect to maritime transport increases for a higher cruise speed for this aircraft concept. An increase of the cruise speed also has an effect on the aircraft design. The sweep angle of the outer wing increases for both aircraft concepts to prevent a significant contribution of the wave drag to the total drag. Next to the sweep angle, the thickness of the wing is usually reduced to prevent an increase in wave drag for a higher cruise Mach number. However, the thickness distribution of the wing is not a design variable and therefore kept constant during the sensitivity analysis. As a result, the effect of a reduction of wing thickness for an increase in cruise speed is not taken into account for this sensitivity analysis. Furthermore, the type of propulsion system could be affected by an increase of the cruise speed. Current contra-rotating turboprop engines have a high propulsive efficiency up to a Mach number of around 0.65 [14]. Therefore, the cruise speed could be increased up to a Mach number of 0.65 for this propulsion system. Otherwise, it would be recommended to use turbofan engines as propulsion system for this aircraft concept. Similar to the cruise speed, a lower cruise altitude results in less fuel weight required to fulfill the mission which is beneficial in terms of operational costs. A lower cruise altitude is also beneficial in terms of sustainability. The impact of water vapor and oxides of nitrogen on the global warming effect reduces with decreas- ing cruise altitude [7]. As a conclusion, the competitiveness in terms of transport efficiency and sustainability with respect to maritime transport increases for a higher cruise speed and lower cruise altitude compared to the initial TLRs stated in Section 1.4. 62 5. Sensitivity Analysis

PFEE

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Fuel weight

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800meters

OEW

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Wing span -15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

L/D

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Figure 5.1: Results of sensitivity study of multifuselage concept. 5.3. Conclusions 63

PFEE

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Fuel weight

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

OEW

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Wing span

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

L/D

-15 -10 -5 0 5 10 15 M=0.45

M=0.55

90 containers

110 containers

h=7200 meters

h=8800 meters

Figure 5.2: Results of sensitivity study of BWB concept.

6 Conceptual Design Trade-Off

Based on the results of the conceptual design study and the sensitivity analysis for the two unconventional aircraft concepts, a quantitative trade-off is performed to evaluate the performance of both aircraft concepts as a competitive and sustainable alternative to maritime transport. This chapter discusses the trade-off between the two conceptual aircraft concepts in detail.

6.1. Trade-off criteria and weights

The first step of a quantitative trade-off is to identify the criteria for the comparison of the two proposed concepts. As discussed in Chapter 4, the aircraft design and performance of both concepts differs from each other with respect to aircraft dimensions, aircraft weight and transport efficiency. Moreover, the sensitivity of the cruise speed on the transport efficiency is different for both aircraft concepts, as discussed in Chapter 5. Therefore, these criteria are selected for the quantitative trade-off. A weight is assigned to each criterion to express the importance of each criterion. The weights range between one and five, where a weight factor of five corresponds to high importance and a weight factor of one to low importance. The criteria and their corresponding weights for the trade-off are discussed below.

Aircraft dimensions The aircraft dimensions have an influence on the operations of the aircraft. The wing span dimension affects the landing gear position and thereby the requirements for the runway width. How- ever, as discussed in Section 4.3.6, the hydrogen powered ultra large cargo aircraft is not able to operate on existing airports because of the excessive aircraft dimensions and MTOW for both aircraft concepts. There- fore, new airports with specialized runways and apron areas have to be build for this aircraft concept. As a result, the difference between the aircraft dimensions of both aircraft concepts will not have a significant influence on the competitiveness and sustainability with respect to maritime transport. Therefore, a weight of one is assigned to this criterion where the aircraft dimensions are quantified in terms of the fuselage length and the wing span dimension.

Aircraft weight The second criterion is the weight of the aircraft in terms of OEW. The weight of the aircraft in terms of OEW is a rough indication of the costs to produce the aircraft. However, the development costs of the aircraft are also inﬂuenced by the complexity of the aircraft shape, which is discussed in Section 6.3. Furthermore, the weight of the aircraft also inﬂuences the landing fees on the airport. As a result, the aircraft weight in terms of OEW has an effect on both the development costs as well as the operational costs. Therefore, a weight of three is assigned to this criterion.

Transport efficiency The third criterion is the transport efficiency of the aircraft in terms of PFEE, which evaluates the amount of energy required to transport a certain amount of containers over a given distance. A higher transport efficiency in terms of PFEE means that less energy and thereby less fuel weight is required to

65 66 6. Conceptual Design Trade-Off transport the containers over a given distance. Therefore, the transport efficiency has a large impact on the operational costs. During the operational and economical study, it was found that the fuel weight required to fulfill the mission and the fuel price of hydrogen have the largest impact on the freight rate of the containers. Because the performance target for this aircraft concept is a freight rate smaller than 250% of the freight rate of current cargo ships, this criterion is the most important criterion to compare both aircraft concepts as a competitive alternative to maritime transport. Furthermore, a higher transport efficiency results in less environmental pollution because less fuel is consumed and thereby less greenhouse gases are emitted. As a result, this criterion gives an indication of the difference in performance between the two concepts in terms of sustainability and competitiveness with respect to maritime transport. Therefore, a weight of five is assigned to this criterion.

Sensitivity to TLRs Based on the results of the sensitivity analysis, it was found that the cruise speed and the cruise altitude have a large inﬂuence on the transport efﬁciency in terms of PFEE for both aircraft concepts. The sensitivity of the cruise altitude on the PFEE is almost the same for both concepts, whereas the sensitivity of the cruise speed on the PFEE is different for both aircraft concepts. This difference is also taken into account for the quantitative trade-off by means of this criterion. Because the sensitivity is only evaluated based on two perturbations of the cruise speed, the accuracy of this sensitivity analysis is relatively low. Therefore, a weight factor of one is assigned to this criterion.

