design study of a fuel-cell powered aircraft 06

CRANFIELD UNIVERSITY

EDOUARD MENARD

DESIGN STUDY OF A FUEL-CELL POWERED AIRCRAFT

Msc AEROSPACE VEHICLE DESIGN

COLLEGE OF AERONAUTICS

Msc THESIS

http://www.cranfield.ac.uk/index.htm

COLLEGE OF AERONAUTICS

DEPARTMENT OF AEROSPACE TECHNOLOGY

Msc Thesis

Academic Year 2005-2006

EDOUARD MENARD

DESIGN STUDY OF A FUEL CELL POWERED AIRCRAFT

Supervisor: Pr John Fielding

September 2006

This Thesis is submitted in partial fulfilment for the

Degree of Master

of Science

©Cranfield University 2006. All rights reserved. No part of this publication may be reproduced without the permission of the copyright owner.

Fuel-cell powered aircraft

Warning This thesis has been assessed as of satisfactory standard for the award of a Master of

Science degree in Aerospace Vehicle Design. This thesis covers the part of the

assessment concerned with the Individual Research Project. Readers must be aware

that the work contained is not necessarily 100 % correct, and caution should be

exercised if the thesis or the data it contains is being used for future work. In doubt,

please refer to the supervisor named in the thesis, or the Department of Aerospace

Technology.

Edouard Ménard - i


Acknowledgements

I would like to thank my supervisor, Pr Fielding for his support and experienced

designer advices, always given in an open-minded manner.

Thank you to Mr David Daggett, Orti and Nieves from Boeing who gave me precious

advices on the importance of compressor and cooling systems.

I also wish to thank all my fellow classmates, housemates and friends who made the

long nights in the system labs more appreciable, and to all the people I met this year

for making this year so rewarding.

Finally I am very grateful to my parents for supporting me morally and financially for

this year at Cranfield University, so that I am now able to begin my professional life

in the very best conditions.

Edouard Ménard - ii


Abstract

Although fuel cell is a technology known since the nineteenth century, and developed for space missions in the 1960’s, it is only in the past decade that energetic dependence to limited fossil fuels and global warming effects accelerated the development of the technology. In the past years sufficient levels of power for automotive applications have been reached, and research is now carried out to introduce fuel cell first as auxiliary power units in civil airliners and in the very long term as primary power plant. This work focuses on smaller applications such as general aviation for which the power requirements are close to be achieved. The initial purpose of this project is to demonstrate the feasibility of a fuel cell powered aircraft, and based on current technology design a light reconnaissance airplane. This design allowed deriving a number of requirements for the use of fuel cell system as primary power plant, and discusses the impact of fuel cell and hydrogen storage on aircraft design. A power plant adapted to the mission requirement was designed, including the air feeding device and cooling device which are as important as the basic fuel cell stack, inverter and electric motor. Liquid hydrogen storage was investigated, and although the numerous safety and technological questions, a viable concept has been developed very close from the current storage system in space launchers. The conceptual design of a fuel cell powered aircraft has been achieved to fulfil reconnaissance missions. Compared to existing competitors the performances are generally equivalent except the range which is slightly reduced. However fuel cells are getting closer and closer from the requirements for light aviation.

Edouard Ménard - iii -


Notations

A Aspect Ratio - a 2D Lift curve slope - a0 3D lift curve slope - b Span m c Chord m c' hydrogen specific consumption Cd 2D drag coefficient - CD 3D drag coefficient - Cl 2D lift coefficient - CL 3D lift coefficient - Cpair Specific Heat of air KJ/Kg.K CpH2 Specific Heat of hydrogen 14000 KJ/Kg.K CpH2O Specific Heat of Water 4180 KJ/Kg.K D Drag N d propeller diameter m η Fuel cell efficiency 0.5 - ηp propeller efficiency - k thermal conductivity W/mK L Lift N μ air viscosity poiseuille Np Number of cells per stack - Q quantity of heat exchanged J

q dynamic pressure Kg.m-1.s-2

qair quantity of air moles qH2 quantity of hydrogen moles qO2 quantity of oxygen moles ρ air density kg/m3

Re Reynolds number - Text External temperature of the aircraft ºC TinFC Inlet temperature of the fuel cell stack ºC Tint Temperature inside ºC TstH2 Temperature of hydrogen storage ºC V Aircraft speed m/s W Aircraft weight (TO) kg WE Zero fuel weight kg

Edouard Ménard - iv -


Table of Contents

Introduction....................................................................................................................1 Literature review............................................................................................................2

1. Context, Why using fuel cells? ..........................................................................2 1.1. Oil prices trends .........................................................................................2 1.2. Pollution.....................................................................................................3 1.3. Noise reduction ..........................................................................................4

2. Fuel cell state of the art......................................................................................4 2.1. Stacks .........................................................................................................4 2.2. Proton Exchange Membrane (PEM)..........................................................5 2.3. Solid Oxide Fuel cell (SOFC)....................................................................7 2.4. Direct Methanol Fuel Cell (DMFC) ..........................................................8

3. Fuels...................................................................................................................9 3.1. Production of hydrogen..............................................................................9 3.2. Storage of Hydrogen................................................................................11 3.3. Towards a hydrogen economy.................................................................12 3.4. Methanol ..................................................................................................12

4. Safety ...............................................................................................................13 4.1. Fuel safety................................................................................................13 4.2. Material compatibility..............................................................................13

5. Fuel cell system................................................................................................14 5.1. Electric Motors.........................................................................................14 5.2. Batteries ...................................................................................................16 5.3. Super capacitors .......................................................................................16

6. Existing projects...............................................................................................17 6.1. Automotive products................................................................................17 6.2. Aircraft.....................................................................................................18 6.2.2. UAV.....................................................................................................18

7. Conclusions......................................................................................................18 Aircraft type and missions ...........................................................................................20

1. Introduction:.....................................................................................................20 2. Reconnaissance airplane: .................................................................................20

2.1. Performances............................................................................................21 2.2. Fuel cell installation.................................................................................22

3. 4 Seat Family aircraft.......................................................................................24 3.1. Performances............................................................................................24 3.2. Fuel Cell installation ................................................................................25

4. UAV.................................................................................................................27 5. Direct Operating Costs.....................................................................................27

5.1. Fuel prices................................................................................................27 5.2. Acquisition costs......................................................................................28

Edouard Ménard - 1 -


5.3. Maintenance costs....................................................................................29 6. Choice ..........................................................................................................29

Requirements and design options ................................................................................30

1. Light aircraft survey.........................................................................................30 2. Basic Requirements .........................................................................................30 3. Design options .................................................................................................34

3.1. Optica design ...........................................................................................34 3.2. High wing, twin pusher propeller ............................................................35 3.3. Canard configuration ...............................................................................35 3.4. Joint wing concept ...................................................................................36

4. Concept choice.................................................................................................36 Power plant Architecture .............................................................................................38

1. Architecture......................................................................................................38 2. Hydrogen fuel system ......................................................................................40

2.1. Hydrogen specific consumption ..............................................................40 2.2. Hydrogen quantity ...................................................................................40 2.3. Hydrogen tanks ........................................................................................41 2.4. Fuelling refuelling....................................................................................43 2.5. Tank Materials .........................................................................................44

3. Cooling system.................................................................................................46 3.1. Fuel Cell Stack cooling............................................................................46 3.2. Inverter.....................................................................................................47 3.3. Electric Motor ..........................................................................................47 3.4. Water pump..............................................................................................47

4. Oxygen intake ..................................................................................................48 4.1. Fuel cell feeding.......................................................................................48 4.2. Compression devices ...............................................................................49 4.3. Supercharger selection and performances ...............................................49

5. Heat balance.....................................................................................................50 5.1. Intake of the Cell......................................................................................50 5.2. Cooling system.........................................................................................50

6. Conclusions and mass summary ......................................................................51 Initial Design................................................................................................................52

1. Introduction......................................................................................................52 2. The baseline optimisation process ...................................................................52 3. Modified spreadsheet .......................................................................................53

3.1. Specific power of power plant .................................................................53 3.2. Fuel cells ..................................................................................................53 3.3. Fuel specific consumption .......................................................................53 3.4. Start of Climb mass and Landing mass....................................................53 3.5. Wing mass penalties ................................................................................54 3.6. Available hydrogen mass onboard...........................................................54

4. Results..............................................................................................................54 5. Wing Selection.................................................................................................57



5.1. Investigated profiles.................................................................................57 5.2. 3D performances......................................................................................58 5.3. Flaps.........................................................................................................60 5.4. Wing selection and design .......................................................................63

6. Conclusion .......................................................................................................64 Design development.....................................................................................................65

1. General Arrangement.......................................................................................65 2. Wings ...............................................................................................................65 3. Fuselage ...........................................................................................................66 4. Tail Unit ...........................................................................................................67 5. Nacelles............................................................................................................68 6. Power plant ......................................................................................................69 7. Cabin ................................................................................................................69 8. Landing gears and ground clearances ..............................................................71 9. Hydrogen tanks ................................................................................................72

Design analysis ............................................................................................................74

1. Mass and CG management ..............................................................................74 2. Drag estimation................................................................................................75 3. Performances....................................................................................................76 4. Lift Drag ratio ..................................................................................................77 5. Propellers efficiency ........................................................................................77 6. Range ...............................................................................................................78 7. Summary of performances...............................................................................78 8. Comparison to competitors ..............................................................................79

Discussion....................................................................................................................81 Recommendations........................................................................................................83 Conclusion ...................................................................................................................84 References....................................................................................................................85 List of Figures ..............................................................................................................90 List of Tables ...............................................................................................................92 List of Appendices .......................................................................................................93


Fuel-cell powered aircraft Introduction

Introduction

Fuel cell technology has been developed for decades, but in order to solve the problems of pollution and dependence to expensive fossil resources, fuel cell technology has been recently developed into stationary power generation and automotive propulsion. As automotive industry invests very large amounts of money in the technology, fuel cells are becoming competitive with gasoline piston engines and will be the alternative solution in the next few years. Aviation industry will have to face the same issues in a few years time and lots of research has been done in the last years to improve specific power of fuel cells, and apply the technology to aircraft. Space industry has always looked at fuel cell and hydrogen technology and was already developing and using fuel cells in Apollo missions in the 1960’s. Hydrogen storage has also been the focus of serious interest as liquid hydrogen is used as reactant in launchers propulsion systems. It is clear that hydrogen and fuel cell technology has been investigated for a long time, and is on the edge of achieving sufficient power levels for light aviation applications. Major technological challenges are still to be solved, but potential of the technology to replace oil economy is widely accepted. The purpose of this thesis is to provide as comprehensive as possible analysis of the current technology development and challenges. Based on available technology, a light reconnaissance aircraft has been designed to demonstrate the feasibility of a fuel cell powered aircraft, derive requirement for fuel cell systems, and on the other hand analyse the impact of this technology on aircraft design. Resulting performances have also been compared to gasoline competitors and other publications.


Fuel-cell powered aircraft Literature review

Literature review

1. Context, Why using fuel cells?

1.1. Oil prices trends In a context of rising oil prices, due to political tensions in Middle-East and growing demand of emerging countries, the crude oil price is not likely to decrease.

Figure 1: Oil prices trends, 2004, (US Department of Energy 2006)

Figure 2 : Recent evolution of the oil barrel price, (www.wikipedia.org)


http://www.wikipedia.org/


In 2004, a report from the US department of energy, published three trends for the evolution of the crude oil barrel price, however in the past months, the barrel price has risen up to around 75 $, far over the worst predictions. Over the past 24 months the crude oil price has risen from 40 $ up to 75 $ a barrel. Due to emerging countries demand for energy, the price of crude oil will stay high or even increase. In that context, alternative energy to fossil fuels become necessary, and a lot of efforts and money has been invested in research and development of fuel cell systems both for power plant applications and automotive applications.

1.2. Pollution The main advantage of a fuel cell powered aircraft would be to reduce significantly the emissions of greenhouse gases to zero or near zero. It must be kept in mind that air transport is responsible for around 10% of the global warming effect and 3% of carbon emissions. The growing market of air transport will pollute even more, releasing more carbon and particles, and NOx components.

Figure 3 : Aviation Carbon release in atmosphere, (Penner, J.E. 2006)

Fuel cells are much more efficient that combustion engines, to compare fuel cell systems have efficiencies around 50 % and could be improved up to 60 %, whereas combustion engines have efficiencies of 15 %. In a sustainable development approach, higher efficiencies up to a ratio of 3 means burning less fuel, and then reduce the energy necessary to produce the fuel (oil refinery or hydrogen production). However, water steam is also a greenhouse gas and to eliminate the pollution, cruise altitude has to be limited so that rejected water turns back to its liquid state. Otherwise, water has to be kept onboard inducing a landing weight heavier than at take-off, which is not desirable.



1.3. Noise reduction The use of fuel cell and electric motor to power an aircraft can bring significant noise reduction compared to a piston engine. FC are still far from replacing turbofans, but on long term prospects FC could reduce noise due to airliners around airports. Airframe noise cannot be reduced by comfort within the cabin will increase significantly.

2. Fuel cell state of the art Fuel cells have been known since the 19th century, and many different types have been developed. This section focuses only on the most promising types of fuel cell for the aviation industry. The main parameter being specific power, proton exchange membrane fuel cell and solid oxide fuel cell are presented. These two types of cells are fed by hydrogen, leading to storage and production issues which are addressed later in section 3.1. A fuel cell using methanol can also be used, it is less weight efficient, but as methanol is widely available and easy to store, this technology has also been investigated.

2.1. Stacks A single fuel cell only produces a voltage of few Volts (0.6-1.2 V), and many cells have to be combined in series to produce a voltage that can be used. To connect fuel cells, anode and cathode of two different cells have to be connected to sum up the voltage produced (and power produced).

Figure 4 : Single Cell assembly (Larminie, J. 2003)

For compact designs, bipolar plates are used, i.e. plates acting as anode and cathode. A stack is a combination of numerous cells connected through bipolar plates.



Figure 5 : Fuel cell stack (Larminie, J. 2003)

2.2. Proton Exchange Membrane (PEM) Proton exchange membrane fuel cells (PEMFC) have first been developed in the 1960’s by general electric for the Gemini space missions. They have the advantage of working at low temperature (60-80º), allowing a quick start time. As were more expensive that other types of fuel cells, the development has been interrupted during the 1970’s and early 1980’s, which explains that the technology is not mature yet. A wide range of power is available, and PEMFC are preferred by car manufacturers for their compact properties.

2.2.1. Operating principle:

Figure 6 : PEMFC principle (Larminie, J. 2003)

In the PEMFC, hydrogen and oxygen taken from air react to produce water, heat and electricity. At the anode, hydrogen dissociates to produce electrons and protons.



Electrons travels trough the electrical system to produce electricity, whereas protons cross the membrane to recombine with oxygen of the air to give warm water. The membrane is made out of sulphured fluoro-polymers. This component is produce from polyethylene in which the hydrogen atoms have been replaced by a fluorine element to get PTFE. Finally a sulphured chain is combined to produce a membrane. Due to its mechanical strength membranes of 40 µm can be manufactured, and its chemical resistance protects the membrane from acids. Electrodes are made out of carbon mainly, both for the anode and the cathode, and a platinum catalyst is used. The platinum is spread as a powder on the electrodes. The high cost of platinum is responsible for the lack of interest in PEMFC in the 1970’s, but platinum content of the FC has been reduced and cost now less that 10$ per fuel cell stack. Water management is one of the issues limiting the development of the fuel cells. The conductivity of the membrane to H+ (and therefore output power) is proportional to the content of water in the fuel cell, so water must be kept in sufficient quantities. However, tan excess of water blocks the gas emission and compromises the process. The flow of air one side of the membrane dries out the membrane, and creates a gradient of water concentration across the membrane. Drying of the membrane is supported by the temperature in the fuel cell. To prevent all these issues, air has to be humidified, and the level of water in the cell carefully controlled. As the efficiency of the fuel cell is (50%-60%), there is some heat produced which has to be dissipated. For relatively low powers the cooling can be done using air, but for larger powers, above 100 W, the coolant agent is liquid (water), but routing the fluid across the cell becomes more difficult for manufacturing reasons. Pressurised operation of the fuel cell is also a possibility to increase the power output of the cell.

2.2.2. Performance PEMFC are currently one of the more mature and weight efficient technology among fuel cell types. PEM technology is being developed by BALLARD, and has reached power densities of 0.89 kW/kg with the 85kW Mark902, and up to 1.24 kW/kg with smaller stacks of 21 kW. NASA report (Alexander, D.S. 2003), also quotes a specific power of 1.24kW/kg realised by General Motors.



2.3. Solid Oxide Fuel cell (SOFC)

2.3.1. Operating Principle

Figure 7 : SOFC operating principle, (Larminie, J. 2003) Solid oxide fuel cell can be fuelled by either a flow of hydrogen or carbon monoxide (CO). It operates at higher temperatures, between 800 and 1000ºC. Some late researches have been oriented towards a lower operating temperature (around 650ºC). The higher internal temperature is a disadvantage because longer starting times are required, but the fuel cell is more tolerant to fuel impurities than any other types of cells. The higher operating temperature means with an efficiency of around 50%, that heat is produced in large quantities, and can be recycled. Bottoming cycles, which is taking advantage of this heat produced by means of exchangers, is then possible. SOFC are used currently on stationary powers where large powers, up to 10MW, are produced in CHP (Combined Heat Power) stations. The electrolyte is a ceramic material named Zirconia (ZrO2), combined with 8-10% of yttria Y2O3 . This electrolyte is a solid material which gives its name to the cell. Electrodes are made out of high porosity materials to enlarge the reaction area. The anode is made out of Nickel and YSZ (yttria-stabilised zirconia), whereas the cathode is LaSr (strontium and lanthanum) combined with manganese oxide MnO3, which is a semiconductor. There are two main lay-outs for the FC: planar and tubular. In the tubular design, anode, electrolyte and cathode are concentric cylinders, thus maximising the reaction surface.



Planar design has proved higher power densities, but leads to sealing problems.

2.3.2. Performances Because of its high power density, and success in high power units SOFC is among the favourite fuel cell types for the NASA for aerospace applications. However this technology is less mature than PEMFC, for example but is believed to enable power densities of 2kW/kg in the next years.

2.4. Direct Methanol Fuel Cell (DMFC) In order to avoid storage, distribution and production issues of hydrogen, fuel cells using methanol as a fuel have been developed. Methanol is stored very easily, is already produced in large quantities, but is more dangerous than hydrogen.

2.4.1. Operating Principle The global reaction taking place in the cell is:

2223 223 COOHOOHCH +→+

This can be decomposed into:

223 66 COeHOHOHCH ++→+ −+ At the anode

OHeHO 22 36623

→++ −+ At the cathode

6 electrons are produced in the process, which increases the energy production; however this potential advantage is reduced by the fact that the fist reaction takes place in several stages and is then a very slow process. This cell is not 100% clean as it releases carbon dioxide, but it can be argued that hydrogen production uses a lot of energy which pollutes as much as carbon dioxide of DMFC. The global structure of the DMFC is nearly the same as PEMFC; an alkaline FC architecture has been investigated and demonstrated major issues. However, the PEMFC architecture is not without its own problems. Methanol mixes very well with water, and travels through the electrolyte. As described in the PEMFC section, cathode is dried out, and water crosses the electrolyte from the anode to the cathode carrying a significant amount of methanol. The methanol reacts at the cathode, but electrons do not travel through the electrical system and recombines immediately with the oxygen, thus fuel is lost. Moreover this



reaction reduces the voltage output of the cathode, reducing the global efficiency of the cell.

2.4.2. Performances Due to the issues presented above, the fuel cross-over from anode to cathode, and the slow reaction at the anode, DMFC have significantly lower efficiencies than other types of cells, but fuel storage is simpler.

3. Fuels From the three systems presented in the previous section, two fuel can be employed both having their advantages and drawbacks. In order to select the best system, production and storage problems must also be addressed, so that a good balance between performances and environmental issues is made through the whole life cycle of the product. For both methanol and hydrogen, production techniques have been considered in terms of environmental impacts, costs, feasibility. Storage problems have been studied looking at weight and volume of storage systems, as well as energy per volume unit. Then the possible installation of infrastructure supporting both fuels has been looked at.