The two aircraft concepts also differs from each other in terms of stability and controllability. However, because both aircraft concepts meet the requirements with respect to stability and controllability and the stability does not have any inﬂuence on the competitiveness with respect to maritime transport, this criterion is excluded from the trade-off.

6.2. Quantitative trade-off table

By making use of the results presented in Chapter 4 and Chapter 5, the performance of both aircraft concepts on the deﬁned trade-off criteria is quantiﬁed. To evaluate the difference between the two concepts on the trade-off criteria, the relative difference of the multifuselage concept with respect to the BWB concept is taken. As a result, the score of the BWB concept is normalized to one, where the scores of the multifuselage concept vary between 0 and 2. A score above one for a criterion corresponds to an improvement in performance compared to the BWB concept, whereas a score lower than one corresponds to a deterioration in performance compared to the BWB concept.

For example, a lower OEW is preferred with respect to the aircraft weight. The relative difference of the multifuselage concept with respect to the BWB concept is 8% higher in terms of OEW. Therefore, a score of 0.92 is given to the multifuselage concept for this criterion. Regarding the criterion of the aircraft dimensions, the relative difference between the wing span and the aircraft length are multiplied to determine the score for this criterion.

Based on the data provided in Table 6.1, the score of the multifuselage concept for each criterion is computed. The scores of both aircraft concepts for the different criteria are indicated as dotted number in Table 6.1. The ﬁnal score of the aircraft concept is determined by Equation 6.1, where wi refers to the weight of each criterion.

wi scorei score · (6.1) total = P wi P 6.3. Discussion

As can be seen in Table 6.1, the overall score of the BWB concept is only slightly better compared to the overall score of the multifuselage concept. The BWB performs signiﬁcantly better in terms of aircraft dimensions because of the smaller aircraft length. Furthermore, the BWB concept is superior over the multifuselage concept in terms of aircraft weight and transport efﬁciency designed for the initial TLRs stated in Section 1.4. 6.3. Discussion 67

Table 6.1: Trade-off table to quantitatively compare the two aircraft concepts.

Criterion Weight BWB Multifuselage Aircraft dimensions 1 1.0 0.28 Wing span [m] 218 206 Aircraft length [m] 86.5 150 Aircraft weight 3 1.0 0.92 OEW [kg] 1185164 1277505 Transport efﬁciency 5 1.0 0.93 PFEE [kg km MJ 1] 152 141 · · ° Sensitivity to TLRs 1 1.0 2.14 ¢PFEE wrt +10% Mcr 4.41 9.45 Total 10 1.0 0.98

On the contrary, the improvement in terms of transport efficiency for an increase in cruise speed is significantly higher for the multifuselage concept compared to the BWB concept. Therefore, the multifuselage concept has the potential to become superior over the BWB concept for a higher cruise speed. However, the sensitivity of the cruise speed with respect to the transport efficiency is measured between a Mach number of 0.45 and 0.55. Based on this research, no conclusions can be drawn about the transport efficiency for a cruise Mach number above 0.55 for both aircraft concepts. As a result, it is difficult to select an aircraft concept based on this research. Therefore, it is recommended for further research to investigate the optimal cruise speed and cruise altitude for both aircraft concepts to maximize the transport efficiency and minimize the environmental impact. Another limitation of this trade-off is that the aircraft concepts are only compared to each other based on four criteria. These criteria are selected because quantitative data is available for both concepts. However, other important criteria, such as development costs, payload handling and design risk are not taken into account. The complexity of the shape of the BWB concept might result in higher development costs despite of the lower OEW compared to the multifuselage concept. Therefore, a costs analysis has to be performed to quantify the development costs of both aircraft concepts. Moreover, the payload handling of the multifuselage concept is expected to be better than the BWB concept because the containers can be loaded/unloaded simultaneously in three straight rows, where the containers have to be distributed along the wing span for the BWB concept. In order to quantify the performance of both aircraft concepts on this criterion, the loading/unloading time including refuelling time has to be analyzed. Finally, design risk is also a criterion to keep in mind for the trade-off, although it is very difficult to quantify this criterion for both aircraft concepts. In order to compare both aircraft concepts, a risk map has to be generated. A risk map is a matrix which categorizes the design risks based on the severity of the consequence and the frequency of occurrence. Based on this discussion, the following conclusions can be drawn from the trade-off. Designed for the TLRs specified at the start of this research, the BWB concept is preferred over the multifuselage concept as competitive and sustainable alternative to maritime transport because of the smaller aircraft dimensions, the lower aircraft weight, and the higher transport efficiency. However, the performance in terms of transport efficiency increases with increasing cruise speed for both aircraft concepts, where the sensitivity of the cruise speed on the PFEE is twice as large for the multifuselage concept compared to the BWB concept. As a result, the multifuselage concept has the potential to become a more transport efficient aircraft concept with respect to the BWB concept for a cruise speed above 0.55, but further research is required to analyze this. Moreover, the trade-off is limited to four criteria where it was not possible to quantify the performance of both aircraft concepts in terms of development costs, payload handling and design risk. Because of these limitations and the small difference in overall score between the two aircraft concepts, it is therefore recommended to investigate the performance of both aircraft concepts in more detail. Recommendations for further research are presented in Chapter 8.