3.1. Production of hydrogen

3.1.1. Steam reforming Hydrogen can be produced by different methods such as methane steam reforming. The reaction is:

224 3HCOOHCH +→+

Then, 222 HCOOHCO +→+

The heat necessary to start the reaction is generally obtained by burning some of the methane flow:

OHCOOCH 2224 22 +→+ This process is not 100% clean, as it produces carbon dioxide, however it is a technique commonly used. Another drawback is that heat produced by methane combustion is lost at 50% into the output steams.



3.1.2. Partial Oxidation Partial oxidation is a faster process that reforming, oxygen and methane are mixed to a catalyst Ni/SiO2 and the following reaction occurs and produces heat:

224 221 HCOOCH +→+

This process is faster and suitable for mobile applications, such as cars. Hydrogen can be stored as methane, avoiding hydrogen storage issues, but adds complexity and weight to the vehicle. Unfortunately carbon monoxide, which is toxic, is produced for each couple of hydrogen molecules produced (twice as in reforming), and compromises the environmental qualities of the fuel cell system.

3.1.3. Water electrolysis

Water electrolysis is exactly the reverse process of a fuel cell, and in this perspective hydrogen is used as an energy storage device. Electricity is provided to a fuel cell as well as water, and the opposite reactions take place:

+−+− +++→+→ HeOOHHOHOH 2221222 222

222 HeH →+ −+

This process is perfectly clean, except from the energy needed. In a hydrogen economy, energy would be provided by sustainable processes such as solar panels, or wind power. Demonstration of this process has been done through the NASA Helios project. This unmanned demonstrator was solar powered at day recharging fuel cells and powered by fuel cells at night. This is the best way to produce hydrogen, but is not currently applicable, at big scales.

3.1.4. Other processes There are means to produce hydrogen from biomass, such as woody products. However, this process produces also carbon dioxide, and is currently focussed on stationary applications, but reaches efficiencies of 48%. Thermal processes are also possible, but require materials able to withstand extreme temperatures, and are not mature enough for current applications.



3.2. Storage of Hydrogen

3.2.1. Gaseous Hydrogen Hydrogen can be stored upon different forms, liquid or as a gas. Both have advantages and drawbacks, and alternatives hybrid systems are currently designed. Under its gaseous form hydrogen storage requires a lot of volume compared to the high volumetric and specific density of liquid storage. To maximise the volumetric efficiency of gaseous hydrogen systems, high pressure devices have been developed. Pressures up to 700 bars have been investigated tested and are currently commercialised. To reduce the weight composite materials are used and the development of winding manufacturing technique allowed lightweight composite tanks to be produced. Power storage densities up to 0.78 kWh/kg have been demonstrated in a 700 bars 85 kg composite tank carrying 2kg of hydrogen. (Larminie, J. 2003). Due to the cylindrical shape of the tanks, the high pressure storage of hydrogen is a good candidate for external fuel tanks, providing that extra weight is permitted.

3.2.2. Liquid storage Liquid hydrogen has a much greater density than in gaseous state, (71 kg/m3) but which is still lower than AVGAS fuel. However storing hydrogen in a liquid state requires maintaining a cryogenic temperature of -253ºC. Maintaining such a temperature requires either a vacuum insulation between tank and exterior, but if the vacuum compartment fails, the insulation is reduced to zero. The other technique is to use solid insulating materials, foam for example, or glass and aluminium mixture. Solid insulation requires as well a layer of reflecting material to prevent heat transfer by radiations. Liquid hydrogen is the best way to store decent quantities of reactant in a given volume. Moreover, space launchers and rockets uses liquid hydrogen tanks, so that the technology has already been developed somewhat.

3.2.3. Metal hybrids Due to the small atomic radius of hydrogen molecules, hydrogen can be absorbed by other materials. Practically, hydrogen molecules penetrate the structure of the metal, and fit between the atoms of metal. Volumetric densities of reactant stored are promising but generally speaking this technology is still immature and metal hybrids storage systems are heavy.

3.2.4. Carbon nano-tubes



Carbon nano-tubes are also a great potential, for the future of hydrogen storage, as some laboratories reported ratio of stored hydrogen mass to carbon mass up to 67%, however, these results cannot been reproduced with complete confidence and the ratio could be lower. The technology is full of opportunities but immature.

3.3. Towards a hydrogen economy Deciding to design an aircraft powered by hydrogen is not without consequences in the use of the aircraft. How hydrogen will be provided and what will the operating prices be? Production of hydrogen issues have been assessed but distribution issues are still to be discussed. If the production technologies are mastered, a large scale infrastructure is far from being operational. Distribution of hydrogen requires hydrogen stations, and production sites (large facilities or decentralised production). Currently hydrogen fuelling stations are used for buses in big cities, and the number of available stations in 2004 was around 25 in the US, 15 in Germany, 6 in Canada and only one in the UK. (Solomon, D.B. 2004) Most of the governments have roadmaps to introduce hydrogen powered vehicles, but except Iceland no country has defined clear deadlines to convert its infrastructure to hydrogen. Iceland is only 300,000 inhabitants but has decided to convert to hydrogen economy for 2030. Concerning aviation powered by hydrogen, the apparition of hydrogen powered aircraft can be planned after 2020, and will probably appear once the infrastructure has been set up for hydrogen powered cars.

3.4. Methanol As explained earlier, methanol fuel cells have been considered because methanol is a very plausible and available alternative to hydrogen. From Larminie,2003, methanol is currently produced at a rate over 20,000,000 tons per year. It is produced from natural gas by the following reaction:

OHCHCOH 322 →+

OHOHCHCOH 23222 +→+ Reactants H2, CO and CO2 are produced by steam reforming as in the hydrogen production process. Overall, the process does not release toxic substances, except to produce the energy needed for the reaction. Methanol can easily be produced from any hydrocarbon fuel but also from biomass, although currently the natural gas reforming is widely spread. The efficiency of the process is 70%, which explains why methanol is available at low costs.



4. Safety

4.1. Fuel safety Using hydrogen as fuel requires additional safety measures to be taken. However hydrogen is not as dangerous as it seems, and compared to other alternatives fuels is relatively safe. Hydrogen is the smallest element in the periodic table, and has a number of other properties, a very low density, a very low viscosity, an extremely high conductivity. Compared to air, hydrogen leaks 3.3 times faster in a hole of identical diameter. Because of its very low density, hydrogen dissipates quickly upwards and does not tend to concentrate in a small area. Unfortunately, hydrogen fire is invisible, and therefore is much more difficult to fight.

Hydrogen Methane Propane Density kg.m-3 0.084 0.65 2.01

Ignition limits in air % (volume) 4.0-77 4.4-16.5 1.7-10.9 Ignition temperature ºC 560 540 487

Min Ignition energy in air MJ 0.02 0.3 0.26 Max combustion rate in air m.s-1 3.46 0.43 0.47

Detonation limits in air % (volume) 18-59 6.3-14 1.1-1.3

Table 1 : Hydrogen, Methane, Propane properties, Larminie,2003 As shown is table 1, hydrogen is mainly dangerous when mixed with air, but ignites at a significantly higher temperature than other fuels. Even if the ignition energy is lower, greater volumes of gas are required to ignite in air compared to methane or propane, which despite the safety concern they raise, are widely used in domestic applications. The risk of explosion is lower with hydrogen than with other gases, as shown in the last row of table 1.

4.2. Material compatibility Due to the small diameter of hydrogen molecules, hydrogen can easily travel through some materials which have large holes in microscopic structures. This is one reason why carbon composite materials are not hydrogen proof. Some metals are also incompatible with hydrogen storage, as hydrogen may penetrate the atomic structure. For example ferrous alloys are not very resistant to hydrogen, (Bekiraris 2002).



Resistance of alloys to cryogenic temperatures is discussed in the power plant chapter.

5. Fuel cell system

5.1. Electric Motors Electric motors quotes very high efficiencies (90-95%), but once again the issue is the weight. And as power requirements tends to stationary power equipments weight is not generally an issue and often not published. Lighter electric motors are generally alternative current, although continuous current designs exist, but have worst performances. Electric motors are made out of two main parts, a rotor and a stator. The stator is fixed and composed of several couples of poles (reluctances which produce a rotating electromagnetic field). The rotor also produces an internal electromagnetic field, either by permanent magnet, or reluctances. The interaction of the electromagnetic fields generates torque and the rotor rotates. Many types of AC motor exists, either using single phase or multi-phase input current, permanent magnets or brushless architecture. The two main candidates for aerospace applications are induction motor and brushless DC motor.

5.1.1. Induction Motors In the induction motor, the input has to be alternative three phases. The stator is equipped with three couple of windings (poles) which produces each a variable electromagnetic field. The sum of theses three fields is a rotating electromagnetic field. The rotor is made out of a cage (aluminium or copper). When immerged in the rotating field, electric currents are induced in the rotor. The currents reacts the surrounding EM field, thus creating torque. The rotation speed depends directly on the input frequency divided by the number of poles. As the number of poles is fixed, the only way to control rotation speed is to control the input frequency of the motor, by acting directly on the inverter.

5.1.2. Brushless DC Motors In BLDC motors, the input is continuous current; however, the flux through the winding is alternated. The rotor is a permanent magnet. As in the previous type of motors the stator is composed of couples of windings. Instead of having a sinusoidal signal in the windings, the current is continuous, and creates a torque because internal magnetic field of the rotor tends to align itself with the magnetic field created by the stator. (Stage 1) As soon as the rotor is aligned with



the stator, the current in the windings is switched off (stage 2). Then inertia rotates the rotor. (Stage 2) Finally, once that rotor and stator are not aligned, the current is switched in the opposite direction, and torque is applied on the rotor. (Stage 3)

Stage 1 Stage 2 Stage 3

Figure 8 : BLDC motor operating principle

5.1.3. HTS Motors As presented in Masson, P.J. (2005 & 2005), a new type of electric motor has been designed for electric aviation purposes, High Temperature Superconductor motors. The basic though is to replace heavy copper windings by superconductor alloys, so that intense electromagnetic fields are created. A new geometry for the rotor has also been designed. Unfortunately, using superconductors implies to operate at low temperature, around 50º K, and therefore requires a complex cooling system. This paper presents the design of an alternative motor for a Cessna 172 Skyhawk, the results are presented below:

Table 2 : HTS motor for Cessna 172 application (Masson, P.J. 2005)



Figure 9 : Power density for HTS motor. (Masson, P.J. 2005)

The results in table 8 does not include the cooling system, however a weight of 28 kg has been achieved together with a compact design (160*220mm). Such a compact design reduces the nacelle size, thus drag. The cooling system weights around 35 kg, so that a resulting system of 63 kg has been designed versus 160 kg for the original gasoline engine. The cooling system weight could be reduced by taking advantage of the liquid hydrogen available onboard. LH2 is stored at 20 º K, and has a very high heat capacity and need to be heated before entering the fuel cell, so that it could be used as coolant fluid.

5.2. Batteries An additional source of energy considered in many automotive studies is the use of batteries and in the case of this design requires batteries, a basis of commercial batteries will be assumed. The batteries selected are manufactured by the company SAFT, which provides batteries for aircraft from Cessna to Airbus A 340. Two typical models have been looked at: 4076 series providing 36Ah at 24V for a weight of 37 kg (Used on small aircraft) and 505CH series providing 50Ah at 24V for 42 kg (A 340). The use of a battery is limited to providing extra power for take off and emergency situations. However considering the power requirements at take-off, a battery would only provide a few extra kilo Watts, which compared to the extra weight is not acceptable.

5.3. Super capacitors To provide sufficient power for take off, batteries can be used, but super-capacitors are an opportunity for the future.



Super-capacitors, or double layer capacitors, are devices designed to store energy as capacitors are. The system BOOOTSCAP, presented in Vielstich, V. 2003, has reached a capacitance of 1500 F. This performance allows storing a significant amount of energy in a limited volume and weight (5.3Wh/kg). By combining a number of super-capacitors, a system providing high power during short durations can be designed. This paper relates that a stack of 90 super-capacitors, reaching 168 kg demonstrated a supply of 50 kW for 15s with an efficiency of 92%. These devices still do not reach the already insufficient batteries specific power, but may be an interesting potential for future developments.

6. Existing projects

6.1. Automotive products Fuel cell application is much more advanced in automotive products that in aviation industry. Every constructor has research project in fuel cell vehicles, and only a few will be presented. Daimler Chrysler is one of the most interesting company, it has invested serious amount of money over 1 billion $ since 1994, and plans to spent the same amount in the next years to sell 100,000 vehicles by 2010. Daimler Chrysler is currently testing 36 buses in the world and has presented car powered by 85kW fuel cell stacks. However it is difficult to find reliable information on the performances of the fuel cell developed.

Figure 10: Hy-Gen Electric Vehicle, General Motors

General Motors claims to have produced fuel cell producing 70 kW for 1.48kW/kg in 2000 but the latest vehicle the Hy-wire only uses a stack producing 94kW continuous, for a specific power of 0.94kW/kg. The main problem in adapting automotive technologies is that weight is not a primary concern for automobile industry as it is in every aerospace application; however the level of power required to power an aeroplane is slightly the same as to power a bus.



6.2. Aircraft

6.2.1. Boeing demonstrator Boeing research centre in Madrid is currently developing a demonstrator based on a 2 seats light aeroplane Diamond HK-36. The fuel cell selected delivers around 15 kW for cruise and is assisted by batteries to feed a 50kW peak electric motor. The fuel is high pressure composite tanks.

Figure 11 : Boeing demonstrator project To compensate the low thrust a glider design has been selected. The project is lead in cooperation with a fuel cell company in the Uk Intelligent Energy.

6.2.2. UAV As presented in (Herwerth, C. 2006), small UAV have been designed and produced to work on fuel cell engine, however, it is easier to implement this technology on a UAV as the payload is lighter as in a passenger aircraft. This source uses a fuel cell providing around 500 W for an 11 kg aircraft. Aero environment and the NASA have developed the Helios project. This aircraft was solar powered during the day and fuel cell powered at night. The fuel cells were regenerative, i.e. generated hydrogen on solar power during the day.

7. Conclusions At the current state of the art, PEMFC have the best performance figures compared to other types of cells, and is undeniably the preferred type for automotive applications. However direct methanol fuel cells are significantly less developed and might be an interesting alternative, as it simplifies the production issues and storage but safety issues remain the same. The current specific powers available from PEMFC are around 1.2 kW/kg, and according to (Wentz, W.H 2005), a specific power of 0.625 kW/kg for the whole propulsion system (including inverter and motor) are required to propel small single engine aircraft, without reducing the payload. These requirements correspond to an “intermediate to advanced” technology, and are demonstrated in (Wentz, W.H 2005).



The power requirements for a business jet and are 0.84 kW/kg, and are not currently met.

Table 3 : Power requirements, (Wentz, W.H 2005)

A lot of information has been provided by automotive publications, but as all car manufacturers invest large amounts of money, state of the art information are not always available. Fuel cell powered cars are much more mature technology than electric aeroplanes, and are likely to impose the hydrogen production and infrastructure requirements before the commercial release of fuel cell powered aircraft. Concerning production of hydrogen, different technologies are available but the best one is electrolyse using renewable energies. Again, this technology is mature enough but requires a lot of infrastructure and investments which can be accelerated by the commercialisation of electric cars.


Fuel-cell powered aircraft Aircraft type and missions

Aircraft type and missions

1. Introduction: After reviewing the different fuel cell technologies available it was then possible to have an idea of the size of aircraft that could be designed using fuel cells. Electric motors and DC/AC converters technologies have also been investigated, so that it was found that components up to 200 kW were easily available. Two aircraft configurations were investigated: a low speed reconnaissance aircraft and a 4 seats family aircraft. For each configuration different fuel cell systems were investigated ranging from commercial materials available at the time to systems based on power densities quoted from literature.

2. Reconnaissance airplane: The baseline aircraft for this case study is the Edgley Optica. This is a low speed aircraft for reconnaissance, with accommodation for 3 and powered by a ducted fan. It has been decided to investigate this airplane because of its relatively low power requirement, the ducted fan, which combined with a fuel cell system, could result into a very quiet propulsion system.



Figure 12: Edgley Optica

2.1. Performances It is powered by a Textron-Lycoming piston engine which delivers 193 kW, and weights 246 kg. It is a relatively old aircraft as it first flew in 1979. The table below presents a summary of the main performances of the aircraft. Engine Power 193 kW TEXTRON LYCOMING 540-4VA5 Cruise speed 103 knots at 70% Weight 248 kg 86 knots at 50% Power density 0.78 kW/kg Power plant mass 26.2 % Empty Weight 948 kg Max T/O 1315 kg Range 619 km Wing span 12 m Ceiling 14,000 ft Aspect ratio 9.09 Length 8.15 m Height 1.98 m Fan Diameter 1.22 m

Table 4: Optica's main performance features



2.2. Fuel cell installation Three types of fuel cell systems have been investigated, each of it uses a fuel cell stack as primary source of DC current, a converter to transform continuous current into alterative current and then an AC motor, as AC motors are much more weight efficient as DC devices. The first option uses a Ballard Mark 902 Fuel cell derivative, as used by Bekiaris in 2002, scaled up to 193 kW. This product is available, as scaling of the basic 902 can be purchased from Ballard. It also uses an integrated drive unit combining converter and electric motor, the A600V250 which produces 225 kW of continuous shaft power at 2100 rpm. As the motor speed is only 2100 rpm no gear box is required, thus saving weight. However this later device has been designed for a bus rather than an airplane so that little consideration has been made for the weight. The second option features Mark 9 SSL stacks, another Ballard product rating 1.24 kW/kg, with two Siemens converters for automotive applications, and two Zytek electric motors both rated at 100 kW. Each stack produces 21kW of electric current for 17 kg, so that a modular design can be achieved, with two independent systems for each motor, so that a more reliable propulsion system is designed. The last option is based on the power densities found in different papers, proved in laboratory, but not currently available. The figure of 1.5 kW/kg has been quoted from Berton et al (2003). According to Bekiaris converters can reach a density of 14.58 kW/kg and a high temperature superconductor motor has been selected, achieving the best performance.



Commercial Solution Fuel Cells Inverter Motor Output

300 A 2100 rpm Ballard Mark 902 scaled up

0.89 kW/kg 700 V

Ballard A600V250 electric drive 225-250 kW 1060 Nm

216.9 kg 411 kg TOTAL 627.8539 kg 66.2 % 0.31 kW/kg Alternative Commercial Solution With two engines Fuel Cells Inverter Motor

Ballard Mark 9 SSL

10 Cells Stack of 21 kW each (17.3kg)

No Inverter required

Two Zytek BLDC electric motors of 100kW each

173 kg 0 kg 42 kg TOTAL 215.00 kg 22.7 % 0.90 kW/kg Advanced Solution Fuel Cells Inverter Motor Output

Advanced FC Stacks ex Gen III

Assumed Power Density : 1.5

Inverter based on Bekiaris research on Zytek products Assumed Power Density : 14.58 kW/kg

HTS High Temperature Superconductor Electric Motor Assumed Power Density : 5.91 kW/kg

128.67 kg 13.24 kg 32.66 kg TOTAL 174.56 kg 18.41 % 1.11 kW/kg

Table 5 : Possible conversion of Optica to Fuel Cell



It can be noticed from Table 5 that current fuel cell technologies allow better performances that piston engine. However it must be kept in mind that the cooling system of the motor and converter have not be accounted for, so that finally fuel cell might be equivalent to piston engine technology, and that this aircraft has been designed in the late 70’s.

3. 4 Seat Family aircraft

3.1. Performances The baseline for this study is a Cessna 172 Skyhawk it is a very old aircraft but has good performances and low power requirements. It is powered by a Lycoming O-320 engine rated at 120 kW. It is widely used for leisure and private use. The performances are summarised below.

Figure 13 : Cessna 172

Engine Power 120 kW Lycoming O-320-H2AD Cruise speed 122 knots max Weight 114 kg Power density 1.05 kW/kg Power plant mass 17.6 % Empty Weight 649 kg Max T/O 1043 kg Range 1065 km Wing span 10.92 m Ceiling 14,100 ft Length 8.21 m Height 2.68 m

Table 6: Cessna 172 Performances



3.2. Fuel Cell installation Again three options have been considered, using commercial data and research publications. For this application as devices sized for 120 kW were not available, commercial products densities have been quoted and devices have been scaled up. The first option uses 6 Mark 9 SSL fuel cells stack, and an electric motor scaled from Zytek products using BLDC motor. The second option is nearly the same but uses a stack designed for automotive applications, by General Motors which is sized for 94 kW continuous and 196 kW peak. It is slightly lighter but does not have better performances as the first option. The last configuration quotes the same density as for the Optica airplane, and results in a small improvement of the power plant which will certainly be overcome by the weight of the cooling system. Fuel Cell has less potential for the 4 seat aeroplane, essentially because of the good performances of the piston engine. In order that a 4 seat fuel cell aircraft is used as a leisure or private aircraft, the fuel cell solution must not be at prohibitive prices. A comparison of the possible operating costs of the aircraft has then been undertaken.



Commercial Solution Fuel Cells Inverter Motor Output

300 A 420 V Ballard Mark 9 SSL 6

Stacks 21 kW each

No inverter required

Based on Zytek products scaled up power density of 5kW/kg

102 kg 0 kg 24 kg TOTAL 126 kg 19.4 % 1.53 kW/kg Alterative Commercial Solution Fuel Cells Inverter Motor Output

300 A 420 V

Gen III from general Motors Rated at 94 kW nominal up to 129 peak power

No inverter required

Based on Zytek products scaled up power density of 5kW/kg

100 kg kg 24 kg TOTAL 124 kg 19.1 % 1.56 kW/kg Advanced Solution Fuel Cells Inverter Motor Output

300 A 420 V 1.5 kW/kg

14.58 kW/kg based on Bekiaris Work

HTS High Temperature Superconductor Electric Motor Assumed Power Density : 5.91 kW/kg

80 kg 8.23 kg 20.30 kg TOTAL 108.535 kg 16.7 % 1.78 kW/kg

Table 7 : Conversion of Cessna 172 to Fuel Cell



4. UAV Designing a fuel cell powered UAV was also one of the possibilities investigated. FC UAV’s have already been designed in the past, mainly prototypes versions such as the Helios project which was solar powered during daytime and fuel cell powered at night. During the day, extra power generated by the solar cells was used to regenerate hydrogen, thus creating a virtually infinite range aircraft. Smaller aircraft can be designed for fuel cell systems, as in Herwerth (2006), but it has been decided that it would be more interesting and innovative to look at a manned aircraft, as the performances allows such design.

5. Direct Operating Costs Before designing a fuel cell powered aircraft, one question must be answered: Why would an electric aircraft be bought? The usage of an electric aircraft is very much dependant on its price, acquisition and operating cost. In the case of a private aircraft, or leased to private pilots the operating cost is critical to the success of an electric airplane. Based on a Cessna 172, the operating costs were compared with a piston engine propulsion system. As hydrogen is not yet widely spread, its cost is still high, but this study is based on projected prices for 2010.

5.1. Fuel prices Assuming a price of 5 $/US gallon, for the 100LL, a range of 834 km for a speed of 222 km/h, and a fuel capacity of 143.8 L without reserves, the operating cost of fuel for a Cessna 172 has been calculated around 50.56$/h. The current price of fuel is 3$/ gallon, and considering what has been discussed in the literature review for oil prices trends, this estimation is reasonable. For the fuel cell system a fuel specific consumption of 2.5 g/s has been selected (This figure has been obtained by the same process presented in the Power plant Chapter. Initially the calculations assumed 5g/s based on an Xcellis engine (Bekiaris, 2002) but the calculations have been updated. For the same performances (speed and range) a consumption of 9 kg/h of hydrogen has been calculated. The price of hydrogen has been investigated for 2010-2015 projections. The price of hydrogen depends on the level of public investments in infrastructures of production and distribution. Predicted cost depends also on the method used to produce hydrogen (steam reforming, electrolysis) and the distribution type (decentralised, centralised plus pipeline network).



Considering the results of Colozza’s studies (2002), a sensitivity analysis has been carried out, looking at direct operating costs for a fuel price ranging from 2.35 $/kg to 6.97 $/kg. The operating cost of running a fuel cell powered aircraft has then been found to be comprised between 18 $/h and 63 $/h. Assuming that the crude oil price continues to rise, and based on a lower estimation of the hydrogen price it can be proved that a fuel cell aircraft would lead to lower operating costs. Moreover, by 2010, the sfc of fuel cells is likely to improve and the value retained for this analysis is quiet conservative. Assuming that the worst price is investigated, fuel cell powered aircraft would remain economically unviable.

Hydrogen price

0

1

2

3

4

5

6

7

8

Remote

refor

mer an

d liqu

ifier (2

7)

Remote

refor

mer an

d liqu

ifier (2

70)

Loca

l refor

mer an

d pipe

line (

27)

Loca

l refor

mer an

d pipe

line (

270)

On site

natur

al ga

s refo

rmer

(2.7)

Partial

oxyd

ation

of oi

l (2.7)

Electro

lysis

of wate

r (3kg

/day)

Methan

ol ref

ormati

on (2

.7)

Method of production (tonnes/day)

Pric

e ($

/kg)

Table 8 : Hydrogen Price, (Colozza, A.J. 2002)

5.2. Acquisition costs Acquisition costs and ownership cost are also part of the operating costs and have been investigated. Again the baseline is a Cessna 172 type aircraft, powered by a Textron Lycoming 320. The airframe is assumed to be the same, and therefore only the power plant acquisition costs have been investigated. The current acquisition cost of such an engine is 9000 $.



According to the US department of energy, and the roadmap established to introduce fuel cell economy, the target price for a propulsion system is 45$/kW by 2010. This leads to an equivalent engine price (stack, converter, electric motor, compressor and cooling system) of 5400 $. According to Ballard Power System, the price of manufacturing only the stack will be by 2010 of 30 $/kW which confirms the previsions of US DOE. Finally, the acquisition cost of a fuel cell system, is likely to be by 2010 less than a piston-engine. This reduction in costs is mainly due to the series production of fuel cell for automotive applications.

5.3. Maintenance costs No reliable data has been found on maintenance costs, which are likely to be high for fuel cells. However US DOE has set a target of 5,000 h life time for commercial fuel cells by 2010 and this target is likely to be achieved as Ballad recently announced having developed a stack working for 2000 h.

6. Choice It has been demonstrated in this chapter that fuel cell system are becoming more and more competitive with piston engines, however this study does not account for cooling system weights and complexity. Compared to relatively old designs, fuel cells can be an improvement, or at least have equivalent performances. It seems that is sufficient public money is invested; a fuel cell powered aircraft may become economically viable. Finally, after discussions with Pr Fielding it has been decided to design an Optica type reconnaissance aircraft, scaled up to accommodate 4 passengers. Such aircraft would need high visibility capabilities and low velocity flight. It could be used for many different missions, ranging from coast guarding reconnaissance, police reconnaissance, and surveillance missions. It could also be used for leisure by private pilots, it would have sightseeing qualities. If the overall design of the Optica is kept, featuring a ducted fan and a fuel cell system, the engine noise could be reduced to levels inaccessible to piston engines, providing maximum comfort and silent surveillance flight.


Fuel-cell powered aircraft Requirements and design options

Requirements and design options

1. Light aircraft survey Once that the aircraft type has been selected, it was important to decide the target performances. In order to be competitive, a fuel cell powered aircraft used as reconnaissance or private vehicle, must achieve equivalent performances as its piston engine competitors. A simple survey has then been carried out in order to select the target performances of the aircraft. The survey takes account for 15 4-seats light aeroplanes, currently produced, and using the latest technologies available. The study has then been completed with reconnaissance aircraft. Reconnaissance types are slightly older, but allow a better understanding of performances required for surveillance. Geometry, power plant, and performances have been listed and maximum, minimum values, and average values have been looked at. Types of aircraft have been separated to emphasize the special features of reconnaissance aircraft. Main data is extracted from Jane’s all the world aircraft database.

2. Basic Requirements

The basic requirements have been drawn from the intended payload. 100 kg have been allocated for each passenger, leading to a payload of 400 kg. Reconnaissance equipments are possible, but allowing for a reduction of the number of passengers.



Empty Weight without Powerplant

0

100

200

300

400

500

600

700

800

900

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5

Passengers

Wei

ght

4 SeatsRecon

Figure 14: Empty Weight without Power plant

Then, empty weight and take off mass have been deduced, from methods given by Stinton (2001), combined with average values of the survey. Allowing for a higher power plant weight of 350 kg, an empty weight of 800 kg has been selected and a takeoff mass of 1250 kg for 200 kW. A parametric study has also been undertaken to identify critical performance parameters, such as wing loading and specific thrust.



Aspect Ratio / Span

0

2

4

6

8

10

12

14

0 2 4 6 8 10

Aspect Ratio

Span

(m)

12

4 SeatsReconnaissance

Figure 15: Aspect Ratio and Wing Span

Following the global trend on a higher aspect ratio for reconnaissance aircraft, an aspect ratio of 9 and a span of 12 meters have been selected as primary values.



Perfomances

0

2

4

6

8

10

12

14

0 20 40 60 80 100 120 140Wing Loading (kg/m2)

Take

off W

eigh

t / R

ated

Pow

er (k

g/kW

)4seatsReconnaissance

Figure 16: Power Loading and Wing Loading

Power / Speeds

0

50

100

150

200

250

300

350

400

0 50 100 150 200 250

Engine Power (kW)

Spee

ds (k

m/h

)

Cruise 4 SeatsCruise ReconnaissanceStall 4 SeatsStall Reconnaissance

Figure 17: Stall and cruise speeds function of engine rated power



Finally, assuming a specific thrust of 8 kW/kg and a wing loading of 686 N/m2, an engine rated between 179 kW and 139 kW was selected, for a wing area between 15.62 and 20.8 m2. As a first estimation, average values of 156 kW and 17 m2 have been selected. Other performance figures are presented in the table below. Fuel Cell engine Cruise speed 190 km/h 52.78 m/s Engine Power 156 kW Stall speed 80 km/h 22.22 m/s Weight 350 kg Maximum 250 km/h 69.44 m/s Empty Weight 800 kg Range 800 km Max T/O 1250 kg Ceiling 15,000 ft Wing span 12 m T/O distance 300 m Length 8.5 m Climb Rate 250 m/min Height 1.65 m

Table 9 : Target performances of a reconnaissance electric aircraft

3. Design options Four different options have been considered, each having advantages and drawbacks. The main parameters driving the design are a high visibility cockpit, and low velocity capabilities. An easy handling of the aircraft with good stability should also be achieved for its use by private pilots. The four designs have been drawn around a common unswept wing, based on a 12 m span at a scale of 1:50.

3.1. Optica design The Optica baseline features an unswept wing, high visibility cockpit, a ducted fan and twin tails booms, supporting a single elevator. It has been described in a previous section. To balance the weigh of the cabin and payload, fuel cell stacks are placed in the fuselage, between the wings, and the electric motor combined with the ducted fan at the rear of the fuselage. The fuel is stored as liquid hydrogen in the wings, avoiding a too large centre of gravity displacements. Due to their small height the fuel cell stacks modules could have been stored in the wings, but to achieve a longer range, stacks are stored in the fuselage. Compared to a



piston engine aircraft, the fuel tanks will need insulation, thus reducing the available volume for the liquid hydrogen. To allow low velocity flight, significant flaps can be used, combined with a wing profile optimised for slow motion. Taking advantage of the forward position of the cabin, the CG can be positioned in front of the aerodynamic centre of the wing so that longitudinal stability is achieved. This existing design allows for a very low noise, as well as an excellent visibility which could be a real asset in police surveillance missions; however the ducted fan has a significant weight penalty over a propeller system but better efficiency at low speed. The available hydrogen storage volume is lower than in other concepts, and the available volume for luggage is very narrow. Besides, the use of a mid-height wing could prevent the use of surveillance devices, such as IR cameras, mainly for ground clearance issues in cross wind situations at take-off.

3.2. High wing, twin pusher propeller The second option is made out of a high wing assembled with a high visibility fuselage. Again, the passenger and the pilot are situated at the very front of the fuselage. The aircraft is powered by two pusher propellers driven by electric motors, which do not interfere with the observer’s vision, even for an observer in the rear seats. This option leaves a great amount of space available in the fuselage, to store hydrogen and might enable to achieve better ranges. Fuel is also stored in the outer wings. The stacks and converter are situated around the centre of gravity in the fuselage just in front of the auxiliary hydrogen tank. The use of pusher propellers prevents from using significant flaps so that this drawback would have to be balanced by a wing profile designed for low speed. One other option would be to accept the visibility penalties from the propellers and then use tractor propellers, allowing for the use of flaps. This design has the advantage of being fairly classical and therefore an easy handling could be appreciated by private pilots. A common design also allows for reduced manufacturing costs by using well known and cheap processes, risks due to innovative techniques are also reduced to minimum. However noise and slow motion performances are lower than with the first design.

3.3. Canard configuration

The canard configuration uses again a high wing profile, combined with a rear engine, and pusher propeller. The foreplanes allow a good visibility and space is available for the power plant at the rear of the cabin.



The main drawback of this design is its high propeller which would result into a thrust offset and therefore a difficult handling due to a negative pitching moment. The canard would allow for a low mass design, but possible advantages would be reduced by the structure supporting the engine nacelle in the tail. Although the visibility is good, the canards still stands in the line of sight of the crew. The canard design is not very common for light aircraft and there might be a risk in guarantying handling capabilities of such an aircraft. This design allows also a more compact fuselage, and a reduced length.

3.4. Joint wing concept This concept uses a “Wolkovich” joint wing, as presented in Appendix A. The electric motor is located at the joint of the aft wings, supported by the fin, and drives a pusher propeller. The stacks can be located in the mid-fuselage, and hydrogen stored in the wings. The need for tail plane or foreplanes is removed, as elevons on the forward wings can be used as elevators when moved together and ailerons when moved differentially. There are also control surfaces on the aft wings, which provide yawing control. The cockpit is again, a high visibility design, located at the front of the aircraft, and balanced by the wings. As the wings are located in the rear part of the aircraft, the pitching moment is negative. A joint wing concept enables to achieve higher lift-to-drag ratio, and better yawing control, thanks to aft wing control surfaces. According to the same reference, the drawbacks of a joint-wing concept are reduced fuel capacity, as the aerofoil section is thinner, and a structural weight penalty. The sfc of a fuel cell powered aircraft is lower than of a piston engine, so that less fuel is required, but as insulation consumes volume, the overall hydrogen capacity of a joint wing aircraft is lower, leading to lower ranges. Again, the high engine mountings may result in a large thrust offset and therefore worst handling capabilities, combined with the specificity of the joint wing concept. This concept holds a great risk potential because of its new technology and handling capabilities, but also promises a larger interior space. Developing such a new concept would also have a price, higher than other designs, and such investments could not be allowed on such a small project. Again for flight safety, standard well known designs are safer than a novel concept.

4. Concept choice To select the concept which would be further developed, a simple decision table has been set up. This decision table is based on different attributes that the ideal aircraft



should have to fulfil its mission. Attributes ranges from producing costs, customer appeal, weight, noise or cockpit visibility. Each of these attributes has been associated with a coefficient measuring its importance, so that by adding all the attributes marks multiplied by the coefficient a reasonable indicator of the concept performances is obtained to allow a reasonable decision. The decision table is presented below, and most of the advantages or drawbacks of each concept are discussed in the individual presentation of each concept.

Attribute Weight Optica Design

High wing, twin engine

Canard Configuration

Joint Wing concept

Low Mass 10 6 8 7 5 Passenger appeal 8 5 7 7 9 Safety 10 6 7 6 6 Risk 8 7 8 6 5 Cost 10 6 8 6 3 Handling 7 6 7 3 4 Visibility 5 9 7 6 7 H2 Storage Volume 5 5 7 7 5 Noise 5 9 4 6 6 TOTAL 433 489 410 365

Table 10: Decision table for concept choice

From this table, it appears clearly that the Optica design, and the high-wing concept are the best available. Mainly for weight and cost considerations the high wing concept has been chosen, to avoid combining airframe high development costs with power plant development costs.


Fuel-cell powered aircraft Power plant Architecture

Power plant Architecture

1. Architecture

The power plant has been designed to provide 156 kW of power, using Ballard fuel cell stacks. The design is based on 8 Mark 9 SSL stacks providing each 21 kW of power with the best specific power available at the moment. The data used in this study have been provided by the commercial datasheet presented in Appendix C, but some figures were missing and assumptions had to be made, such as the efficiency of the stack. Minimum assumptions were made and other missing figures were calculated. Manufacturers were contacted for more information but some of them did not accept to give more information. Among the assumptions made:

Fuel Cell efficiency of 50 % Operating pressure of 1.5 bar Operating Temperature 80ºC Electric motor efficiency 93 %

A closer attention has been paid to cooling and compression units, as advised by Mr Orti and Nieves from Boeing. Therefore accounting for the efficiencies of different systems, the dissipated powers have been calculated and the requirements for cooling have been investigated.




Fuel Cell Stacks 168 kW

Efficiency 0.5 80ºC

LH2 Tank

-253ºC

Heat Exchanger

43.6ºC -253 ºC 2.902 g/s

Supercharger

Eaton M62

Ram Air 100g/s, 20ºC

100 g/s179ºC

43.6ºC

Inverter 168 kW

560 VDC300 A

22 kW AC

Electric Motor

Efficiency 0.93

H2O, 80ºC 26.1 g/s

Air Electric Power Hydrogen Hot water Hot Coolant Cold Coolant

Heat Exchanger Radiator Coolant,

Water, 1.07 l/s

Propeller

Humidifier

Onboard Systems 146 kW 134 kW

Figure 18 : Fuel cell powered aircraft power plant architecture


2. Hydrogen fuel system

2.1. Hydrogen specific consumption

The fuel specific consumption is a critical value for the design of an aircraft, but specific consumption of Ballard fuel cell stacks were not found on the datasheets. A simple estimation has been performed. According to Colozza (2002), the specific fuel consumption of a stack can be calculated by the equation:

965002×

×=

IANsfc p

H Mol/s

Where Np is the number of cells in the stack, IA the current density times the area i.e. the current. The number 2 stands for the number of electron in the chemical reaction of the fuel cell, and 96500 J is the energy of a mole of electrons. The number of cells is calculated dividing the stack voltage (70 V), by the voltage of a single cell. The voltage of a single cell is derived from the efficiency (assumed of 50%):

η×= 2.1V = 0.6 V Leading to Np = 117 The sfc for a single stack is 0.363 g/s, leading to a global sfc:

Sfc = 2.902 g/s

2.2. Hydrogen quantity To achieve a range of 800 km at a speed of 190 km/h, the travel time is just over 4 hours. The required hydrogen mass stored onboard, allowing for 10 % reserve has been calculated 55 kg of hydrogen. To store such an amount of fuel, liquid hydrogen storage is the only available solution, for available space aspect. It has been decided to store the hydrogen in the wings, which would comprise integral fuel tanks and additional thermal insulation. It must be noticed that the quantity of hydrogen stored onboard has been derived referring to the maximum take-off power. As it is unlikely that the engine is used at full power during cruise, the sfc should be less, and consequently the range of the aircraft increased.



2.3. Hydrogen tanks The hydrogen is stored in wing tanks. In order that hydrogen is kept liquid, the temperature inside the tank must be kept at -253 ºC, and tanks are protected by a layer of insulating material. The insulating materials consume space normally used to store fuel. Colozza (2002) presents a number of available materials to insulate hydrogen tank. Clearly from this work the best material available is an Evacuated aluminium foil and glass paper laminate, or Evacuated aluminium foil separated with fluffy glass mats. Results are presented in Fig 19 below.