Conclusions

A research has been initiated to investigate the feasibility of a hydrogen powered ultra large cargo aircraft as a competitive and sustainable alternative to maritime transport for intercontinental transportation of containerized goods. Due to the ability to transport 100 lightweight standardized containers, signiﬁcant costs savings are expected for this new aircraft concept with respect to current cargo aircraft. The scope of this research is to investigate the design and the performance of this aircraft concept by performing a conceptual design study, where the competitiveness of this aircraft concept in terms of freight rate and environmental impact is investigated by an operational and economical study which is performed in parallel to this research. The following conclusions can be drawn from this research.

7.1. Potential aircraft concepts

A literature study has been performed to identify potential aircraft configurations for the application of a hydrogen powered ultra large cargo aircraft. The main outcome of this literature study is that the multifuselage configuration and the BWB configuration are potential aircraft configurations for this application because of the efficient storage of the large payload requirement of 100 lightweight TEU containers and the large pressurized fuel tanks required to store hydrogen. By distributing the containers along the wing span, these two unconventional aircraft configurations require a lower structural weight compared to a conventional aircraft configuration due to the bending moment relief of the payload on the wing. The multifuselage concept has an operational benefit with respect to the BWB concept by loading and unloading the containers simultaneously, whilst the BWB concept has a potential higher aerodynamic efficiency compared to the multifuselage concept because the BWB concept only consists of lifting surfaces. Regarding the propulsion system, it was concluded that turboprop engines with contra-rotating propellers are most suitable for this aircraft concept because of its high propulsive efficiency and high power output.

7.2. Conceptual design framework

In order to make a quantitative trade-off between the two concepts, a conceptual design framework has to be used which is able to investigate the design and performance of both concepts. In its current state of development, it was concluded that the Initiator is not suitable as conceptual design framework for this research, because the current Initiator is not able to support the synthesis and analysis of these unconventional aircraft concepts. Furthermore, the current Initiator is only limited to passenger aircraft powered by kerosene where alternative fuels such as hydrogen are not taken into account. Therefore, a conceptual design framework has been developed which is able to design and analyze both aircraft concepts. A MDO approach has been applied to integrate multiple disciplines for the design and analysis of both unconventional aircraft concepts, as well as to take into account the mutual interactions between the different disciplines already in the conceptual design phase. Because of the signiﬁcant design features of these unconventional aircraft concepts compared to conventional aircraft conﬁgurations, the simple and fast statistical methods which are based on empirical data of existing aircraft are unreliable. Therefore, quasi-analytical and semi-empirical methods

69 70 7. Conclusions are used for the design and analysis of both aircraft concepts. A drawback of these methods is the increased computational time compared to empirical methods. Therefore, only a limited number of design variables are selected for the MDO of both aircraft concepts. By making use of this conceptual design framework, both aircraft concepts are designed and analyzed successfully by low ﬁdelity methods. The level of design accuracy of the conceptual design framework is relatively low compared to the Initiator, although sufﬁcient to compare the two proposed aircraft concepts inside a relatively large design space.

7.3. Aircraft design and performance

Designed for the same TLRs, the main differences in design and performance between the two proposed concepts are summarized in Table 7.1. It can be seen that both aircraft concepts have a wing span larger than 200 meters. The multifuselage concept consists of two large fuselages with a length of 150 meters, where the length of the BWB concept is 86.5 meters. An illustration of the geometric differences between the two concepts and the Antonov An-225 is shown in Figure 7.1. As can be seen in this ﬁgure, the dimensions of both aircraft concepts are signiﬁcantly larger compared to the dimensions of the Antonov An-225. Due to these extreme large aircraft dimensions and excessive payload weight requirement, this aircraft concept is not able to operate on existing airports. Therefore, large investment costs are required to build new airports and hydrogen facilities for the aircraft operations of this new concept.

Figure 7.1: Illustration of the geometric difference between the multifuselage, BWB concept and the Antonov An-225.

In order to cover the main route of the world wide transportation of containerized goods, ﬁve airports have to be build at the following locations: Shanghai region (China), Rotterdam (the Netherlands), Pennsylvania (USA), Novosibirsk (Russia), and Anchorage (USA). The latter two airports are only used for refuelling where the ﬁrst three airport locations are used as full service airports (loading and unloading containers). To minimize the infrastructure and environmental impact, hydrogen is produced locally near the airport locations by making use of onshore wind energy.

In terms of fuel efﬁciency, the BWB concept requires 8.2% less fuel weight compared to the multifuselage concept. This reduction in fuel weight is caused by two reasons. Firstly, the aerodynamic performance in terms of L/D is 7.1% higher for the BWB concept. Secondly, the OEW is 7.8% lower for the BWB concept compared to the multifuselage concept caused by a signiﬁcant weight saving due to the absence of the fuselages.

Compared to current large cargo aircraft, it was found that the BWB concept is 8.6% more transport efficient in terms of PFEE, where the multifuselage concept performs slightly better (1.4%) to current large cargo aircraft. With respect to maritime transport, it was found that the transport efficiency of an average container ship along the Europe-Far East trade line has a transport efficiency of around 47 and 51 times higher compared to the BWB concept and multifuselage concept, respectively. On the other hand, the transportation time is significantly larger for a cargo ship. The sailing time of a container ship from Shanghai to Rotterdam is 20.6 7.4. Sensitivity analysis 71 days, where the transportation time of the hydrogen powered ultra large cargo aircraft is approximately one day. Regarding environmental impact, the hydrogen powered ultra large cargo aircraft is significantly lower compared to existing aircraft and container ships due to the absence of CO2 emission for this aircraft concept. Furthermore, the environmental impact of the water vapor emission is small due to the relatively low cruise altitude. A quantitative assessment of the environmental impact for the introduction of the hydrogen powered ultra large cargo aircraft is performed by the operational and economical study.