Insulation Materials Properties

0 50 100 150 200

Polymeth

acryl

imide

Polyure

thane

Polyvin

alchlo

ride

Polyure

thane

and c

hopp

ed gl

ass f

iber

Aluminu

m foil s

epara

ted w

ith gl

ass

Aluminu

m foil a

nd gl

ass p

aper

lamina

te

Evacu

ated s

ilica p

owde

r

Density (kg/m3)

0 0.002 0.004 0.006 0.008 0.01 0.012Thermal Conductivity (W/m.K)

DensityThermal Conductivity

Figure 19: Material Properties, from Colozza (2002)

The hydrogen is stored at a pressure of 20 psi (1.44 bar), so that no oxygen penetrates the tank by pressure or refuel valves. A preliminary study has been performed to determine the amount of insulation required, and the associated weight penalties. Based on the wing geometry of the Optica aircraft, available quantity has been calculated, and space available for insulation materials deduced. Assuming a quantity of 55 kg, (corresponding to 4.5 hours flight time, at maximum power), 0.39 m3 were required per wing. The hydrogen storage concept has been derived from those proposed by Bekiaris. The fuel is stored in the wings between the spars.



LH2

-253ºC

Front spar (10% chord)

Insulating material (2cm)

Rear spar (65% chord)

Figure 20 : Liquid Hydrogen storage concept

The heat transferred is assumed to be only transferred by conductivity, (no fluid exchange with outside). The heat transferred to the hydrogen is then proportional to tank area and temperature difference.

)( inttan TTkAQ extk −××=

Liquid hydrogen is stored near its vaporisation point; therefore temperature rise energy is neglected compared to vaporisation heat.

22

HH C

QN =

Assuming a differential temperature of 273 ºC (outside temperature of 20 ºC), it has been calculated that 0.73 kg of hydrogen were transformed to gaseous from during the flight time. It would be unsafe to park a hydrogen aircraft with full fuel tanks for longer than 5 hours, because vaporised hydrogen would raise the pressure in tanks, and create fire risks. This would imply to fuel before each flight and de-fuel after each cycle, which leads to higher maintenance compared to a gasoline aircraft. Assuming a front spar at 10% chord and a rear spar at 65% chord, a tank chord of 0.715m has been deduced. As the maximum thickness of the wing is 0.2m, the average thickness in the tank has been arbitrary assumed to be 0.13m for a span of 6m. Therefore, a volume of 0.35 m3, and a surface of 10.35 m2 for each tank have been calculated. The hydrogen tank penalty is then 16.52 kg for both wings, and 37 % of the fuel tank volume is used for insulation. A detailed analysis has been performed on the insulating material thickness, and the results are plotted below:



Hydrogen storage weight and losses

0

5

10

15

20

25

30

35

0.01 0.015 0.02 0.025 0.03 0.035 0.04

Insulating material thickness (m)

Wei

ght (

kg)

0

1

2

3

4

5

6

7

Hyd

roge

n Lo

sses

(%)

Weight Penalty Volume Stored Hydrogen losses

Figure 21 : Hydrogen storage weight and losses

2.4. Fuelling refuelling It should be kept in mind that hydrogen is mostly dangerous in gaseous phase mixed with air. Therefore, air must be prevented to penetrate the fuel tank. Ideally when empty the hydrogen tank has to be filled with an inert gas such as nitrogen. When the tanks are filled, heat is transmitted to the hydrogen through the insulating material at a low rate. The heated hydrogen turns to gaseous phase, thus raising the pressure in the tank. Ideally this gaseous heated hydrogen should be consumed by the fuel cell. Heat transfer from the outside to the tank creates hydrogen losses, but is necessary to maintain the pressure inside the tank, as hydrogen in consumed. Eventually at the end of the flight, almost all hydrogen would have been consumed but the remaining hydrogen would be mixture of gaseous and liquid phases. The tank has to be equipped with pressure release valves. As soon as too much hydrogen is turned into gas, the pressure rises above 20 psi. Pressure valves should then open to let the exceeding hydrogen out. Hopefully, hydrogen leaks very fast due to its small molecular radius, and small viscosity, so that very low heat losses should happen when the valves open. Before maintenance it is essential to empty the tank by purging it with nitrogen, in an open air area, so that hydrogen remaining is exhausted and dissipates upwards. The



same fire rules should apply when purging the hydrogen tank as when working with gasoline. After maintenance, air has to be exhausted before refuelling with hydrogen.

2.5. Tank Materials Storing hydrogen in liquid state at -253ºC, creates some thermal constrains on the materials from which the tank is made. Very low temperatures result in different material properties in strength and fatigue properties. Fuelling and de-fuelling the hydrogen tank will results in thermal constrains cycles, which may have detrimental effect if the fatigue properties are insufficient. This paragraph deals with both metallic and composite materials, even if the selected insulating material is aluminium based, there has been significant work undertaken in the past years to reduce tank weight by using carbon fibres laminates.

2.5.1. Tensile properties For metallic materials, aluminium alloys have been investigated mainly for weight purposes. Generally speaking tensile properties of allows improve at cryogenic temperatures, both yielding strength and ultimate tensile strength. These results are confirmed by both Vander Kooi et al (1999) for C-458 Al-Li alloy and Tetsumi et al (2001) for 5000 series aluminium alloys. According to Nettles et al (1996) who studied the impact of cryogenic temperatures on composite materials, a quasi-isotropic laminate would have deteriorated properties at cryogenics temperatures. The average tensile modulus is slightly increased (4%) at cryogenic temperature for liquid nitrogen storage. However the tensile strength is reduced by 9 %. It can be assumed that strength properties will continue to decrease as the temperature for hydrogen storage (20º K) is lower than for nitrogen storage (110 ºK), unfortunately data were not available in the time of this project.

2.5.2. Fatigue properties Tetsumi et al (2001) have tested two manganese based aluminium alloy A5083 and A5183 for fatigue properties at cryogenic temperatures. The tests demonstrate that these two alloys have better fatigue properties at cryogenic temperatures, and allow higher repeated stress levels at cryogenics temperatures. Moreover, compared to stainless steels, these two alloys have better fatigue properties at low temperature, because steels properties decrease rapidly at cryogenics temperatures.



Figure 22 : Fatigue properties of 5000 series Aluminium alloys at cryogenic temperatures, From Tetsumi et al (2001)

Concerning composite materials, Nettleset al (1996) focussed on matrix-dominated properties of the composites, and claims that thermoplastics resins have better properties in terms of strength and damage tolerance. Thermosets resins on the other hand have lower properties at cryogenics temperatures. Initiation of micro-cracks in the structure is reduced at lower temperatures.

2.5.3. Fracture properties One major concern when using materials at cryogenic temperatures is the toughness at cryogenic temperatures. For fracture properties, only data on the Al-Li alloy C-458 has been found. Vander Kooi et al (1999) studied the fracture toughness of this new alloy from ambient temperature to cryogenic temperatures (4º K, corresponding to liquid helium). Al-Li being slightly anisotropic, toughness has been tested in two directions, moreover ageing has been investigated by exposing some of the samples to an ageing of 1000 h at 355º K. Results are reproduced below, and show that toughness reduces at cryogenic temperatures, but the changes are minor. Compared to previous Al-Li alloy, the reduction of toughness is less important. Ageing does not have a significant effect on results



Figure 23 : Fracture toughness at cryogenic temperatures for Al-Li alloy, Vander Kooi et al

(1999)

2.5.4. Thermal compression Another issue to be addressed is the thermal compression at low temperatures of materials. Allowance for compression should be made in the design of the hydrogen tanks, which may create some issues for the attachments of the tank to the primary structure. The effect of temperature drop when filling the tank after an empty period should be investigated, because rapid temperature changes will induce important stresses in the structure of the fuel tank.

3. Cooling system As any power plant, the fuel cell system is not 100 % efficient and therefore produces heat which needs to be evacuated. A water system has been investigated, because of compatibility issues with the stack material, however any coolant not deteriorating the stack can be used, the higher the specific heat is, the better the cooling system will be.

3.1. Fuel Cell Stack cooling Fuel cell stacks produce heat as well as electric power, and this heat needs to be evacuated. Assuming an efficiency of 50 % means that as muck power is produced as heat as in its electrical form. The stacks are water-cooled: water is routed within the metallic electrodes and cool the stack down. Very simple calculations have been performed to determine the water flow required to cool down the system, in order to account for the weight and energy penalties of such a system. Such a system activated by a pump and electric motor, consumes energy



which is taken from the stack, just the same way shaft power or bleed air is taken from a turbofan engine, leading to penalties. The power taken by a flow of fluid depends on the intake temperature and stack temperature of the cooling system. An inlet temperature of 20ºC has been assumed for the water, and stack temperature of 80ºC. The exchanged quantity of energy is then:

)(2 inoutOH TTCpqP −××=

q stands for the water flow, Cp is the specific heat of water. To dissipate 168 kW of heat a flow of 0.67 kg/s is required.

3.2. Inverter The inverter was designed on the basis of a TIM-600 inverter. This baseline converts a continuous current between 80 and 400 V into an alternative current and is rated for 100 kW. The inverter used is a derivative from this model, scaled up to 168 kW. The weight of the scaled up unit is 16.8 kg. The unit is water cooled and, again, scaled up requires 0.22 kg/s of water mixed with glycol.

3.3. Electric Motor The same process as for the stack has been used, to define the water flow required to cool the electric motor. An efficiency of 0.93 has been assumed, according to Larminie, J. (2003), leading to heat power of 17 kW to be evacuated. There are different options for the cooling of the electric motor, either air cooled or oil cooled, or water cooled. The motor is assumed to be a Zytek product delivering 100 kW for 21 kg. Unfortunately, the manufacturer did not provide any additional information, although they have been contacted several times.

3.4. Water pump Once the total flow of coolant has been calculated, a water pump has been selected. The pump selected hasn’t been designed for aerospace applications, and unfortunately makes extensive use of steel. However, the unit weight a reasonable amount of 6.8 kg. The pump provides 58.6 l/min (0.97 l/s) at a differential pressure of around 1 bar. It supports fluids from -15 to 100 ºC. Detailed specifications are presented in annex. It should be noticed that at full power, this device uses 400 W which are also taken from the fuel cell stack.



Figure 24: Water pump for fuel cell stack cooling system

As the device weight can easily be optimised using aluminium instead of steel, so a reasonable weight of 12 kg including the water itself can be assumed for the whole cooling system, excluding the heat exchanger.

4. Oxygen intake

4.1. Fuel cell feeding As in piston engine, in order that the chemical reaction happens and produces power, oxygen must be fed into the engine. The fuel cell stack has to be fed with oxygen, taken from ram air. Oxygen has to be fed in the right proportions corresponding to the reaction of the fuel cell. Half as much oxygen has to be provided, unfortunately oxygen is heavier than hydrogen, and is only 21 % of ambient air; consequently the intake mass flow of air is high. It has been calculated before that 1.45 mol/s of hydrogen were consumed at full throttle:

73.021

202 =×= Hqq mol/s

Therefore,

45.321

1002 =×= Oair qq mol/s of air

The atomic mass of air is 29 g/mol

qair = 0.1002 kg/s In order to achieve higher levels of specific power, fuel cells operate at low pressure. Although the operating pressure cannot be found for the stack selected, it has been assumed that the pressure is 1.5 bars, as in Mark 902 stack, another product from Ballard.



4.2. Compression devices Again, this is another commonality from fuel cell with piston engine, is the use of pressurised air to increase performances. Such devices are named superchargers, and are used in powerful cars, such as the Lotus Elise. The technology is quite mature compared to fuel cell engines, and different types of superchargers are available. A roots supercharger is the simplest type air is compressed by two rotating gear wheels. The air stream is normal to the axis of the rotating devices. Only small pressure ratios can be achieved. The Lysholm or screw supercharger works on the same principle, but the air stream is parallel to the rotating axis of the screws. Screw manufacturing needs precise tools and therefore the device costs more. In the centrifugal type, a wheel rotates at high speed; air is provided at the centre of the wheel, slides against the wheel and comes out at high velocity. The velocity is then transformed into pressure energy in the outlet of the supercharger. The efficiency can be good but only for the nominal air flow. Axial superchargers are exactly the opposite of turbines; a fan rotates at high speed and compresses the gas.

Type

Pressure

ratio Noise Efficiency Flow rates Price

Roots 1.8 - good wide Cheap Lysholm (screw) 8 - good wide Expensive

Centrifugal 4 + good for optimum flow

rate narrow Cheap

Axial 2.2 - good for optimum flow

rate narrow ExpensiveTable 11: Summary of supercharger types and performances

Considering the performances of the different types of supercharger, a roots type is clearly the best for our application. The pressure ratio required is only 1.5, the air flow is likely to change as the throttle setting changes, and it is cheap to manufacture. As the atmospheric pressure drops at higher altitudes, the cell stack pressure is likely to drop as well, but this is no major concerns as the power requirements are not so high during cruise.

4.3. Supercharger selection and performances The selection of the supercharger depends very much on the available data, but an Eaton M62 supercharger, satisfies the flow and differential pressure requirements.



The use of the performance charts, provided by the manufacturer leads to a rotational speed of 8600 rpm, a requirement of 17 hp (22 kW). As the air receives energy through the compression unit, its temperature rises of approximately 145ºC. This energy is then used to heat the incoming hydrogen flow.

Figure 25 : M62 Eaton supercharger

The unit weights 20.86 kg.

5. Heat balance

5.1. Intake of the Cell The fuel cell operates at 80ºC, but if order to start up and the fuels (oxygen and hydrogen) have to be provided at a temperature between 2 and 65ºC. As the hydrogen is stored at -253ºC and the hydrogen comes out of the supercharger at 165ºC, intake temperature has to be regulated. A heat exchanger between air and hydrogen is necessary. Under the assumptions previously explained, a heat balance calculation has been performed.

( )( ) ( ) 0argsup2222 =−××+−×+× inFCererchairairinFCstHHHH TTCpqTTCpCpq

=∴ inFCT 43.6ºC

It should be kept in mind that this calculation has been performed for stoechiometric proportions of fuels, so that if throttle is reduced, hydrogen and air flow should be reduced. If the proportions are kept, the equation is still correct, and the inlet temperature for the fuel cell stack stays 43.6ºC, for any throttle setting.

5.2. Cooling system The stack, the inverter, and the electric motor are water cooled, however the heat has to be evacuated from the coolant. It is necessary to use a radiator, to dissipate the heat, or recycle a small amount of the heat in the ECS for example.



The question of the coolant can then be asked. Why using an intermediate fluid (water) to dissipate heat in air? This system answers to a inherent constraint of fuel cells, in order to maintain performances, a high humidity level must be kept in the stack, whereas air cooling would dry the stack too much compromising the performances.

6. Conclusions and mass summary It has been confirmed that auxiliary systems such as oxygen feed system and cooling devices did not have a negligible weight, and that there were serious heat related issues. Most of the calculations done in this section have been performed on take-off configuration at maximum power; however fuel flows and heat transfers must be controlled at any time of the flight, to ensure that the stack is wet enough to achieve best performances and reach its design life.

kg % Fuel Cell Stacks 138 52.15 Inverter 16.8 6.35 Electric Motor 35 13.23 Supercharger 20.8 7.86 Air conducts and Radiator 20 7.56 Water cooling system 12 4.54 Heat exchanger 5 1.89 Hydrogen tanks penalties 17 6.42 TOTAL 264.6 Global Power to Weight ratio 0.59

Table 12 : Power plant mass summary


Fuel-cell powered aircraft Initial Design

Initial Design 1. Introduction Once the basic geometry and performances of the airplane have been decided, it was essential to design into more detail the aircraft itself. Empirical methods have been used to size different elements of the design, such as empennage areas, wing loading and power requirements. The methodology used has been derived from Howe, (2000), which is based on empirical equations to size main members of the aircraft. The methodology equations are presented in a Microsoft Excel spreadsheet, adapted for different types of aircraft, (piston engine, turbo propellers, and turbofan engine aircraft). The piston engine spreadsheet has been used; however, it has been significantly modified to reflect the specificities of the fuel cell power plant.

2. The baseline optimisation process The baseline optimisation process accounts for the basic requirements of the aircraft: (Landing and take-off length, stall speed, design speed, cruise speed, g limits, and payload) and derives for each mission stage the thrust to weight ratio as function of the wing loading. Eventually, wing loading is defined by the landing case, which itself depends on the approach speed. Approach speed is taken to be the minimum of 1.25 times the stall speed or a calculated approach speed depending on the landing length. In our case, the low stall speed required for the loitering capability had a significant impact on the wing loading. Selecting a higher stall speed would have allowed to increase the wing loading and therefore have a smaller wing area. Specific thrust is defined by the intersection of take-off case characteristic with the previously selected wing loading. Then using user-defined coefficient the mass of each sub-component is estimated, as well as the maximum take-off mass. The two main parameters are the aspect ratio and the thickness to chord ratio. The general purpose of this spreadsheet is to optimise the take-off weight in respect to these two parameters, using the Microsoft Excel solver. The general requirements of the aircraft have been filled using the figures derived from the parametric study and general aviation survey.



3. Modified spreadsheet The purpose of this chapter is to present and justify the modifications of the original spreadsheet, referring to the fuel cell engine. These modifications have been partially derived from Bekiaris (2002), and improved referring to the present state of the art. It might also be used for further work in the future, with up-to-date figures.

3.1. Specific power of power plant Although a more refined analysis of the power plant weight has been carried out later on, the specific power of the installation has been set to 0.59 kW/kg for a first analysis, reflecting the whole fuel cell system weight.

3.2. Fuel cells A specific input has been created in the spreadsheet to reflect the number of Ballard Mark 9 SSL fuel cells installed onboard. From this number the rated output of the power plant is estimated and can be compared to the power requirement estimated by the original spreadsheet. The output power accounts for efficiencies of the stack, inverter and motor. The weight of the power plant is also estimated as a function of the number of stacks and replaces the power plant mass which was previously derived from an empirical coefficient which only reflected piston engines.

3.3. Fuel specific consumption Hydrogen specific fuel consumption is also estimated referring to the number of stacks installed onboard and the efficiency assumed to be 50%. The derivation of the sfc is presented in the power plant chapter. The sfc is significantly lower than with a standard piston engine, so that the need for fuel storage is reduced.

3.4. Start of Climb mass and Landing mass As discussed by Bekiaris, the ratio of start of climb mass to take-off mass has been assumed to be unity, to reflect the low sfc of the engine during taxing and take-off acceleration. The landing mass has been assumed to be 95 % of the take-off mass, due to the small hydrogen weight (55 kg) compared to take-off mass.



3.5. Wing mass penalties The penalty due to the introduction of insulating materials in hydrogen tanks has been estimated in chapter 4, so that a mass penalty of 20% has been decided, lower than the penalty of 50% assumed in previous works. The figure of 20% has been obtained using the 16kg penalty calculated on the Optica baseline wings and on the 77 kg wing mass obtained as a first estimation with the original spreadsheet.

3.6. Available hydrogen mass onboard The available volume for hydrogen storage has been modified. First the original formula has been corrected accounting for the density ratio between AVGAS and hydrogen, hydrogen being nearly 6 times lighter. Moreover a volume penalty of 20 % has been assumed for the insulating materials, as calculated in chapter 4.

4. Results The output of the spreadsheet is mainly the curve of specific thrust plotted against wing loading. The design wing loading is decided by the landing operation, which leads to a loading of 613 N/m2, which is lower than common values for similar aircraft (comprised between 735 and 800 Nm2). The specific thrust is set by the take-off, so as the aircraft is designed for a low speed capability, a low specific thrust of 0.49 is necessary. Then given some geometry of the aircraft, a summary of mass repartition is given:



Weight % Wing 83.81 7.12 Fuselage 161.33 13.70 Tail Unit 20.11 1.71 Landing Gear 61.30 5.21 Power Plant 245.88 20.88 Systems 134.86 11.45 Operational Items 12.00 1.02 OEW 719.29 61.08 Payload 400.00 33.97 Hydrogen 58.33 4.95 MTOW 1177.62 100.00

Table 13: Mass distribution of optimised aircraft

Wing Area 27.10 m2 Wing Apex 2.40 m Wing span 15.62 m Horizontal tail area 7.09 m2 Mean chord 1.74 m Vertical tail area 4.88 m2 Aspect ratio 9.00 Prop diameter 1.73 m Sweep angle 0.00 t/c 0.17 Power 123.44 kW Wing Loading (Mg/S)o 613.00 Specific Thrust (T/Mg)o 0.49

Table 14: Main dimensions of optimised aircraft

Figure 26 : (next page) specific thrust plotter against wing loading


Fuel Cell Aircraft Performances

0.9

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

400 500 600 700 800

Wing Loading (Mg/S)o

Spec

ific

Thru

st (T

/Mg)

o

Take offAcc.StopSec.SegClimbEnd ClmbCruiseMax.speedManLandingGust Sen


Using this analysis the required power of the fuel cells has been found to be slightly lower than expected, and therefore possibilities of reducing the number of stacks has been investigated. For a number of stacks comprised between 5 and 10, the aircraft mass, optimised aspect ratio, and thickness to chord ratio have been determined. It has been found that the optimal number of cells would be comprised between 7 and 8, accounting for inverter and motor efficiency. The final number of stacks is then 8, to provide enough power to feed the cooling and supercharger devices. The curves of available power (from fuel cell number) and required power (from optimisation process) are presented in graph below, together with the overall take-off weight of the aircraft.