Table 7.1: Performance overview of BWB concept and multifuselage concept

Parameter Symbol Unit BWB Multifuselage Difference Wing span b m 218 206 5.5% Length L m 86.5 150 -73% Maximum take-off weight MTOW kg 2,871,834 3,004,081 -4.6% Operational empty weight OEW kg 1,185,164 1,277,505 -7.8% Payload weight PW kg 1,200,000 1,200,000 - Fuel weight FW kg 486,415 526,321 -8.2% Lift-over-drag ratio L/Dcruise - 24.1 22.4 7.1% Payload Fuel Energy Efﬁciency PFEE kg km MJ 1 152 141 7.2% · · °

7.4. Sensitivity analysis Subsequently, a sensitivity study is performed to investigate the influence of TLRs on the aircraft design and performance of the aircraft. It was found that the aircraft dimensions of the multifuselage concept are mainly influenced by the payload weight requirement, whilst the wing span of the BWB concept is mainly influenced by the cruise speed and the cruise altitude. In terms of transport efficiency, it was found that both aircraft concepts become more efficient for a higher cruise speed and a lower cruise altitude. The sensitivity of the cruise altitude on the transport efficiency in terms of PFEE is almost the same for both concepts, whereas an increase in cruise Mach number is around two times more effective for the multifuselage concept compared to the BWB concept. As a result, it is recommended for further research to investigate the optimal cruise speed and cruise altitude for both aircraft concepts to maximize the transport efficiency and minimize the environmental impact.

7.5. Conceptual design trade-off Finally, a quantitative trade-off is performed to evaluate the performance of both aircraft concepts as a competitive and sustainable alternative to maritime transport based on the results of the conceptual design study and the sensitivity study. It was found that the BWB concept has a slightly higher overall score compared to the multifuselage concept based on the performance evaluation of the aircraft dimensions, aircraft weight, transport efﬁciency and the sensitivity to TLRs. However, the quantitative trade-off is limited to four trade-off criteria due to the limited analysis of both aircraft designs. As a result, important trade-off criteria, such as development costs, payload handling, and design risk, are not taken into account for this trade-off. There- fore, recommendations for further research are provided to analyze both aircraft concepts in more detail to complete the conceptual design. As a result, a more grounded trade-off between the two aircraft concepts will be established and an aircraft concept can be selected as baseline design to enter the preliminary design phase.

8 Recommendations

The scope of this research was to investigate the aircraft design and performance of a hydrogen powered ultra large cargo aircraft. A comparison was made between two early conceptual designs of two unconventional aircraft concepts. As can be concluded from the former chapter, further research is required to complete the conceptual design study and to select an aircraft concept as baseline design to enter the preliminary design phase. The recommendations for further research are presented in this chapter.

Reassessment of top level requirements

As can be concluded from the sensitivity analysis, the transport efficiency of the hydrogen powered ultra large cargo aircraft is mainly influenced by the cruise speed and cruise altitude. The transport efficiency increases with a higher cruise Mach number and a lower cruise altitude. A lower cruise altitude also results in a reduction of environmental impact of the greenhouse gas emissions (H2O and NOx ). Therefore, it is recommended to reassess the TLRs and investigate the optimal cruise speed and cruise altitude with respect to environmental impact and competitiveness compared to maritime transport and existing cargo aircraft.

Wing planform

A limitation of this research is that only a limited amount of design variables are selected to reduce the computational time of the conceptual design framework. The conceptual designs presented in Chapter 4 can be improved by introducing more design variables during the numerical optimization. By introducing more design variables, the shape of the wing could be analyzed and optimized in more detail. For example, it is recommended to deﬁne a smooth curved leading edge and trailing edge along the wing span. Furthermore, it is recommended to optimize the airfoil shape and the twist distribution along the wing span in order to minimize the induced drag.

It would also be interesting to include the rear spar position as design variable for the BWB concept. This position has an influence on the container distribution for this aircraft concept. As can be seen in Figure 4.10, there is sufficient volume at the trailing edge section of the inner wing section to place containers. In terms of volume efficiency, it would be beneficial to place the back spar more aft, but on the other hand the structural weight of the wing increases and the CoG position moves aft which can result in difficulties with respect to stability and controllability. Based on this trade-off it would be interesting to determine the optimal location of the rear spar position for this concept.

Next to introducing more design variables into the MDO framework, the aircraft design and performance could be improved by investigating the aerodynamic beneﬁt of winglets for this aircraft concept. The main beneﬁt of these vertical surfaces at the wing tip is the reduction in induced drag. Furthermore, the winglets can be used for lateral directional stability and controllability for the BWB concept.

73 74 8. Recommendations

Fuel tank and container distribution

An initial container and fuel tank distribution has been considered for both concepts. Further research is required to determine the optimal distribution of the containers and the fuel tanks to use the volume in the aircraft as efficient as possible. For example, there is sufficient volume available to place containers on top of each other at the inner wing section for the BWB concept, as can be seen in Figure 4.10. Therefore, multiple floors instead of one floor for the container distribution can be considered to use the volume in the wing more efficient. Furthermore, it is recommended for further research to investigate the optimal orientation of the containers along the wing span.

Based on the distribution of the containers and the placement of the fuel tanks, an additional study is required to design the access doors for loading and unloading the containers as efﬁcient as possible. Subsequently, the loading and unloading time of both aircraft concepts can be determined, which can be used for the trade-off between the two aircraft concepts.