Fuel cell aircraft optimisation

0

20

40

60

80

100

120

140

160

180

200

5 6 7 8 9 10

Number of stacks

Pow

er (k

W)

1000

1020

1040

1060

1080

1100

1120

1140

1160

1180

1200

MTO

W (k

g)

Power rated Power required Weight

Figure 27: Rated power and Power requirements

5. Wing Selection

5.1. Investigated profiles Initially the wing profile indicated by the empirical method presented below featured a thickness to chord ratio of 0.26 for an aspect ratio of 7.56. However even if these figures are the best mathematical solution to the equation set, they might not be the better solution. A 26 % thick wing is a very thick profile compared to conventional profiles, for example the Optica wing is a GA(W)-1 profile, 17 % thick and the Cessna 172 Skyhawk is a 12 % thick 2412 NACA profile.



Such a profile would have more drag, due to a thicker profile, adding to the drag induced by the braced wing configuration. It has then been decided to investigate the use of well known wing profiles, such as the NASA developed GA(W)-1 and GA(W)-2. The General Aviation GA(W)-1 wing profile has been designed and studied in the early 70, (McGhee, R 1973) and is 17 % thick. A later version GA(W)-2, is thinner, 13 % thickness to chord ratio, and has slightly better performances at low speeds. This design has been tested in wind tunnel facilities by the NASA (McGhee, R. 1977). Aerodynamic performances of these two profiles are well known and easily available, on a two-dimension basis. A latter study also investigates the use of single slotted flaps and fowler flaps, so that all the required information of theses profiles was available. For our wing profile, the Reynolds number is:

μρVLRe = = 4.18 106

Where, ρ the cruise air density 0.909 kg/m3

V design speed 52.7 m/s L typical dimension (wing chord) 1540 mm μ Viscosity of air 17.4 10-6 Pa.s

5.2. 3D performances The data found in the literature are based on an ideal 2D profile, and Stinton, 2001, gives a simple calculation set to derive 3D performances. Drag is corrected by adding induced drag as a function of aspect ratio:

Π×+=

ACCC l

dD

2

Lift is corrected by modifying the lift curve slope, which is multiplied by a constant factor.

20 +××=

AAfaa w



Corrected Drag Characteristic GA(W)-1, No flaps

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

-1 -0.5 0 0.5 1 1.5 2

Cl

Cd

Infinite Aspect Ratio Aspect Ratio 9

Figure 28: Corrected Drag Characteristic GA(W)-1

The angle of attack corresponding to zero lift remains unchanged, but the 3D model leads to lower lift coefficient. The higher the aspect ratio is, the lower the performances are.

Corrected Lift Curve Slope, GA(W)-1, No flaps

-1

-0.5

0

0.5

1

1.5

2

-15 -10 -5 0 5 10 15 20 25

Angle of attack

Lift

Coe

ffic

ient

Aspect Ratio 9 2D profile (infinite aspect Ratio)

Figure 29: Corrected Lift curve slope GA(W)-1



5.3. Flaps Cruising at low speed for surveillance mission is an important performance, which influences the choice of the wing, but cruising at low speed means using flaps. Different types of flaps have been investigated for both profiles; however, the analysis was limited by the available data. Single slotted flaps were investigated for GA(W)-2 profile, and Fowler flaps were investigated for both options. The different configurations were investigated:

Single slotted flaps, 25% chord, on GA(W)-2 profile, 40º deflection Fowler flaps, 30 % chord, GA(W)-1 profile, 35 º deflection Fowler flaps, 30 % chord, GA(W)-1 profile, 20 º deflection (extrapolated) Fowler flaps, 30 % chord, GA(W)-2 profile, 40 º deflection Fowler flaps, 30 % chord, GA(W)-2 profile, 20 º deflection

Single slotted flaps are simpler, and offer good lift performances, but at the cost of high drag. Fowler flaps are more complex, but have the advantage of providing a first setting where the flaps extends trough the rear of the profile, thus increasing wing area, and lift, without a great drag component. The second setting, full deflection offers a maximum lift for landing, and high drag. It should be noticed that the Optica uses Fowler flaps up to 50 º deflection on a GA-1 profile. The lift curves slopes of each investigated configuration are plotted below:



Lift Curves Slopes

-1

-0.5

0

0.5

1

1.5

2

2.5

3

-15 -10 -5 0 5 10 15 20 25

Angle of Attack

Cl

GA(W)-1 Clean GA(W)-1 Fowler Flap 35% GA(W)-2 CleanGA(W)-2 Slotted Flap 40% GA(W)-2 Fowler Flap 40% GA(W)-2 Fowler Flap 20%

0

Figure 30: Lift Curves slopes for different flap configurations

Some irregular points are due to the clarity of the paper they are taken from.



Drag Curves

-0.2

0

0.2

0.4

0.6

0.8

1

-2 -1 0 1 2 3 4

Lift Coefficient

Dra

g C

oeffi

cien

t

0

0.05

0.1

0.15

0.2

0.25

5

GA(W)-1 Fowler Flap 35% GA(W)-2 Slotted Flap GA(W)-2 owler Flap 40% GA(W)-2 Fowler Flap 20% GA(W)-1 Clean GA(W)-2 Clean

Figure 31: Drag plots of different configurations

Lift Curve GA(W)-1

-1

-0.5

0

0.5

1

1.5

2

2.5

3

-15 -10 -5 0 5 10 15 20 25

Angle of attack

Lift

Coe

ffici

ent

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

Dra

g, P

itchi

ng M

omen

t

GA(W)-1 Clean GA(W)-1 35 deg deflection Extrapolated at 20 deg Drag, no Flaps Pitching Moment, No flaps

Figure 32: Characteristics of the selected wing



5.4. Wing selection and design First of all, the selection of the wing profile among the general aviation NASA profile, had to satisfy the low speed requirements, and also accommodate the hydrogen tanks. At low speed, both profiles have very similar performances, but the GA(W)-2 profile has a slightly lower drag. However, attention must be paid when comparing the two profiles based on the lift and drag characteristics, because the curves presented here have been derived from NASA papers, McGhee, Robert J. 1973, McGhee, Robert J. 1977 are plotted for slightly different Reynolds numbers. (3.9 106 for the GA(W)-1 profile and 4.1 106 for GA(W)-2 profile). For the flaps configurations, the same Reynolds Number applies (2.2 106) but different deflections are investigated in Wentz, W.H (1977, 1983). As presented in the power plant chapter, hydrogen storage is quite inefficient compared to gasoline storage, and requires a greater amount of space. A layer of insulating material is required combined with a structure able to withstand a differential pressure of 20 psi, twice the usual working pressure on airliners’ fuselage. The shape of the tanks needs to be cylindrical or at least made out of several circles to withstand pressure. The GA(W)-1 profile, thicker, offers a greater space to store the tanks and hydrogen. According to Howe, D. 2000, the optimum combination for minimum mass is a 26 % thick wing with an aspect ratio of 7.56, but for thinner wings the aspect ratio prescribed is 6.42 for a GA(W)-1 profile and 6.02 for a GA(W)-2 version. Unfortunately, small aspect ratios are more typical of aerobatic aircraft, than reconnaissance aeroplanes. Small aspect ratios enable higher lift coefficients, but correspond to small wing areas which do not allow low speed flights.

GA(W)-1,

Cruise

GA(W)-2

Cruise

GA(W)-1,

Fowler

Flaps 35º

deflection

GA(W)-1,

Fowler

Flaps, 20º

deflection

GA(W)-2 Single Slotted

Flaps, 40º deflection

GA(W)-2 Fowler Flaps,

20º deflection

GA(W)-2 Fowler Flaps,

40º deflection

GA(W)-1, Fowler Flaps,

20º deflection

Aspect Ratio 6.42 6.02 6.42 6.42 6.02 6.02 6.02 9Wing Area 19.62 19.66 19.62 19.62 19.66 19.66 19.66 27.1Lift Coefficient 0.434 0.467 2.669 1.797 2.454 1.459 2.819 1.889Aoa 1 1 10 10 13 12 10Speed m/s 52.7 52.7 21.97 26.77 22.89 29.68 21.36 22.22Speed Km/h 189.7 189.7 79.1 96.4 82.4 106.9 76.9 80.0Drag coefficient 0.0293 0.0243 0.839 0.436 0.769 0.763 0.93 0.335Drag 1095.2 972.6 3675.3 2882.6 3669.5 6124.8 3850.5 2102.9Trottle Setting 0.43 0.38 0.60 0.58 0.63 1.36 0.61 0.35

10

Table 15 : Summary of the performances for different configurations. As shown in this table, the GA(W)-2 wing has slightly lower drag on the geometry prescribed by Howe, 2000. The row corresponding to speed indicates the cruise speed for the two first configurations, and the stall speed for the flaps configurations. The purpose of this table is to show the reader the ranges of speeds and thrust requirements for each configuration.



It can be noticed that with a deflection of 20 º, none of the profiles enables the loitering capability the aircraft is designed for. Moreover the GA(W)-2 profile creates too much drag at 106.9 m/s, to be compensated by the fuel cell engine. Under the initial geometry, both profiles allow to fly at low speed but with a full flap deflection which leads to a throttle setting around 60%. Designing the aircraft for loitering with full flap deflection, would seriously compromise the range by increasing the fuel consumption. Eventually, it was chosen by the designer to come back to higher aspect ratio, as most of reconnaissance aircraft, and to allow for a bigger wing area (27.1 m2). By changing the geometry of the wing, the designer makes a compromise between low mass and low speed capability.

6. Conclusion The initial calculations presented above gave early results which were used as guidance for the CAD modelling which follows. Some of the dimensions have been modified later in the design process, but calculations and CAD model were kept up to date. Aircraft design is made of compromises between low mass, performances and costs. Even if the optimum solution for a minimum weight requires a thick (26 %) wing, it has been decided that it would be a better compromise to reduce drag by selecting a thinner wing and reduce costs and risks by using a standard profile. The similar mission of our fuel cell aircraft and of the Optica, lead to the same choice of wing: a 17 % thick GA(W)-1 profile with 30% chord Fowler flaps, and an aspect ratio of 9.


Fuel-cell powered aircraft Design Development

Design development

1. General Arrangement

Figure 33 : General arrangement of the reconnaissance electric aircraft

The general arrangement of the aircraft has been designed to achieve a maximum visibility cockpit, combined with a standard wing profile and conventional tail unit, to provide an easy handling. The model was built to define the internal arrangement of the fuselage, and help to CG calculations which are presented later on. Internal cabin has been designed to accommodate all sizes of passengers, from typical 5% Japanese female to 95% American male.

2. Wings The wings have been designed based on the geometry defined in previous chapter, with struts to reduce structural weight. The main structure is made of two spars located at 15 % chord and 60 % chord, allowing space for 30 % fowler flap and ailerons. The hydrogen tanks are located between the spars.



Figure 34 : Wing concept, showing front and rear spar.

3. Fuselage The fuselage has been designed to accommodate the power plant and the high visibility cabin. 9 meters long, its maximum width is 1.7 m. Two versions have been designed. The first version included a low tail, which was the source of ground clearance issues in combined roll and pitch situations. A second sketch has then been drawn, accounting for these issues. Such an obvious issue would have been addressed on the first draw by an experienced designer, but in this manner the author learnt more on the design process.



Figure 35: Fuselage Left View

4. Tail Unit The tail unit has fist been designed as a conventional unit, based on the areas given by empirical formulas. Root and tip chords, span have been decided arbitrary to provide enough control area, so that root chord was always smaller than the wing chord. A T-tail would have been better for ground clearance considerations, but worst on a structural aspect and downwash stream would have imposed greater areas.

Figure 36 : Tail Unit

Fin has been designed based on a symmetrical NACA profile, and scaled up to match the dimension requirements.



Figure 36 shows the design development of the tail. The left image corresponds to wing geometry with a lower aspect ratio, and un-swept surfaces. However, increasing the wing area to enable the low speed capability resulted in larger stabiliser areas. In order to maintain ground clearances and reduce the areas, swept surfaces have been designed. Swept surfaces allow a shift in the centre of pressure backwards, thus increasing the arm so that smaller surfaces are required. Eventually, the root chord of the fin has been increased significantly so that a smaller span was required.

5. Nacelles Nacelles have been designed to accommodate small electric motors. Unfortunately, the precise dimensions of the selected motors were not available, although the manufacturer has been contacted several times. The nacelles have been sized based on other manufacturers dimension and can accommodate motors up to 400x200x400 approximately. The nacelle should also provide space for a cooling device.

Figure 37 : Nacelles for electric motors

In order to provide maximum lift, nacelles have been placed under the wing. According to CS-23, sufficient clearance has been left between the propeller and the fuselage structure, as well as with the wing. Mounted on a high wing, there are no ground clearance issues.



6. Power plant Power plant has been designed in a very simple way, in order to model the general arrangement of the fuselage interior. The dimension for the fuel cell stacks, inverter, cooling systems and compressor were available so that a reasonable space representation could be set up.

Figure 38 : Power plant location Global positioning of the power plant was decided by CG considerations, which are explained further on this report. The stacks are the heavier component and have been organised into two layers. The power plant is quiet a long way aft in order to balance the large offset and high weight of the cabin and payload.

7. Cabin The cabin has been designed to accommodate up to 4 passengers along with their luggage. The seats are arranged in two rows of two. In order to achieve a maximum visibility, the accommodation for passengers is located as forward as possible. The windshield had been designed for reconnaissance and observation missions. The design is inspired from helicopters canopies. Large windows allow observation in large angles of view and floor windows for the crew allow to spot targets under the aircraft.



Figure 39 : Cabin isometric and left views Basic arrangement of the cabin has been checked in CATIA v5 with the human activity tool. This tool allows to check movements of the crew as well as to give a representation of the view the passengers would have. Figure 40 is a digital mock-up of the pilot line of sight when looking through the windshield.

Figure 40 : Pilot line of sight This design only gives an idea of the visibility of the cockpit, and windshield transparent area could be reduced for structural considerations, however based on helicopters geometry, the visibility achieved is competitive.



8. Landing gears and ground clearances Landing gears have been designed based on the methodology provided in Young, 2004. The landing gear is a conventional tri-cycle configuration with a nose wheel, as for competitor aircraft the landing gear is fixed. More drag is produced, but lower mass is achieved. Moreover, sufficient space in the forward fuselage was not available to store a retractable landing gear. The design assumes that the wheels are the same as baseline aircraft Optica and Cessna 172, and have a single wheel per unit. The nose wheel is a 5x5.00 of external diameter 330 mm. The main landing gear is a 6x6.00 wheel of external diameter 386 mm. The main gear has been located at 60 % chord, and the vertical and horizontal positions have been chosen to achieve sufficient ground clearance.

Figure 41 : Front and Left view of the landing gear

In the take-off or landing phases, strong cross-wings or gusts can happen, causing a sudden change of attitude of the aircraft. In order to keep the integrity of the structure in these cases and avoid collisions, ground clearances were checked. The clearance requirements are as follow:

15º in pitch, allowing for rotation in take-off and landing attitudes 20 º in roll, allowing for cross wing effects, such a high value is recommended

for light aircraft. Combined effects, 15 º pitch and 20 º roll.



Figure 42 : Pitch and Roll ground clearances

Figure 43: Combined Ground Clearance

As explained above, the fuselage had to be re-designed to achieve sufficient ground clearances, and the resulting fuselage is quite high above the ground as shown figure 43. Therefore the cabin door may have to accommodate stairs.

9. Hydrogen tanks As explained in chapter 3 covering the power plant architecture, hydrogen has to be stored in wings to carry sufficient volume.



Figure 44 : detailed hydrogen storage concept

As the hydrogen is stored at relatively high pressure, (20 psi) the wing structure would have to resist the pressure loads. To avoid a significant weight penalty, the wing tanks have been re-designed to take advantage of circle arcs. Cylinder is the most weight efficient way to resist pressure loads. However in order to maintain a significant insulating material thickness, a great amount of volume is lost, and more material is used leading to further weight penalties. A simple hoop stress calculation showed that, assuming that the insulating material is aluminium, the minimum tank thickness would be less that 0.2 mm thanks to the very small radius on which the pressure applies. The full span equipped with hydrogen tanks is designed to accommodate up to 1.87 m3, which corresponds to 132 kg of hydrogen. Two options are possible for maintenance of the hydrogen tanks. One option would be to design a integral fuel tank. The skin (upper or lower) would be removable, and the tank could be removed or replaced in one single piece. The tank would be made of components limited to the ribs d the structure. The other option would be to have a tank made out of two parts, (upper and lower) to allow maintenance under a removable skin.


Fuel-cell powered aircraft Design Analysis

Design analysis

1. Mass and CG management A major consequence of designing a reconnaissance aircraft with such a forward located cabin was a difficult CG management. In a full payload configuration, the centre of gravity would be shifted a long way forward, due to the position of the cabin. In a empty configuration, with a very light pilot alone, the cg would have been moved backwards by the fuel cell stack important weight. To solve this issue, the two critical design cases have been identified, and used to locate precisely the power plant location within the fuselage, and the baggage compartment.

Critical forward case: Full payload, no luggage, no hydrogen Critical aft case: Light pilot only, full luggage, full hydrogen

In order to keep the aircraft under control at any time, it is essential to have a CG position which does not vary much and which is located between 15 and 35 % chord, to be as close as possible from the wing centre of pressure. The mass and CG balance has been performed by collecting the CG location components from the CAD model and accounting for each component weight (from the optimisation spreadsheet). Table 16 below summarise the CG calculation for the critical forward case.

Element Mass Wing apex 2370x y z x y z Chord 1540

Wings 84 3023 21 1012 253932 1764 85008Fuselage 161 3165 0 219 510609 0 35331Cabin Payload 495 1744 160 -7 863280 79200 -3465Landing gears 61 2084 -63 -868 126374 -3820.32 -52636 FC position 3537.257Empennage 20 7965 0 737 159300 0 14740 Baggage position 3000Fuel cell stacks 136 3537.257 0 -60 481067 0 -8160Inverter 17 3712.257 0 57 62366 0 958Cooling system 17 3252.257 -15 94 55288 -255 1598Supercharger 40 3677.257 96 258 147090 3840 10320Electric motors and propellers 36 3213 0 1006 114062 0 35713Baggage compartment 0 3000 0 0 0 0 0Hydrogen 0 2947 0 942 0 0 0

MASS 1066 CG Position 2601 76 112 15.000 % Chord

CG from CATIA Mass X distance

Table 16 : Foward Cg calculation case

All distances are taken from the nose of the aircraft. In order to achieve CG control, the location of the power plant has been selected to achieve a forward cg of 15% chord.