Landing gear design

The landing gear considered for this research is designed based on empirical data of existing aircraft wheels. Because of the excessive payload weight requirement compared to the payload weight of existing aircraft, a large amount of wheels are required to withstand the loads during ground operations. Therefore, it is recommended to investigate a new innovating landing gear design for this aircraft concept. Furthermore, the optimal location of the landing gear system with respect to balance and steering capabilities during ground operations could be analyzed in further research.

Take-off and landing performance

The take-off and landing performance of the aircraft is roughly estimated by the wing power loading diagram. In order to complete the conceptual design study, the take-off and landing performance including the ground performance has to be analyzed in more detail. Therefore, the stability and controllability of both aircraft concepts have to be assessed during these conditions. This has an inﬂuence on the maximum forward position of the CoG, which is not taken into account during this research. Therefore, it is recommended to investigate the take-off and landing performance of the aircraft in more detail.

Fuel weight analysis

The fuel weight is estimated based on the fuel fraction method. The fuel consumption during cruise and loiter condition are estimated by making use of the Brequet equations, where the fuel consumption of the other mission segments are estimated based on statistical data. Because this aircraft concept is powered by hydrogen instead of kerosene, it is recommended to determine the fuel weight of the mission segments analytically instead of making use of statistical data. The Initiator has a mission analysis module where the fuel consumption per mission segment is determined based on a step by step simulation [69]. Therefore, it is recommended for further research to use this mission analysis module instead of the fuel fraction method.

Engine performance analysis

Another limitation of this research is the simplification of the analysis of the design and performance of the propulsion system. Turboprop engines with very large contra-rotating propellers are required as propulsion system for this aircraft concept. Because the dimensions and the power of the propulsion system considered for this research exceed the dimensions and the power output of current turboprop engines, it is recommended to investigate the performance of this propulsion system in more detail by making use of higher fidelity methods, such as XROTOR. As discussed in Chapter 5, this aircraft concept becomes more transport efficient for a higher cruise speed. Therefore, it is very useful to determine the propulsive efficiency as a function of the cruise speed for this propulsion system. Furthermore, it is recommended to investigate the influence of the propeller slipstream on the aerodynamic performance and stability of the aircraft. 75

Costs analysis As discussed in Chapter 6, a cost analysis module has to be implemented in the conceptual design framework to compare the development costs of the two aircraft concepts quantitatively. Zijp [70] developed a cost analysis module for the Initiator to estimate the development costs and the operational costs for both conventional as well unconventional aircraft concepts. Therefore, a recommendation for further research is to investigate the applicability to implement this module into this conceptual design framework. Subsequently, it is recommended to deﬁne the minimization of the development and operational costs as objective for the MDO framework to maximize the competitiveness with respect to current maritime transport and air cargo transport.

Aerodynamic analysis The aerodynamic solver Q3D is used for this research to determine the aerodynamic performance during cruise condition. Because this solver is not able to take into account the three dimensional effects that occur at the inner wing section of the BWB concept, semi-empirical relationships are used to estimate the drag over the inner wing section. It is recommended to verify this drag estimation over the inner wing section by making use of higher ﬁdelity methods, such as computational ﬂuid dynamics (CFD) methods.

Weight analysis The quasi-analytical method EMWET is used to determine the required wing box weight for this aircraft concept, where the secondary weight of the wing is estimated based on semi-empirical relationships. Because the wing weight is the largest contributor to the OEW of the aircraft, it is recommended to verify this weight estimation by making use of a higher ﬁdelity method, such as a ﬁnite element method (FEM). In order to minimize the weight of the wing, it would also be interesting to investigate the possibility to design a composite wing instead of an aluminum wing as considered for this research.

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This appendix provides a description how the model is executed and where the main files are stored in the folder structure. An overview of the folder structure is given in Figure A.1. The MDO framework is executed by running the file RUN_BWB.m for the BWB concept or RUN_multifuselage.m for the multifuselage concept in the main folder. This MATLAB file initiates the SUMO toolbox by go.m, which is used to perform the DOE and the construction of the surrogate models as discussed in Section 3.2. The set-up of the SUMO toolbox is defined in a xml file, stored in the folder ’Config’. The xml file specifies the amount of output variables (and thereby the amount of surrogate models), the sampling technique, the amount of samples and the surrogate modelling technique. Two xml files are defined for each configuration, named as 1.xml and 2.xml. The first xml file is used to define the DOE, whereas the second xml file is used to update the surrogate models based on the data of the DOE and the data of the numerical optimization. The SUMO toolbox starts to generate the predefined amount of unique design vectors based on the Latin Hypercube Sampling method, where the design variables including the lower and upper bound are specified in .xml in the folder /examples/Design/. Subsequently, a conceptual design is generated and analyzed for each design vector by the MATLAB file ConceptualDesign.m stored in the same folder. The design and analysis methods and tools explained in Section 3.3 are executed by making use of multiple .m files, where the TLRs are specified in InputParameters.m. External tools which are used during the MDA are stored in the ’Storage’ folder. The output of ConceptualDesign.m are the fuel weight, L/D, MTOW, wing weight, induced drag and the constraint evaluation. After all design vectors are analyzed, a surrogate model is built for each output variable. The surrogate models are stored in the folder /examples/Design//output and used during the numerical optimization. The data generated during the DOE is stored into the text file data.txt. This text file contains all the information of each design vector: the values of the design variables, the output of ConceptualDesign.m, and the constraint evaluation. The last step of the MDO framework is to generate a report containing all the information of the aircraft concept. This report, including all figures illustrating the performance of the aircraft, are stored in the folder /examples/Design//Figures.

81 82 A. User manual

Figure A.1: Overview of the folder structure for the BWB concept. B Database of large cargo aircraft

The following database is developed to make an estimation of the OEW as a fraction of the MTOW for the Class I weight estimation. The data of existing aircraft is provided by Janes [28], where the information about the PELICAN is given in [47].

Table B.1: Database of OEW and MTOW for different aircraft (in kilograms).

Aircraft MTOW OEW Airbus A400M 137500 76500 Antonov AN22 250002 120000 Tupolev TU-114 175000 93000 Douglas C133B 129727 54596 Shorts Belfast 104326 58967 Lockheed C-130J-30 Super Hercules 79380 39765 Airbus A380 Cargo 590000 250560 Boeing 747-8F 447695 197130 Antonov An-124 402000 181000 Boeing C-17 Globemaster III 265320 125415 Airbus A300 600F 170500 81525 Kawasaki C-2 141000 60800 Pelican ULTRA 2721554 979760

C Veriﬁcation of wing power loading diagram

A wing power loading diagram is constructed for the Lockheed Super Hercules C130J, a dedicated cargo aircraft powered by turboprop engines, to verify the wing power loading diagram module of the MDA. The specifications of this aircraft are provided by Jane’s all the world’s aircraft [28], where estimations of the maximum lift coefficient during take-off and landing, climb rate and Oswald efficiency factor are made based on the handbook of Raymer [36]. The specifications are listed in Table C.1. Based on the reference data of Jane’s all the world’s aircraft [28], the same power loading was found for a wing loading of 4500 N/m2. The wing power loading diagram is illustrated in Figure C.1, where the design space is indicated in light bue. As can be seen in this figure, the landing distance imposes a limit on the maximum wing loading of the aircraft. The maximum power loading is constrained by the other performance constraints for a given wing loading. In case of the Lockheed Super Hercules C130J, the power loading is limited by the take-off distance requirement.

Table C.1: Speciﬁcations of Lockheed Super Hercules C130J based on Jane’s all the world’s aircraft [28] and the handbook of Raymer [36].

Parameter Symbol Value Unit Aspect ratio AR 10 - Maximum lift coefficient (landing) CLmax,landing 2.8 - Maximum lift coefficient (take-off) CLmax,TO 2.2 - Climb rate crate 5 m/s Oswald efficiency factor e 0.8 - Cruise altitude hcruise 8535 m Cruise speed Vcruise 177 m/s Landing speed Vlanding 60 m/s Landing distance slanding 778 m Take-off distance sTO 930 m Wing loading W /S 4500 N/m2 Power loading W /P 0.0534 -

85 86 C. Veriﬁcation of wing power loading diagram

0.5 Landing distance s=778 m

0.45 Take-o, sTO=930 m Cruise speed V= 177 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure C.1: Wing power loading diagram for the Lockheed Super Hercules C130J. D Final report of multifuselage concept

(a) Top view (b) Front view

Figure D.1: Aircraft geometry (all dimensions in meters)

General Characteristics

Aircraft HULC generated by version 8.4.0.150421 (R2014b). The aircraft is a multifuselage concept with a wing aspect ratio of 6.8. The aircraft is designed to transport 100 containers with a total payload of 1200000kg over 7400km.

87 88 D. Final report of multifuselage concept

Speciﬁcation

Table D.1: Speciﬁcation Mission Requirements

Containers 100 - Payload mass 1200000 kg Cruise Mach 0.5 - Altitude 8000 m Range 7400 km Takeoff distance 3300 m Landing distance 3300 m

Operational Performance

0.5 Landing distance s=3300 m

0.45 Take-o, sTO=3300 m Cruise speed V= 154.031 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure D.2: Wing Power Loading Diagram

Result: Wing loading at MTOM: 4566 N/m2 Power loading at MTOW: 0.062 -

Table D.2: Performance results

L/Dcruise 22.4 - Cruise altitude 8000 m Range 7400 km 89

Weight estimation

Landing gear (2%)

Wing (24%)

Payload (40%)

VT (< 1%) HT (1%)

Fuselage (7%)

Fuel tank (3%)

Engine (2%) Fixed Equipment (2%)

Fuel (18%)

Figure D.3: Mass Distribution

Table D.3: Mass summary

Maximum take-off mass 3004081 kg Operational empty mass 1277505 kg Payload mass 1200000 kg Fuel mass 526321 kg

Table D.4: Masses of aircraft components

Wing weight 731623 kg Horizontal tail weight 40139 kg Vertical tail weight 30031 kg Fuselage weight 221075 kg Landing gear weight 51332 kg Fuel tank weight 81210 kg Total engine weight 74902 kg Fuel system weight 9916 kg Flight control system weight 1328 kg Auxilary Power Unit (APU) weight 8204 kg Instruments weight 603 kg Hydraulic and/or pneumatic system weight 635 kg Electrical system weight 830 kg Avionics weight 971 kg Air-conditioning weight 196 kg Anti- and de-icing systems weight 5306 kg Handling gear weight 796 kg Cargo handling equipment weight 18407 kg 90 D. Final report of multifuselage concept

Center of Gravity

Table D.5: Location of neutral point and CG

CG at MTOW 14.2 % MAC CG at Cruise 17.3 % MAC CG at ZFW 20.7 % MAC CG at OEW 56.3 % MAC Neutral point 30.9 % MAC Static Margin 10.2 %

Aerodynamics

Table D.6: Aerodynamic properties at cruise

Angle of attack 7.4 ° CL,cruise 0.67 - CD,cruise 0.0297 - L/Dcruise 22.4 -

CDi 0.0212 -

CD0 (Total) 0.00843 -

CD0 (Wing) 0.00412 -

CD0 (Fuselage) 0.00314 -

CD0 (Horizontal Tail) 0.000561 -

CD0 (Vertical Tail) 0.000595 -

CD0 (Nacelle) 2.05e-05 -

CLmax,clean 1.58 -

CLmax,take-off 2.02 -

CLmax,landing 2.15 - 91

0.8 30

0.75 25 0.7

0.65 20 l l

C 0.6 c*C 0.55 15

0.5 10 0.45

0.4 5 0 20 40 60 80 100 120 0 20 40 60 80 100 120 Spanwise position Spanwise position (a) C b/2 (b) c C b/2 l ° · l °

Figure D.4: Lift distribution along the wing span

1.5 l

C 1

0.5

0 0 5 10 15 20 Angle of Attack

Figure D.5: C Æ curve l °

Propulsion

Table D.7: Characteristics of propulsion system.