The same calculations have been performed for the rear case, but as the power plant location was already set, therefore the CG was adjusted by the location of the luggage. It has been found that, in the rear critical case, in order to balance the rear located and heavy fuel cell system and fix the CG at 35% chord, the luggage had to be positioned at 3m from the nose apex. Practically, this means that luggage is placed on the rear seat. If the baggage location remains the same, the CG moves back to 38 % chord, which is still acceptable but limiting the CG shifts to 15-35% chord is strongly advised. Centre of gravity displacements during flight has also been investigated and results are plotted below:

CG displacements

255026002650270027502800285029002950

0 20 40 60Hydrogen Weight (kg)

CG

long

itudi

nal p

ositi

o

80

n(m

m)

MAX PAYLOAD PILOT ONLY 35 % chord 15% chord

Figure 45 : CG diplacements as function of remaining fuel

The design of a fuel cell powered aircraft leads to difficult CG management because of the high weight of the stacks, which ideally would be located near the quarter chord. However, as the reconnaissance aircraft features a cabin located far forward, the stack is used to trim the aircraft. The displacements due to hydrogen consumption are small because the wing is unswept and the total fuel weight is relatively small compared to piston engine aircraft.

2. Drag estimation For a better estimation of the performances of the aircraft, and especially the stall and loitering speed, a better approximation of the drag was necessary. The analysis was conducted on the methodology introduced in Stinton, (2001), combined with the information found on the wing profile.



The methodology accounts for each component of the aircraft (fuselage, struts, landing gears, tail, and nacelles). For each component the parasitic drag is obtained by multiplying the cross sectional area by an empirical coefficient and the dynamic pressure.

ACq

DpD ×= π

q is the dynamic pressure, Dp the parasitic drag (N), and Cdπ the empirical coefficient. The area A is the frontal area, i.e. the area of the projection of the object on the plan perpendicular to the longitudinal axis of the aircraft. However for control surfaces, flaps, A is the area of the surface itself. For the landing gear, the parasitic drag is obtained by multiplying the sum of all Dp/q factors by a constant.

∑∑ ××=×=components

Dcomponents

p ACq

Dq

Dpπ25.125.1

This equation corresponds in our case at a drag coefficient of 0.06, so that the total drag coefficient is 0.3133.

Component Cd Drag Wing 0.016226 397.55 49.80 %

Fuselage 0.11574 146.10 18.30 % Tail Unit 0.108927 137.50 17.22 %

Struts 0.01992 25.14 3.15 % Nacelles 0.00612 7.73 0.97 %

Landing Gears 0.066733 84.24 10.55 %

Table 17 : Drag distribution between components in cruise configuration

3. Performances Once that the drag of the aircraft had been analysed, it was possible to check if the low speed reconnaissance capability was fulfilled. The analysis has been partially presented in the wing selection chapter, Table 15. The graph below presents the different drag characteristics of the aircraft as a function of speed, in different configurations



Drag and Thrust

-1000

1000

3000

5000

7000

9000

11000

13000

15000

0 10 20 30 40 50 60 70 80

Speed (m/s)

Dra

g (N

)

Cruise, No flaps, Aoa 0

Cruise, No flaps, Aoa 5

Cruise, No Flaps, Aoa 10

Loitering, Flaps 20, Aoa 0



Take off, Full Flaps, Aoa 5

Take off, Full Flaps, Aoa 10

Available Power

Figure 46 : Drag and Thrust requirements for different configurations

In this plot the cruise and loitering configurations take place at an altitude of 10,000 ft, whereas the take-off configuration is obviously designed for sea level. It should be noticed that at lower altitudes, loitering requires more power than showed in this graph. The power curve represents the available power at the electric motor, and therefore does not allow for the propeller efficiency.

4. Lift Drag ratio The lift drag ratio has been estimated for cruise and loitering configurations, and following values have been found:

MaxCruise DL

DL

⎟⎠⎞

⎜⎝⎛==⎟

⎠⎞

⎜⎝⎛ 17.14

44.5=⎟⎠⎞

⎜⎝⎛

LoiteringDL

5. Propellers efficiency According to Stinton, (2001) the efficiency of the propellers is:



839.0)/(1

85.02

max

=××+

=dq

DLW

pη

6. Range The calculation of the aircraft range has been performed using the Breget equation. Hopefully this equation does not assume any fuel type. Using the hydrogen specific consumption calculated earlier in the design process, the range could be estimated.

⎟⎟⎠

⎞⎜⎜⎝

⎛⎟⎠⎞

⎜⎝⎛×⎟⎟

⎠

⎞⎜⎜⎝

⎛=

E

p

WW

DL

cR log

'750

max

η

Unfortunately assuming a hydrogen quantity of 70 kg as previously estimated, and the maximum Lift Drag ratio, the maximum range is 379 nm (735 km). However this value is close to the initial objective of 431 nm (800 km). In order to achieve the target range, a quantity of hydrogen of 77 kg should be carried. This capacity would require a hydrogen tank over 9m span, for a tank weight of 21.3 kg based on the latest geometry presented in the design development chapter. The designer also looked at an extended range design, in which the full span of the aircraft would be filled. The maximum quantity which can be carried is 132 kg of hydrogen. This capacity corresponds to a fuel tank of 36.8 kg. Such a configuration could lead to a range of 709 nm (1,313 km). However the take-off mass would be significantly increased, which means lower performances, longer take-off distance, higher stall speed, and therefore higher loitering speed. This configuration corresponds to a private use of the aircraft, for leisure piloting, the second mission of the aircraft. Thanks to the low density of hydrogen, if the aircraft is operated with only 3 passengers, the design take-off mass of 1171 kg is not exceeded and target performances can be met with the extended range configuration.

7. Summary of performances The fuel cell aircraft reaches its target performance of low speed characteristics, but misses from a few miles the target range. Moreover, the Breguet range equation assumes a full flight at maximum Lift Drag ratio, which is clearly not practical as soon as the crew decides to fly at low speed. However, there is clearly an improvement potential in the fuselage design. The fuselage is very large compared to competitors aircraft, (1.7m) and re-designing a narrow fuselage could reduce significantly the cross sectional area and thus reduce drag of the



aircraft. Lift Drag ratio would be improved and again proper range performances could be achieved.

8. Comparison to competitors

Manufacturer Aviaimpex JSC

Robinson Helicopter

Vertical Aviation

Technologies Cranfield Edgley Cessna

Model KT-112 Yanhol R 44 HUMMINGBIRD Fuel cell

aircraft Optica 172 Skyhawk

Span 8.06 10.06 10.06 15.62 12 11 Max T/O Weight 900 1134 1225 1171 1350 1111 Rated Power 147 224 186 134 193 119 DOC 70 (2004) 43.3 (2004) - - - - Cost 150000 307000 170000 - - 150000 Range 420 643 603 735 (-) 619 1070 Speed 161 217 145 190 191 226

Table 18 : Comparison of Fuel Cell aircraft to competitors

Table 18 summarises the comparison between competitor helicopters and fuel cell powered aircraft, along with the two design baselines. The fuel cell aircraft has a much wider span than helicopters and baseline aircraft because of the low speed capability which imposed large wing area.

Figure 47 : Competitors helicopters (left to right : Hummingbird, R44, Yanhol) .



The take-off mass is within the range of competitors, although Optica and Cessna are quite old designs. The fuel cell powered aircraft requires less power than equivalent helicopters. However the capability of a low speed aircraft and helicopters are different, and are designed for slightly different missions. It should be noticed that the direct operating costs are figures from 2004 and do not account for the steady increase in oil price since 2004, (figure 1) and in near future (figure 2). Moreover direct operating costs of rotor wing aircraft are usually higher than fixed wings, so that a fuel cell powered aircraft would be economically advantageous around 2010-2015. In terms of range, the fuel cell aircraft allows a longer range, but loitering at low speed significantly reduces the autonomy. In the extended range configuration, performance is comparable to baseline aircraft. For pure reconnaissance missions, the aircraft designed is competitive for cockpit visibility, range and direct operating costs.


Fuel-cell powered aircraft Discussion

Discussion During the whole duration of this project important information or figure were missing, and assumptions had to be made. Fuel cell is a recent technology, and many patents are protecting existing techniques or knowledge. Moreover automotive industry has invested very large amounts of money in fuel cell research. Consequently information on fuel cell systems is not always available, and latest performances for example are not available to the public. Assumptions had to be made on essential figures, such as fuel-cell efficiency, however assumptions were always justified with available literature. The choice of fuel cell type and fuel has been concluded from a performance review of each possibility. Proton exchange membrane is currently the more mature technology, and is generally selected for all mobile modern applications. Fortunately this is also the technology for which the data is relatively more easily available. It must be kept in mind that direct operating costs are based on forecasts of hydrogen prices, which depends very much on public investment. Public investment in fuel cell technology will depend on government energetic policy and public opinion. The figures quoted therefore are only forecasts, and are subject to revision in next years. In the same manner, pollution resulting in fuel cell powered aviation will depend on the production method and distribution scheme of hydrogen which is adopted. These parameters of the analysis presented in this thesis are again influenced by policy and public opinion. It is most likely that the first production methods to be implemented at large scale will be the cheaper and more polluting one such as oil reforming. Producing hydrogen from oil or natural gas can only be a temporary solution, as the dependence on limited fossil fuel resources will be maintained. The path to hydrogen economy will probably be opened by the introduction of fuel cell powered cars and buses. Increasing use of these vehicles will create a need for production infrastructure and develop distribution schemes, so that when hydrogen aviation may be ready, the infrastructure will exist thanks to automotive market. Some hydrogen stations already exist and operate a fuel cell powered fleet of buses in major cities. Such projects will, hopefully, greatly help to demonstrate the viability of hydrogen infrastructure. All the government have acknowledged that solution to oil dependence will be through hydrogen economy, and some have already set roadmaps to enable the transition presented above, but in all cases hydrogen technology is seriously investigated. Many concerns may be formulated on the choice for a liquid hydrogen storage system. However, liquid hydrogen storage is not a new technology. Liquid hydrogen has been used in rocket and launcher for decades. The space shuttle launcher currently uses liquid hydrogen tanks with a solid foam insulation layer, to protect from aerodynamic heating. Hydrogen suffers from the bad reputation associated with the Hindenburg accident in 1937, but many references have demonstrated that hydrogen is not as dangerous as methane or propane, gases currently used on domestic applications.


Fuel-cell powered aircraft Discussion

Liquid hydrogen storage onboard creates many technological challenges on materials, handling methods, or maintenance issues. The design of hydrogen wing tank will impose serious material constraints for thermal expansion, thermal stresses, toughness, fatigue and strength at cryogenic temperatures. These issues may be handled by the use of new aluminium alloys but again little information is available on effect of repeated thermal cycles on materials properties. Using liquid hydrogen as fuel will have important impact on refuelling procedures and fuelling infrastructures in airfields. To avoid mixture of gaseous hydrogen with oxygen, inert gas such as nitrogen will have to be widely available in every airfield. Maintenance procedure will also have to be modified to prevent any risk and also prevent leakage. Under the assumptions made, a power plant delivering 134 kW of power has been designed, for a power density of 0.59 kW/kg. According to Wentz (2005) the required power density to achieve equivalent performances to gasoline power plant on piston engine aircraft was 0.625 kW/kg. The target is almost achieved, which explains why the design range has not been achieved with the fuel cell engine. The resulting design of the reconnaissance electric aircraft is hybrid between fixed and rotor wing aircraft. It combines a cabin design inspired from helicopters cockpits, and a power plant located in the fuselage with a wing typical of low speed applications. After many configuration investigated, the wing profile is the same as the Optica, (NASA GA(W)-1) with a larger area (27.1 m2 versus 16 m in the original design). However new materials will alloy to reduce weight compared to design from the 80’s. Consequently the wing loading is much lower, and the specific power required is 30 % lower. Compared to the Cessna baseline, the opposite comparison applies, the wing area is larger to achieve low speed capability, and wing loading is lower. However the specific thrust is smaller. As presented in Berton et al (2003), the range is lower than initially defined and a small reduction in payload is necessary to achieve identical performances. This is due to the low density of hydrogen which does not allow storing much energy in wing tanks compared to gasoline baseline.


Fuel-cell powered aircraft Recommendations

Recommendations In order to solve the data collection issue, the author would suggest, as Bekiaris (2002) did, to set up agreements with electric equipment manufacturers interested in aerospace applications. Fuel cell information was limited, and even if several companies have been contacted none answered positively. Defining an agreement which would probably include confidentiality considerations, would allow the next study to be based on consistent data, and therefore avoid multiplication of sources, assumptions or extrapolated figures. Moreover a great amount of time may be saved and unsuccessful efforts to find precise data maybe spared. The author would also suggest that the design of a liquid hydrogen tank for aviation purposes could be the topic of a whole Msc thesis. There are interesting conclusions to be drawn from already existing technologies, especially in launchers application, which could be assessed by an Msc student. Issues concerning material properties at cryogenic temperatures have been partially addressed in this paper, but again investigations were limited by timeframe of the project and the literature review could have been more developed on this particular point as studies have been carried out since the 1950’s. A risk and failure analysis could also be carried out once the design is frozen. The power plant designed in this work is based on existing technology and a significant number of the components used have been designed for automotive applications, for which weight is not critical. A more detailed analysis and weight saving analysis could be carried out. A more detailed analysis of the fuel cell performances could also be performed, but this work would require more chemical and electronic skills than aerospace design skills.


Fuel-cell powered aircraft Conclusions

Conclusion It is widely accepted that moving towards hydrogen economy is the only viable option to address the energetic dependence to fossil fuels issues. However to solve the pollution problem and limit greenhouse gases emissions, hydrogen has to be produced from renewable energies which would initially lead to infrastructure problems. It has been demonstrated that in the short term, hydrogen engine aircraft may lead to lower direct operating costs compared to gasoline engines. A fuel-cell aircraft has been designed to achieve reconnaissance and surveillance missions, and alternatively be used for private piloting. A power plant matching the requirements has been designed accounting for cooling and oxygen feeding devices. Eventually a specific power of 0.59 kW/kg has been achieved. Liquid hydrogen storage concept was developed and material compatibility issues were addressed, and heat transfers accounted for. It was demonstrated that it was possible to design a light hydrogen tank which would reduce the boil-off loses to less than 1 % of capacity over a 4.5 hour journey. Storing energy as liquid hydrogen onboard will have important impact on safety measures, fuelling and de-fuelling or maintenance procedures for the aircraft, but the system is still viable. A reconnaissance aircraft has been designed to accommodate a fuel cell engine power plant based on current technology, and as expected, performances are almost equivalent to gasoline competitors, as the fuel cell specific power tends to reach the level of maturity required. However, application of fuel cell to airliners is still a long term prospect, as the technology is far from being competitive with turbo-prop or gas-turbine engines. Development of production and distribution infrastructures and safety procedures will probably be accelerated by the introduction of hydrogen automotive products. For example fuelling stations are already being installed to fuel a fleet of electric buses. Hopefully, the infrastructure will be ready to provide hydrogen for aviation applications by the time the technology will achieve sufficient power density for larger applications.


Fuel-cell powered aircraft References and Tables

References

ADAMSON, K.A. and CRAWLEY, G., 2006. Fuel Cell Today Market Survey: Light Duty Vehicles. fuelcelltoday.com: Fuel cell today.

ALEXANDER, D.S., 2005. Advanced Energetics for Aeronautical Applications: Volume II. NASA/CR-2005-213749. NASA.

ALEXANDER, D.S., 2003. Advanced Energetics for Aeronautical Applications. NASA/CR-2003-212169.

ALEXANDER, D., LEE, Y.M., GUYNN, M. and BUSHNELL, D., 2002. EMISSIONLESS AIRCRAFT STUDY. AIAA 2002-4056. American Institute of Aeronautics and Astronautics.

ANANTHACHAR, V. and DUFFY, J.D., 2005. Efficiencies of hydrogen storage systems onboard fuel cell vehicles.

ARUN K. SEHRA, WOODROW WHITLOW JR., 2004. Propulsion and power for 21st century aviation. Progress in Aerospace Sciences 40 (2004) 199–235.

BEKIARIS, N., 2002. Conceptual Design of a Fuel Cell powered Aircraft, Cranfield University.

BLOOM, I. and POLZIN, E., 2005. FY 2005 Progress Report, US DOE Hydrogen Program. US Department of Energy.

BOSSEL, U.G., 2003. Efficiency of Hydrogen Fuel Cell, Diesel-SOFC-Hybrid and Battery Electric Vehicles, European Fuel cell forum, 2003.

BOSSEL, U.G., 1999. Solid Oxide Fuel Cells for Transportation, 3rd European SOFC Forum, 1999.

BOUQUET, F. and SANDERS, H.M., 2003. SPACE APPLICATIONS OF HYDROGEN STORAGE IN CARBON NANOSTRUCTURES. AIAA 2003-4733. American Institute of Aeronautics and Astronautics.

COLOZZA, A.J., 2002. Hydrogen Storage for Aircraft Applications Overview. NASA/CR—2002-211867. NASA.



DAGGETT, D.L., EELMAN, S. and KRISTIANSSON, C., 2003. FUEL CELL APU FOR COMMERCIAL AIRCRAFT. AIAA 2003-2660. American Institute of Aeronautics and Astronautics.

FIELDING, J., 1999. Introduction to aircraft design. Cambridge: Cambridge University Press.

FRANGOPOULOS, C.A. and NAKOS, L.G., 2006. Development of a model for thermo-economic design and operation optimization of a PEM fuel cell system.

FRIEND, M.G. and DAGGETT, D.L., 2003. FUEL CELL DEMONSTRATOR AIRPLANE. AIAA 2003-2868. American Institute of Aeronautics and Astronautics.

FUJII, H. and ICHIKAWA, T., 2006. Recent development on hydrogen storage materials composed of light elements.

HERWERTH, C., OFOMA, U., WU, C., MATSUYAMA, S. and CLARK, S., 2006. Development of a Fuel Cell Powered UAV for Environmental Research. AIAA 2006-237. American Institute of Aeronautics and Astronautics.

HIIBNER, W., GRADT, T., SCHNEIDER, T. and BORNER, B., 1998. Tribological behavior of materials at cryogenic temperatures.

HOWE, D., 2000. Aircraft conceptual design synthesis. London: Professional Engineering Publishing.

J. W. YOUNGBLOOD, T. A. TALAY, AND R. J. PEGG, 1984. Design of Long-Endurance Unmanned Airplanes Incorporating Solar and Fuel-Cell Propulsion. 84-1430.

JEFFREY J. BERTON, JOSHUA E. FREEH, AND TIMOTHY J. WICKENHEISER, 2003. An Analytical Performance Assessment of a Fuel Cell-Powered, Small Electric Airplane. NASA/TM—2003-212393. NASA.

K. HARALDSSON, P. ALVFORS, 2005. Effects of ambient conditions on fuel cell vehicle performance. Journal of Power Sources 145 (2005) 298–306.

KOHOUT, L.L. and SCHMITZ, P.C., 2003. Fuel Cell Propulsion Systems for an All-Electric Personal Air Vehicle. NASA/TM—2003-212354. NASA.

LARMINIE, J. and DICKS, A., 2003. Fuel Cell Systems explained Second Edition. Chichester: Whiley.



MARK D. GUYNN, JOSHUA E. FREEH, ERIK D. OLSON, 2004. Evaluation of a Hydrogen Fuel Cell Powered Blended-Wing-Body Aircraft Concept for Reduced Noise and Emissions. NASA 2004-212989. NASA.

MASSON, P.J. and LUONGO, C.A., 2005. High power density superconducting motor for all-electric aircraft propulsion. IEEE Transactions on Applied Superconductivity, 15(2 PART II), pp. 2226-2229.

MASSON, P.J., SOBAN, D.S., UPTON, E., PIENKOS, J.E. and LUONGO, C.A., 2005. HTS motors in aircraft propulsion: Design considerations. IEEE Transactions on Applied Superconductivity, 15(2 PART II), pp. 2218-2221.

MCCONNELL, B.W., 2005. Applications of High Temperature Super-conductors to Direct Current Electric Power Transmission and Distribution. IEEE TRANSACTIONS ON APPLIED SUPERCONDUCTIVITY.

MCGHEE, R.J., 1973. LOW-SPEED AERODYNAMIC CHARACTERISTICSOF A 17-PERCENT-THICK AIRFOIL SECTIONDESIGNED FOR GENERAL AVIATION APPLICATIONS. NASA TN D-7428. NASA.

MCGHEE, R.J., BEASLEY, W.D. and SOMERS, D., 1977. LOW-SPEED AERODYNAMIC CHARACTERISTICSOF A 13 PER CENT THICK AIRFOILSECTION DESIGNED FOR GENERAL AVIATIONAPPLICATIONS. NASA TMX-72697. NASA.