Number of engines 6 - SFCcruise 2.16e-08 kg/J Engine Diameter 0.97 m Engine Length 8 m Propeller radius 7.5 m 92 D. Final report of multifuselage concept

Aircraft Geometry

Table D.8: Main Wing

AR 6.8 - Span 206.1 m Table D.9: Fuselage Planform Area 6280 m2 Length 150 m MAC 33.5 m Diameter 13.6 m Root chord 42.3 m Slenderness cabin 11 - Root t/c 0.15 - Slenderness nose 1.5 - Tip chord 12.7 m Slenderness tail cone 2 - Tip t/c 0.083 - Sweep LE 0.047 ° Taper ratio 0.3 -

Table D.10: Horizontal stabilizer Table D.11: Vertical Stabilizer

Span 50.1 m Span 20.3 m Planform Area 628 m2 Planform Area 316 m2 MAC 13 m MAC 17.1 m Root chord 16.7 m Root chord 24 m Root t/c 0.083 - Root t/c 0.083 - Tip chord 8.35 m Tip chord 7.2 m Tip t/c 0.083 - Tip t/c 0.083 - Sweep 0.25c 4.76 ° Sweep 0.25c 22.5 ° Taper ratio 0.5 - Taper ratio 0.3 -

(a) 3D (b) Side

Figure D.6: Plot of the fuselage including container and fuel tank. 93

Conclusions

Table D.12: Final Overview

W/S 4566 N/M 2 W/P 0.062 - AR 6.8 - Span 206.1 m Length 150 m Planform Area 6280 m2 Maximum take-off mass 3004081 kg Operational empty mass 1277505 kg Payload mass 1200000 kg Fuel mass 526321 kg Wing weight 731623 kg Horizontal tail weight 40139 kg Vertical tail weight 30031 kg Fuselage weight 221075 kg L/Dcruise 22.4 - Angle of attack 7.4 ° CL,cruise 0.67 - CD,cruise 0.0297 -

CDi 0.0212 -

CD0 (Cruise) 0.00843 -

CLmax,clean (Clean) 1.58 -

CLmax,TO (TO) 2.02 -

CLmax,landing (Landing) 2.15 - SM 10 %

CnØ 0-

ClØ -0.035 - ±rudder 14.7 ° 2 W /Smax 5912 N/m PFEE 141 kg km/MJ ·

E Final report of blended-wing-body concept

(a) Top view (b) Front view

Figure E.1: Aircraft geometry (all dimensions in meters)

General Characteristics

Aircraft HULC generated by version 8.4.0.150421 (R2014b). The aircraft is a blended-wing-body concept with a wing aspect ratio of 5.1. The aircraft is designed to transport 100 containers with a total payload of 1200000kg over 7400km.

95 96 E. Final report of blended-wing-body concept

Speciﬁcation

Table E.1: Speciﬁcation Mission Requirements

Containers 100 - Payload mass 1200000 kg Cruise Mach 0.5 - Altitude 8000 m Range 7400 km Takeoff distance 3300 m Landing distance 3300 m

Operational Performance

0.5 Landing distance s=3300 m

0.45 Take-o, sTO=3300 m Cruise speed V= 154.031 m/s

0.4 Takeo,, First segment, FAR 25.111 (OEI)

Takeo,, Second segment, FAR 25.121 (OEI)

0.35 Takeo,, Third segment, FAR 25.121 (OEI)

Landing, First segment, FAR 25.119 (AEO) 0.3 Landing, First segment, FAR 25.121 (OEI)

Climb Rate c = 5 m/s 0.25 Design Space

Design Point 0.2 Power loading (WP) [-]

0.15

0.1

0.05

0 0 1000 2000 3000 4000 5000 6000 7000 8000 Wing loading (WS) [N/m2]

Figure E.2: Wing Power Loading Diagram

Result: Wing loading at MTOM: 2928 N/m2 Power loading at MTOW: 0.066 -

Table E.2: Performance results

L/Dcruise 24.1 - Cruise altitude 8000 m Range 7400 km 97

Weight estimation

Landing gear (5%)

Wing (28%) Payload (42%)

VT (< 1%) Fuel tank (2%) Engine (2%) Fixed Equipment (2%)

Fuel (17%)

Figure E.3: Mass Distribution

Table E.3: Mass summary

Maximum take-off mass 2871834 kg Operational empty mass 1185164 kg Payload mass 1200000 kg Fuel mass 486415 kg

Table E.4: Masses of aircraft components

Wing weight 816729 kg Horizontal tail weight 0 kg Vertical tail weight 23564 kg Fuselage weight 0 kg Landing gear weight 156722 kg Fuel tank weight 65334 kg Total engine weight 69083 kg Fuel system weight 9643 kg Flight control system weight 2093 kg Auxilary Power Unit (APU) weight 7789 kg Instrumentation weight 557 kg Hydraulic and/or pneumatic system weight 548 kg Electrical system weight 686 kg Avionics weight 971 kg Air-conditioning weight 196 kg Anti- and de-icing systems weight 5036 kg Handling gear weight 755 kg Cargo handling equipment weight 25453 kg 98 E. Final report of blended-wing-body concept

Center of Gravity

Table E.5: Location of neutral point and CG as a percentage of MAC

CG at MTOW 21.5 % MAC CG at Cruise 24.4 % MAC CG at ZFW 27.5 % MAC CG at OEW 31 % MAC Neutral point 32.5 % MAC Static Margin 5 % MAC

Aerodynamics

Table E.6: Aerodynamic properties at cruise

Angle of attack 7 ° CL,cruise 0.43 - CD,cruise 0.0178 - L/Dcruise 24.1 -

CDi 0.01191 -

CD0 (Cruise) 0.00588 -

CD0 (Inner Wing) 0.00393 -

CD0 (Outer Wing) 0.00141 -

CD0 (Vertical Tail) 0.000505 -

CD0 (Nacelle) 2.66e-05 -

CLmax,clean 1.6 -

CLmax,take-off 1.6 -

CLmax,landing 1.6 -

1 25

0.9

0.8 20

0.7 l l

C 0.6 15 c*C 0.5

0.4 10

0.3

0.2 5 0 20 40 60 80 100 120 0 20 40 60 80 100 120 Spanwise position Spanwise position (a) C b/2 (b) c C b/2 l ° · l °

Figure E.4: Lift distribution along the wing span 99

1.8

1.6

1.4

1.2

l 1 C

0.8

0.6

0.4

0.2

0 0 2 4 6 8 10 12 14 16 18 20 Angle of Attack

Figure E.5: C Æ curve l °

Propulsion

Table E.7: Characteristics of propulsion system.

Number of engines 6 - SFCcruise 2.16e-08 kg/J Engine diameter 0.95 m Engine length 7.7 m Propeller radius 7.1 m

Aircraft Geometry

Table E.8: Main Wing Table E.9: Vertical Stabilizer AR 5.1 - Span 217.8 m Span 22.5 m Planform Area 9256 m2 Planform Area 389.3 m2 MAC 56.2 m MAC 19 m Root chord 86.5 m Root chord 26.6 m Root t/c 0.16 - Root t/c 0.1 - Tip chord 10.4 m Tip chord 7.99 m Tip t/c 0.12 - Tip t/c 0.1 - Sweep LE 55,10 ° Sweep 0.25c 22.5 ° Taper ratio 0.40,0.30 - Taper ratio 0.3 - b fraction 0.377 - 100 E. Final report of blended-wing-body concept

-10

-20

-30

-40

Y-axis [m] -50

-60

-70

-80

-90 0 20 40 60 80 100 120 X-axis [m]

Figure E.6: Overview of the placement of containers 101

30 30 30 30 20 20 20 20 10 10 10 10

t [m] 0 t [m] 0 t [m] 0 t [m] 0

-10 -10 -10 -10 -20 -20 -20 -20 -30 0 20 40 60 80 0 20 40 60 80 0 20 40 60 0 20 40 60 c [m] c [m] c [m] c [m] (a) Fuel tank (b) Row 1 (c) Row 2 (d) Row 3

30 20 20 20 20

10 10 10 10

0 0 t [m] 0 t [m] 0 t [m] t [m]

-10 -10 -10 -10

-20 -20 -20 -20 0 20 40 60 0 20 40 60 0 20 40 60 0 20 40 60 c [m] c [m] c [m] c [m] (e) Row 4 (f) Row 5 (g) Row 6 (h) Row 7

15 15 10 10 10 10 5 5 5 5

0 0 0 0 t [m] t [m] t [m] t [m]

-5 -5 -5 -5 -10 -10 -10 -10 -15 0 10 20 30 40 0 10 20 30 0 10 20 30 0 10 20 30 c [m] c [m] c [m] c [m] (m) Row 12 (n) Row 13 (o) Row 14 (p) Row 15

10 10 10 10

5 5 5 5

0 0 0 0 t [m] t [m] t [m] t [m]

-5 -5 -5 -5

-10 -10 -10 -10 0 10 20 30 0 10 20 30 0 10 20 30 0 10 20 c [m] c [m] c [m] c [m] (q) Row 16 (r) Row 17 (s) Row 18 (t) Row 19

10 10

5 5

0 0 t [m] t [m]

-5 -5

-10 0 10 20 0 10 20 c [m] c [m] (u) Row 20 (v) Row 21

Figure E.7: Cut-outs of the aircraft to show the placement of the fuel tank and the containers, starting from the centerline towards the wing tip. 102 E. Final report of blended-wing-body concept

Conclusions

Table E.10: Final Overview

W/S 2928 N/M 2 W/P 0.066 - AR 5.1 - Span 217.8 m Length 86.5 m Planform Area 9256 m2 Maximum take-off mass 2871834 kg Payload mass 1200000 kg Fuel mass 486415 kg Operational empty mass 1185164 kg Wing weight 816729 kg Horizontal tail weight 0 kg Vertical tail weight 23564 kg Fuselage weight 0 kg L/Dcruise 24.1 - Angle of attack 7 ° CL,cruise 0.428 - CD,cruise 0.0178 -

CDi 0.0119 -

CD0 (Cruise) 0.00588 -

CLmax,clean (Clean) 1.6 -

CLmax,TO (TO) 1.6 -

CLmax,landing (Landing) 1.6 - SM 5 %

CnØ 0.028 -

ClØ -0.054 - ±rudder 25 ° 2 W /Smax 4404 N/m PFEE 152 kg km/MJ ·