MOFFITT, B.A., BRADLEY, T.H., PAREKH, D.E. and MAVRIS, D., 2006. Design and Performance Validation of a Fuel Cell Unmanned Aerial Vehicle. AIAA 2006-823. American Institute of Aeronautics and Astronautics.

MOORE, R.B. and RAMAN, V., 1998. HYDROGEN INFRASTRUCTURE FOR FUEL CELL TRANSPORTATION. J. Hydrogen Energy.

MURAKAMI, Y. and UCHIBORI, K., 2006. Development of Fuel Cell Vehicle with Next-generation Fuel Cell Stack. 2006-01-0034. Society of Automotive Engineers.

NAM, T., SOBANY, D.S. and MAVRIS, D., 2005. A Generalized Aircraft Sizing Method and Application to Electric Aircraft. AIAA 2005-5574. American Institute of Aeronautics and Astronautics.

NETTLES, A. and BISS, E.J., 1996. Low Temperature Mechanical Testing of Carbon-Fiber/Epoxy-Resin Composite Materials. 3663. NASA.

O'HAYRE, R.P., ed, 2006. Fuel cell fundamentals. John Wiley & Sons.



PENNER, J.E., LISTER, D.H., GRIGGS, D.J., DOKKEN, D.J. and MCFARLAND, M., 2006. Aviation and the Global Atmosphere. Intergovernmental Panel on Climate Change.

PRATT, J.W., BROUWER, J. and SAMUELSEN, G.S., 2005. Experimental Performance of an Air-Breathing PEM Fuel Cell at High Altitude Conditions, 43rd AIAA Aerospace Sciences Meeting and Exhibit; Reno, NV; USA; 10-13 Ja. 2005, 2005, American Institute of Aeronautics and Astronautics.

REYNOLDS, T.W., 1955. Aircraft Fuel tank for liquid hydrogen. NACA RM E55F22. National Advisory Committee for aeronautics.

SAXE, M. and ALVFORS, P., 2005. Advantages of integration with industry for electrolytic hydrogen production.

SOLOMON, D.B. and BANERJEE, A., 2004. A global survey of hydrogen energy research, development and policy.

SORENSEN, B., 2005. Hydrogen and Fuel Cells. Elsevier Academic Press.

STINTON, D., 2001. The design of the aeroplane . Oxford: Blackwell Science.

STINTON, D., 2001. The design of the aeroplane . Oxford: Blackwell Science.

STINTON, D., 1998. The anatomy of the airplane. Second Edition edn. Oxford: Blackwell Science.

TAEWOO, N., SOBANY , D. and MAVRIS, D., 2005. A Generalized Aircraft Sizing Method and Application to Electric Aircraft. AIAA 2005-5574. American Institute of Aeronautics and Astronautics.

TETSUMI, Y., OGATA, T., SAITO, M. and HIRAYAMA, Y., 2001. Effect of welding structure on high-cycle and low-cycle fatigue properties for MIG welded a 5083 aluminum alloy at cryogenic temperatures.

THRING, R.H., 2004. Fuel Cells for Automotive applications. Towbridge UK: Cromwell Press.

TSENGA, P., LEEB, J. and FRILEY, P., 2005. A hydrogen economy: opportunities and challenges.

US DEPARTMENT OF ENERGY, 2006. Annual Energy Outlook 2006. US Department of Energy.



US DEPARTMENT OF ENERGY, 2005. DOE Hydrogen Program VII.I.7 Fuel Cell Testing at Argonne National Laboratory. US DOE.

US DEPARTMENT OF ENERGY, 2002. NATIONAL HYDROGENENERGY ROADMAP, PRODUCTION • DELIVERY • STORAGE • CONVERSION• APPLICATIONS • PUBLIC EDUCATION AND OUTREACH. USDOE.

VANDER KOOI, D.C., PARK, W. and HILTON, M.R., 1999. CHARACTERIZATION OF CRYOGENIC MECHANICALPROPERTIES OF ALUMINUM-LITHIUM ALLOY C-458.

VIELSTICH, V., LAMM, A. and GASTEIGER, H.A., 2003. Handbook of Fuel Cells, Vol 1-4. Chichester: Whiley.

WENTZ, W.H. and FISCKO, K.A., 1978. Pressure Distributions for the GA( W)-2 Airfoil With 20% Aileron, 25% Slotted Flap and 30% Fowler Flap. NASA CR 2948. NASA.

WENTZ, W.H., FISCKO, K.A. and SEETHARAM, H.C., 1977. FORCE AND PRESSURE TESTS OF THE GA(W)-1 AIRFOIL WITH A 20%AILERON AND PRESSURE TESTS WITH A 30% FOWLER FLAP. NASA CR 2833. NASA.

WENTZ, W.H., MYOSE, R.Y. and MOHAMED, S.A., 2005. Hydrogen-Fueled General Aviation Airplanes. AIAA 2005-7324. American Institute of Aeronautics and Astronautics.

WENTZ, W.H. and OSTOWARI, C., 1983. Additional Flow Field Studies of the GA(W)- 1 Airfoil With 30 -Percent Chord Fowler Flap Including Slot-Gap Variations and Cove Shape Modifications. NASA CR 3687. NASA.

YOUNG, D., 2004. Landing gear design lecture notes. Cranfield University.



List of Figures

Figure 1: Oil prices trends, 2004, (US Department of Energy 2006)............................... 2 Figure 2 : Recent evolution of the oil barrel price, (www.wikipedia.org) ....................... 2 Figure 3 : Aviation Carbon release in atmosphere, (Penner,J.E. 2006) ........................... 3 Figure 4 : Single Cell assembly (Larminie,J. 2003) ......................................................... 4 Figure 5 : Fuel cell stack (Larminie,J. 2003).................................................................... 5 Figure 6 : PEMFC principle (Larminie,J. 2003) .............................................................. 5 Figure 7 : SOFC operating principle, (Larminie,J. 2003) ................................................ 7 Figure 8 : BLDC motor operating principle ................................................................... 15 Figure 9 : Power density for HTS motor. (Masson, P.J. 2005) ...................................... 16 Figure 10: Hy-Gen Electric Vehicle, General Motors.................................................... 17 Figure 11 : Boeing demonstrator project....................................................................... 18 Figure 12: Edgley Optica................................................................................................ 21 Figure 13 : Cessna 172 ................................................................................................... 24 Figure 14: Empty Weight without Power plant.............................................................. 31 Figure 15: Aspect Ratio and Wing Span ........................................................................ 32 Figure 16: Power Loading and Wing Loading ............................................................... 33 Figure 17: Stall and cruise speeds function of engine rated power ................................ 33 Figure 18 : Fuel cell powered aircraft power plant architecture..................................... 39 Figure 19: Material Properties, from Colozza (2002) .................................................... 41 Figure 20 : Liquid Hydrogen storage concept ................................................................ 42 Figure 21 : Hydrogen storage weight and losses............................................................ 43 Figure 22 : Fatigue properties of 5000 series Aluminium alloys at cryogenic temperatures, From Tetsumi et al (2001) ....................................................................... 45 Figure 23 : Fracture toughness at cryogenic temperatures for Al-Li alloy, Vander Kooi et al (1999)...................................................................................................................... 46 Figure 24: Water pump for fuel cell stack cooling system............................................. 48 Figure 25 : M62 Eaton supercharger .............................................................................. 50 Figure 26 : (next page) specific thrust plotter against wing loading .............................. 55 Figure 27: Rated power and Power requirements .......................................................... 57 Figure 28: Corrected Drag Characteristic GA(W)-1 ...................................................... 59 Figure 29: Corrected Lift curve slope GA(W)-1 ............................................................ 59 Figure 30: Lift Curves slopes for different flap configurations...................................... 61 Figure 31: Drag plots of different configurations........................................................... 62 Figure 32: Characteristics of the selected wing.............................................................. 62 Figure 33 : General arrangement of the reconnaissance electric aircraft ....................... 65 Figure 34 : Wing concept, showing front and rear spar. ................................................ 66 Figure 35: Fuselage Left View ....................................................................................... 67 Figure 36 : Tail Unit ...................................................................................................... 67 Figure 37 : Nacelles for electric motors ......................................................................... 68 Figure 38 : Power plant location .................................................................................... 69 Figure 39 : Cabin isometric and left views..................................................................... 70 Figure 40 : Pilot line of sight .......................................................................................... 70



Figure 41 : Front and Left view of the landing gear....................................................... 71 Figure 42 : Pitch and Roll ground clearances................................................................. 72 Figure 43: Combined Ground Clearance........................................................................ 72 Figure 44 : detailed hydrogen storage concept............................................................... 73 Figure 45 : CG diplacements as function of remaining fuel........................................... 75 Figure 46 : Drag and Thrust requirements for different configurations ......................... 77 Figure 47 : Competitors helicopters (left to right : Hummingbird, R44, Yanhol) . ....... 79



List of Tables

Table 1 : Hydrogen, Methane, Propane properties, Larminie,2003 ............................... 13 Table 2 : HTS motor for Cessna 172 application (Masson, P.J. 2005) .......................... 15 Table 3 : Power requirements, (Wentz, W.H 2005) ....................................................... 19 Table 4: Optica's main performance features ................................................................. 21 Table 5 : Possible conversion of Optica to Fuel Cell .................................................... 23 Table 6: Cessna 172 Performances................................................................................. 24 Table 7 : Conversion of Cessna 172 to Fuel Cell ........................................................... 26 Table 8 : Hydrogen Price, (Colozza, A.J. 2002)............................................................. 28 Table 9 : Target performances of a reconnaissance electric aircraft .............................. 34 Table 10: Decision table for concept choice .................................................................. 37 Table 11: Summary of supercharger types and performances ....................................... 49 Table 12 : Power plant mass summary........................................................................... 51 Table 13: Mass distribution of optimised aircraft .......................................................... 55 Table 14: Main dimensions of optimised aircraft........................................................... 55 Table 15 : Summary of the performances for different configurations.......................... 63 Table 16 : Foward Cg calculation case........................................................................... 74 Table 17 : Drag distribution between components in cruise configuration.................... 76 Table 18 : Comparison of Fuel Cell aircraft to competitors........................................... 79


Fuel-cell powered aircraft Appendices

List of Appendices

Appendix A Conceptual hand drawings Appendix B Fuel cell aircraft engineering drawing Appendix C Power plant components datasheets Appendix D Light aircraft and reconnaissance aircraft study Appendix E Modified Optimisation Spreadsheet


APPENDIX A

CONCEPTUAL HAND DRAWINGS

APPENDIX B

FUEL CELL AIRCRAFT ENGINEERING DRAWING

APPENDIX C

POWER PLANT COMPONENTS DATASHEETS

Specifications and descriptions in this document were in effect at the time of publication. Ballard Power Systems Inc. reserves the right to change specifications, product appearance or to discontinue products at any time (10/05). Ballard, and Power to Change the World are registered trademarks of Ballard Power Systems Inc. 2004 Ballard Power Systems Inc. SPC5101006-0D PRINTED IN CANADA

Ballard® fuel cell power

Mark9 SSLTM

Mark9 SSL Ballard Power Systems offers a proton exchange membrane (PEM) fuel cell stack based on our proven, fourth generation, Mark 902 transportation stack technology. Available now to customers with fuel cell stack integration capabilities, the Mark9 SSL is designed to perform in rugged conditions and is available in a scalable series depending upon customer requirements. The Mark9 SSL can be configured for motive or stationary power applications. Stacks are available in power increments from 4 kilowatts to 21 kilowatts. The Mark9 SSL provides stable electrical power to a system over a wide range of operating and environmental conditions. A liquid-cooled, hydrogen-fueled product, the Mark9 SSL uses Ballard’s off-the-shelf fuel cell components. The Mark9 SSL features fast, dynamic response, robust and reliable operation and durable packaging. The Mark9 SSL establishes a new standard of performance by optimizing reliability, power density and compatibility with customer system requirements. Please contact us for product availability and pricing.

t) 604-454-0900 f) 604-412-4700 www.ballard.com

Ballard Power Systems Inc. 4343 North Fraser Way Burnaby, British Columbia Canada V5J 5J9

Specifications

Rated Power [kW] 1 4 9 14 21

DC Voltage (at 300A) 1 16 32 48 70

Mass (with no coolant) [kg] 7.2 10 13 17

Stack core length [mm] 108 172 234 326

Stack core width [mm] 760 760 760 760

Stack core height [mm] 60 60 60 60

Type: PEM (Proton Exchange Membrane) fuel cell stack

Performance: Maximum current 300A

Shock and vibration Automotive 2

Fuel: Fuel composition (pre-humidification)

≥85% H2 3

≤15% inert 3

Oxidant: Oxidant composition (pre-humidification)

Compressed ambient

Stack Temperatures: Storage temperature

2 to 40° C (36 to 104° F)

Start up temperature ≥2° C (≥36° F)

Fluid inlet temperature (operating)

2 to 65° C 1 (36 to 150° F) 1

External stack temperature (operating)

-25 to 75° C 4 (-13 to 167° F) 4

Additional information available upon request.

1 Values achieved at Ballard-specified conditions at beginning of operational life. 2 Vibration: USABC 10. Shock: 5g sections of IEC 600068-2-27 Ea and

IEC 600068-2-29 Eb. 3 Measured at the inlet, on a dry, molar basis. When using a recirculating fuel loop,

incoming hydrogen purity of ≥99.99% H2 is recommended. Purging must be sufficient to maintain inert concentrations below stated maximum.

4 Insulate stack whenever possible to minimize heat loss.

Batterypack

AC motorPC Pedals

Auxiliarybattery

Instrumentationpanel

Batterymanagement

system

AC Induction Motor Drive

TIM ( Traction Inverter Module ) 600 is a vector control AC motor drive especially designed for electric and hybridvehicles. The main features are:- Digital Signal Processing based control- Smooth Motor Control- Complete adjustable regenerative braking levels on releasing the accelerator and on pressing the brake pedal.- The parameter can be update in real-time for adjustable the performance on own special features, the programming is quickly achieved with a standard laptop computer by RS-232 and by CAN - BUS (optional)- Total management of selecting gear change ( Drive, Economy, Reverse, Parking)- Economy mode: reducing the performance for increasing the autonomy.- Programmable digital tachometer output.- Autotuning of motor characteristics.- Under and Over Voltage Protection- Overcurrent Protection- Overtemperature Protection

The information contained herewith is subject to change without notice

WaterCooling

CAN BUS2.0 Boption

Type TIM 600

TIM 600

115

ø8.5 ø2

0

212

232

248

270

326

Status: Marzo '03

D i v i s i o n e E n e r g i e A l t e r n a t i v e

- D E A

TEL: +41 (0)91 6415311 FAX: +41(0)91 6415333 E-mai l : in [email protected] internet : ht tp: / / www.mes-dea.ch Via Laveggio, 15 CH - 6855 Stabio - Switzerland

C o m p o n e n t s f o r E l e c t r i c V e h i c l e s

Technical dataInput voltage: 80÷400 VDC

Input voltage Service: 12 VDC

(24 VDC optional)Nominal output current: 225 AmpsMax. output current: 340 ArmsControl type: Vector controlModulation PWMSwitching frequency: 3-9 kHzProtection: IP54Communication: RS-232

CAN – BUS ( optional )Cooling: 50% H2O + 50% GlicolDrop pressare cooling : 93mBar @ 8 Liter/min.Least flow rate cooling: 8 Liter/min.Operative Temperature: -20 to +65 °CWeight: 10Kg.

57

Horizontal Multi-Stage Close Coupled Pumpsin stainless steelMXHMXH

ConstructionHorizontal multi-stage close coupled pumps in chrome-nickel stainless steel.Compact and robust construction, without protruding flangeand with single-piece lantern bracket and base.Single-piece barrel casing, with front suction port abovepumps axis and radial delivery at top.Filling and draining plugs on the middle of the pump, acces-sible from any side (like the terminal box).

ApplicationsFor water supply.For clean liquids, without abrasives, which are non-aggressivefor stainless steel (with suitable seal materials, on request).Universal pump, for domestic use, for civil and industrialapplications, for garden use and irrigation.

Operating conditionsLiquid temperature from - 15 °C to + 110 °C.Ambient temperature up to 40 °C.Maximum permissible pressure in the pump casing: 10 bar.Continuous duty.

Motor2-pole induction motor, 50 Hz (n = 2900 rpm).MXH: three-phase 230/400 V ± 10%.MXHM: single-phase 230 V ± 10%, with thermal protector.

Capacitor inside the terminal box.Insulation class F.Protection IP 54.Constructed in accordance with: IEC 34;

IEC 38; IEC 335-1, EN 60335-1;IEC 335-2-41, EN 60335-2-41;

EN 60529.

Special features on request- Other voltages. - Frequency 60 Hz (as per 60 Hz data sheet).- Protection IP 55. - Special mechanical seal- Pump casing seal rings in FPM (Viton).- Higher or lower liquid or ambient temperatures.

MaterialsComponent Material

Pump casing Chrome-nickel steel 1.4301 EN 10088 (AISI 304)Stage casing Chrome-nickel steel 1.4301 EN 10088 (AISI 304)Wear ring PTFE (Teflon)Impeller Chrome-nickel steel 1.4301 EN 10088 (AISI 304)Casing cover Chrome-nickel steel 1.4301 EN 10088 (AISI 304)Spacer sleeve Chrome-nickel steel 1.4301 EN 10088 (AISI 304)

Pump shaft Chrome-nickel steel 1.4305 EN 10088 (AISI 303)Plug Chrome-nickel steel 1.4305 EN 10088 (AISI 303)

Mechanical seal with seat Ceramic alumina, carbon, EPDMaccording to ISO 3069 (Other materials on request)

m /hl/min

3

Q

H

MXH 8MXH 4MXH 2

0

0

00 14

72.978

4 8 122 6 10

2010 40 60

2010 40

30 50

30 50

U.S. g.p.m.

0 50 100 150 200

Imp. g.p.m.

20

10

30

40

50

60

180

200

160

140

120

100

80

60

40

20

0

mft

Coverage chart n ≈ 2900 rpm

6

C R E A T I V E T E C H N O L O G Y

58


MXH 802E

MXH 803

MXH 804

MXH 805

A

3,7

5

6,4

7,5

MXHM 802E

MXHM 803

MXHM 804

Qm /h

l/min

Hm

3

3 ~ 230 V400 V 1 ~230 V P1 P2

22,5

48

60

20,5

42,5

54

20

41

52 49,5

39

19

47

37

18

43,5

34,5

16,5

39,5

32

15

35

28

13

29,5

24

11

24

19,5

8,5

36 32 30,5 29 27,5 25,5 23 20 17 14

166133

0

0

5

83,3

6

100

7

116

8 9

150

10 11

183

12

200

13

216A

2,2

2,9

3,7

4,3

A

5,7

7,4

9,2

kW

1,2

1,5

2

kW

0,75

1,1

1,5

1,8

HP

1

1,5

2

2,5

MXH 205E 3,5 2 MXHM 205E 5,4 1,2 0,75 1

MXH 204E 2,8 1,6 MXHM 204E 4,2 0,9 0,55 0,75

MXH 203E 2,4 1,4 MXHM 203E 3 0,65 0,45 0,6

MXH 202E 1,7 1 MXHM 202E 2,3 0,5 0,33 0,45 22

33

45

57

22,5

33

44,5

56,5

20

30

40,5

52

19,5

29

38

50

18,5

27,5

36,5

47,5

17,5

26

35

45,5

16

24,5

33

43

15

23

31

40

12,5

19,5

26

33,5

9,5

15

20

26

6

9,5

12,5

16,5

20,5

31

42,5

53,5

19

29

40

50,5

18

27

37,5

47

16

24

34

43

14

21,5

30

38

11,5

18

25,5

32,5

9

14,5

21

26,5

8

12,5

18

23

A A A kW kW HP 0 16,6 25 33,3 41,6 50 58,3 66,6 70,8

3 ~ 230 V 400 V 1 ~ 230 V P1 1 1,5 2 2,5 3 3,5 4,25P2 0 4Q

m /h

l/min

Hm

3

MXH 402E

MXH 403E

MXH 404E

MXH 405E

A

2,4

2,8

3,5

4,7

A

1,4

1,6

2

2,7

MXHM 402E

MXHM 403E

MXHM 404E

MXHM 405

A

3

4,2

5,4

7,4

kW

0,65

0,9

1,2

1,5

kW

0,45

0,55

0,75

1,1

HP

0,6

0,75

1

1,5

Qm /h

l/min

Hm

3

3 ~ 230 V400 V 1 ~230 V P1 P2

83,366,6

0

0

2,25

37,5

3

50

3,5

58,3

4 4,5

75

5 6

100

7

116

8

133

TYPE

TYPE

(1)�Filling (2)�Draining (3) MXHM

331

331

381

405

331

357

381

405

381

MXH 202E - MXHM 202E







MXH 405E


94

94

118

142

94

94

118

142

118

182

182

206

230

182

182

206

230

206

L L1 L2

mm

MXH MXHM

6,8

7,6

10

11,5

7,6

9,3

10,8

13

10,6

6,9

7,7

11

12,5

7,7

10,3

11,8

11,6

kg

176

176

189

189

176

189

189

189

189

H

98,5

98,5

112

112

98,5

112

112

112

112

w

464

440

470

500

MXHM 405

MXH 803 - MXHM 803

MXH 804 - MXHM 804

MXH 805

142

118

148

178

230

206

236

266

L L1 L2

mm

MXH MXHM

15,8

18,2

19

18

16,9

19,2

kg

G 1

L1

L

G 1

1/2

(G

1 1

/4 M

XH

M 4

05

)

L2 155

31

888

ISO 228

ISO

22

8

3.93.020

10 9

112

12

7 18

4

19

2

146

160

92

30

(3)

(1)

(2)

ISO

22

8

ISO 228

88

28

4.93.209

8

L2 w

G 1

1/4

(

G 1

1/2

MX

H 8

02

)

(1)

L

G 1

L1

10

30

112

146

18

4

H

9

160

12

7

(2)

Performance n ≈ 2900 rpm

Dimensions and weights

P1 Max. power input. Test results with clean cold water, without gas content.P2 Rated motor power output. Tolerances according to ISO 9906, annex A.


59

0 1 2 3 4 5

0 20 40 60 80

30

40

50

20

30

40

50

60

10

50

100

150

ft

5 10 15Imp. g.p.m.0

0 50

2

4

1 2 3 4 72.976 72.9770

10

Hm

m /h�

l/min

3

Q

NPSH

ft

m /h��

3Q

MXH 205E

MXH 204E

MXH 203E

MXH 202E

"%

m��

0 2 4 6 8

0 50 100

40

50

60

20

30

40

50

60

10

0 10

50

100

150

ft

20Imp. g.p.m.

00

2

4

2 4 6 80

10

Hm

m /h�

l/minQ

NPSH

3Q

ft

MXH 405E

MXH 404E

MXH 403E

MXH 402E

3

%"

m

0 14

0

50

60

70

20

30

40

50

60

10

Imp g.p.m.0

01

2

3

4

4 6 8 10 12

144 6 8 10 12 72.397/1

50 100 150 200

6

8

20 30 40 50

50

100

150

200

ft

ft

MXH 805

MXH 804

MXH 803

MXH 802E

Hm

m /h�

l/minQ

NPSH

3

%"

m

m /hQ 3

Characteristic curves n ≈ 2900 rpm


6

Test results with clean cold water, without gas content.+ 0,5 m security margin on NPSH-value is necessary.Tolerances according to ISO 9906, annex A.


60


3.93.001

Extra safetyagainst running dry, with the suction port above pump axis.

ReliableAll hydraulic parts in contact with the pumped liquid are of stainless steel.

For liquids from -15 °C to 110 °C.

RobustSingle-piece, thick barrel casing.

CompactSingle-piece lantern bracket and base.

Without protruding flange.

Greater protectionagainst leakage, with the pump casing cover separated from the motor

shield. Possibility of inspecting the seal through the side apertures

between the two walls.

Greater protection against water entering the motor from outside provided

by an extension of the pump casing around the lantern bracket.

Features


EATON M62 SUPERCHARGER From Eaton website

APPENDIX D

LIGHT AIRCRAFT AND RECONNAISSANCE AIRCRAFT STUDY

Country Manufacturer ModelFirst Flight Year

Empty Weight

Max T/O Weight Payload Engine Power Propelle

r bladesPropeller Diameter

Cruise speed

Max speed

Stall speed Climb rate Span Length Aspect

RatioT/O

Distance range Ceiling PAX Engine Weight

Empty W without Engine

kW/kg

Empty W without Engine /

PAX

T/O Weight / Rated Power

Wing area gross

Wing Loading

[kg] [kg] [kg] [kW] [m] [km/h] [km/h] [km/h] [m/min] [m] [m] [m] [km] [ft] [kg] [kg] [kW/kg] [kg] [kg/kW] [m2] [N/m2]

US Commander 115 2000 953 1474 Lycoming IO-540-T4B5 194 3 1.96 296 304 112 326 9.98 7.59 7.1 349 1583 16800 4 176 777 0.132 194.25 7.598 14.12 1024.08

US Zodiac CH-640 2001 520 998 Lycoming IO-360-A1A 134 2 241 253 94 290 9.6 7.01 6.6 290 820 12800 4 133 387 0.134 96.75 7.448 13.94 702.32

US Tiger AG-5B 2001 680 1089 Lycoming O-360-A4K 134 2 1.93 265 274 104 259 9.6 6.71 7.1 260 1261 13800 4 133 547 0.123 136.75 8.127 13.02 820.51

US Cessna 172 Skyhawk 1986 735 1111 Lycoming IO-360-L2A 119 2 1.9 226 227 95 219 11 8.28 7.5 288 1074 13500 4 133 602 0.107 150.5 9.336 16.17 674.02

US Lancair Columbia 350 2000 1043 1542 Continental IO-550-N 231 3 1.96 352 435* 132 427 10.97 7.67 9.2 397 2444 14000 4 186.9 856.1 0.150 214.025 6.675 13.12 1152.97

US Luscombe 11E 2002 658 1034 Continental IO-360-ES4 138 2 - 217 241 87 267 11.73 7.32 8.9 274 926 16000 4 145.8 512.2 0.133 128.05 7.493 15.51 654.00

Russia Ilyushin Il-103 1995 900 1310 270 Continental IO-360-ES2B 157 2 1.93 180 220 117 190 10.56 8 7.6 380 800 9800 4 145.8 754.2 0.120 188.55 8.344 14.71 873.63

Austria Diamond DA-40D 1997 735 1150 4 Pax + 30kg Centurion 1.7 turbo-diesel 99 3 - 198-239 239 97 262 11.94 8 10.53 310 - - 4 - - 0.086 - 11.616 13.54 833.20

France DynAero MCR4S 2000 350 750 4Pax + 40kg Rotax 914 UL 84.6 2 1.7 272 287 100 228 8.72 9.2 6.7 350 1846 - 4 56.6 293.4 0.113 73.35 8.865 8.30 886.45

India Nat aerosp lab HANSA-3 1993 485 750 SMA SR305-230 123 2 1.7 213 - 89 198 10.47 8.8 7.6 413 842 - 4 - - 0.164 - 6.098 12.47 590.02

France Socata TB-9GT 2000 684 1060 Lycoming O-360-A1AD 134 2 1.88 194 213 107 203 10.01 8 7.72 570 1046 11000 4 133 551 0.126 137.75 7.910 11.90 873.83

Canada CLASS Bush Caddy 2000 562 1134 Lycoming IO-360 134 2 - 193 217 68 366 10.97 6.9 7.75 76 1480 13000 4 133 429 0.118 107.25 8.463 15.61 712.65

Poland PZL-Swidnik I-23 Manager 1999 690 1150 330 Lycoming IO-360-A1A 147 2 1.83 225 300 125 330 8.95 7.1 8 250 1450 - 4 133 557 0.128 139.25 7.823 10.00 1128.15

Germany Ruschmeyer R 90-230 RG 1990 898 1350 452 Lycoming IO-540-C4D5 171.5 4 1.88 - 322 124 347 9.5 7 7.93 260 1378 16060 4 170 728 0.127 182 7.872 12.94 1023.45

France Robin Aviation Dr 500 President 1997 560 1059 4pax + 60kg Lycoming IO-360-A1B6 149 2 1.88 - 265 272 255 - 7.22 - 463 1842 - 4 133 427 0.141 106.75 7.107 - -

MAX 1043.00 1542.00 231.00 1.96 352.00 322.00 272.00 427.00 11.94 9.20 10.53 570.00 2444.00 16800 187 856 0.164 214.03 11.62 16.17 1152.97AVERAGE 696.87 1130.73 143.27 1.87 239.50 258.62 114.87 277.80 10.29 7.65 7.87 328.67 1342.29 13676 139 571 0.127 142.71 8.05 13.24 853.52MIN 350.00 750.00 84.60 1.70 180.00 213.00 68.00 190.00 8.72 6.71 6.60 76.00 800.00 9800 57 293 0.086 73.35 6.10 8.30 590.02

UK Edgley Optica OA7-300 1979 948 1315 231 Lycoming IO-540-V4A5D 194 - 1.22 191 259 108 247 12 8.15 9.09 331 619 14000 3 170 778.00 0.148 259.33 6.778 15.84 814.40

Australia Seabird SB7L-360 Seeker 1989 604 897 2 pax + 45kg Lycoming O-360-B2C 125 2 1.77 207 243* 97 248 11.07 7.01 9.4 265 881 15250 2 122.5 481.50 0.139 240.75 7.176 13.05 674.30

ItaliaSocieta

Aeronautica Italia

SAI G97 SPOTTER 1998 298 450 2pax Rotax 912 59.6 2 1.66 175 200 169 280 8.25 6.7 6.24 85 800 11500 2 56.6

241.40 0.132 120.70 7.55010.30

428.59

US Schweitzer SA 2-37 1985 1157 1950 322 Lycoming TIO-540-AB1AD 186 3 - 157 305* 125 - 21.7 8.79 25 449 370 24000 2 - - - - 10.484 18.68 1024.06

MAX 1157.00 1950.00 194.00 1.77 207.00 259.00 169.00 280.00 21.70 8.79 25.00 449.00 881.00 24000 170 778 0.148 259 10 19 1024AVERAGE 751.75 1153.00 141.15 1.55 182.50 229.50 124.75 258.33 13.26 7.66 12.43 282.50 667.50 16188 116 500 0.140 207 8 14 735MIN 298.00 450.00 59.60 1.22 157.00 200.00 97.00 247.00 8.25 6.70 6.24 85.00 370.00 11500 57 241 0.132 121 7 10 429

4 Seater Aircraft and Reconnaissance Aircraft Survey

APPENDIX E

MODIFIED OPTIMISATION SPREADSHEET

Spreadsheet

PARAMETERS REQUIREMENTS ANALISISA t/c N eng ToLength LLength Vstall Vd Vmax Vcruise Sigma cr MaxRge Vv V-A(man) N Pay/crew Take off Acc.Stop Sec Seg. Climb[-] [-] [-] [m] [m] [m/s] [m/s] [m/s] [m/s] [kg/m3] [km] [m/s] [m/s] [m/s-2] [-] (Mg/S)o 1st.app (T/Mg)o (T/Mg)o Tau co (Cd)co/Clus (T/Mg)o Y ceil; Y cr9 0.17 2 300 550 20 70 65 52.7 0.909 800 10 52 4.125 4 [N/m2] [-] [-] [-] [-] [-] [-] [kg/m3] [kg/m3]

400 0.441707 0.379142 0.200057 7.5352 0.0420307 0.404003 0.768 0.909ASSUMED 450 0.507035 0.410765 0.223501 7.76495 0.0420307 0.409031 0.768 0.909

M1/Mo S^-0.1 Rw Type Fac TE flap SS alpha SSgamma Flap Fac a Cruise T/Mg ass Tan gam des Mu G Sigma ceil Cl Fac Z 500 0.577083 0.438591 0.247597 7.976398 0.0420307 0.413581 0.768 0.909[-] [-] [-] [-] [-] [-] [-] [-] [m/s] [-] [-] [-] [kg/m3] [-] 550 0.652611 0.462363 0.272481 8.172633 0.0420307 0.417741 0.768 0.9091 0.8 4 2.25 SglSlot(L) 2.74 0.02 1 322.16 0.25 0.0524 0.38 0.768 1 600 0.73452 0.481873 0.298293 8.355993 0.0420307 0.421575 0.768 0.909

0.718957 650 0.823895 0.496963 0.325179 8.528299 0.0420307 0.425133 0.768 0.909INITIAL CALCULATED VALUES 700 0.922068 0.507533 0.353293 8.690995 0.0420307 0.428454 0.768 0.909Ml/Mo Mcr/Mo (Clm)o ae Del flp(TO) Del flap(L) (Cl use)o (Cdz)o Landg (L) Man(Cl )o (Clm)o TO (Cl)us o (Clmax) o (Cl)a o ML/M0 750 1.030699 0.513538 0.382808 8.84525 0.0420307 0.431569 0.768 0.9091 0.997142 1.5 0 1 0.65 0.037928 7.66944 1.5 1.5 1.2 2.5 1.6 0.944 800 1.15189 0.514986 0.413917 8.992024 0.0420307 0.434503 0.768 0.909

No reverse Thrust 613 0.756991 0.486227 0.305174 8.401792 0.0420307 0.422524 0.768 0.909PRELIMINARY CALCULATIONScos delta Delta deg Wave Dr.F (Cdz)cr (Cd)co (Kv)o (Kv)cr Mean R/C to Cruise Alt Landing-rev.thrust1 0 1.61E-14 0.036687 0.050437 0.044169 0.044169 (Mg/S)o Fact.Qv f(drag) 1st.app X (T/Mg) 1 Tau Cl (T/Mg)o (T/Mg)o Landg L Llength

[N/m2] [m/s] [-] [-] [-] [-] [-] [-] [-] [-] [m/s]Cl max(L) Cl us Cl a Cl use Va calc Va (Mg/S)o ld (Mg/S)o gt Cor.Llegth Man(Mg/S) Man Cl 400 20.77629 0.155784 0.428814 0.985269 0.434005 0.193189 0.41147 0.614813 5.296197 684.5035568[-] [-] [-] [-] [m/s] [m/s] [N/m2] [N/m2] [m] [N/m2] [-] 450 22.03658 0.155784 0.409168 0.985064 0.414134 0.187473 0.400321 0.550619 5.445404 686.88154712.5 1.2 1.6 0.65 42.60 25 613 990.133 722.3271 906.7108 1.5 500 23.2286 0.155784 0.392558 0.984869 0.397335 0.182503 0.390775 0.499694 5.573992 688.9309099

550 24.36236 0.155784 0.378278 0.984684 0.382892 0.178121 0.382477 0.462363 5.674675 690.5355466600 25.44565 0.155784 0.36583 0.984507 0.370303 0.174213 0.375173 0.481873 5.621343 689.6855759650 26.48467 0.155784 0.354855 0.984338 0.359203 0.170693 0.368677 0.496963 5.581167 689.0452614700 27.48445 0.155784 0.345085 0.984175 0.349321 0.167497 0.362851 0.507533 5.553563 688.6053277750 28.4491 0.155784 0.336314 0.984017 0.340451 0.164576 0.357585 0.513538 5.538072 688.3584429800 29.38211 0.155784 0.328384 0.983865 0.332431 0.16189 0.352795 0.514986 5.534358 688.2992506613 25.71984 0.155784 0.362847 0.984463 0.367286 0.173263 0.373412 0.486227 5.609657 689.4993345

End of climb Cruise Max speed(Mg/S)o Fact Qv (T/Mg)1 Tau ceil. (T/Mg)o Cl Cd L/D (T/Mg)cr (T/Mg)o Cl Cd L/D (T/Mg) max (T/Mg)o

[N/m2] [m/s] [-] [-] [-] [-] [-] [-] [-] [-] [-] [-] [-] [-] [-]400 20.77629 0.143223 0.193189 0.160353 0.257734 0.039621 6.504931 0.1537295 0.401303 0.169420452 0.037955 4.46370785 0.224029 0.614813450 22.03658 0.142002 0.187473 0.161172 0.28995 0.040401 7.176881 0.1393363 0.362669 0.190598009 0.038292 4.97750843 0.2009037 0.550619500 23.2286 0.14097 0.182503 0.161957 0.322167 0.041272 7.806016 0.1281063 0.332374 0.211775565 0.038668 5.47673354 0.1825906 0.499694550 24.36236 0.140082 0.178121 0.16271 0.354384 0.042234 8.39089 0.1191769 0.308139 0.232953122 0.039084 5.96028611 0.1677772 0.458418600 25.44565 0.139309 0.174213 0.163433 0.386601 0.043289 8.930738 0.1119728 0.288444 0.254130678 0.03954 6.42720727 0.1555886 0.424377650 26.48467 0.138627 0.170693 0.164126 0.418817 0.044435 9.425424 0.106096 0.272236 0.275308235 0.040035 6.87667863 0.145419 0.395899

Efficiency 0.5 List of Modifications 700 27.48445 0.138019 0.167497 0.164792 0.451034 0.045673 9.875369 0.101262 0.258763 0.296485791 0.04057 7.30802244 0.1368359 0.371791Number of fuel cell used 8 750 28.4491 0.137474 0.164576 0.165433 0.483251 0.047002 10.28147 0.0972623 0.247475 0.317663348 0.041144 7.72069959 0.1295219 0.351177Power of each stack 21 kW T/O T/Mg approximations 800 29.38211 0.136982 0.16189 0.16605 0.515468 0.048423 10.64504 0.0939404 0.237957 0.338840905 0.041758 8.11430588 0.1232391 0.3334

168 kW 134.24 End of Climb Tau ceiling 613 25.71984 0.139123 0.173263 0.163616 0.394977 0.043578 9.063693 0.1103303 0.283929 0.259636843 0.039665 6.54577834 0.1527702 0.416492T/Mg at cruise SUMMARY (T/Mg)o

sfc 0.002922 kg/s Available hydrogen mass stored onboard (Mg/S)o Take off Acc.Stop Sec.Seg Climb End Clmb Cruise Max.speed Man Landing Gust Sen2.92 g/s Sfc 400 0.379142 0.200057 0.404003 0.41147 0.160353 0.401303 0.61481260.61 N/kWs 450 0.410765 0.223501 0.409031 0.400321 0.161172 0.362669 0.5506189

Ratio Start of Climb/ T/O mass 500 0.438591 0.247597 0.413581 0.390775 0.161957 0.332374 0.4996942Weight Specific power of powerplant 550 0.462363 0.272481 0.417741 0.382477 0.16271 0.308139 0.4584181Fuel cells 136 kg Wing Weight penalty of 20% dur to fuel tanks 600 0.481873 0.298293 0.421575 0.375173 0.163433 0.288444 0.4243767Inverter 16.8 kg 650 0.496963 0.325179 0.425133 0.368677 0.164126 0.272236 0.3958988Electric Motor 35.28 kg 700 0.507533 0.353293 0.428454 0.362851 0.164792 0.258763 0.3717907Supercharger 20.8 kg 750 0.513538 0.382808 0.431569 0.357585 0.165433 0.247475 0.3511768Air conducts and Radiator 20 kg 800 0.514986 0.413917 0.434503 0.352795 0.16605 0.237957 0.3334003Water cooling system 12 kg 850 0.486227 0.305174 0.422524 0.373412 0.163616 0.283929 0.4164923Heat exchanger 5 kg 906.7108 0

906.7108 0.8TOTAL 245.88 kg 613 0

613 0.8990.133 0990.133 0.8

RESULT(Mg/S)o (T/Mg)o L/D Case Cl (Cd)z o (Cd)z cr (Kv)o (Kv)cr Rev th lnd SP (P/Mg)

[N/m2] [-] [-] [-] [-] [-] [-] [-]613 0.486227 9.063693 0.394977 0.037928 0.036687 0.044169 0.0441687 689.4993 12.48824 0.010684908

CorrectedCLIMB PATHClimb EAS ClEAS H2 ClEAS sig Clmb Mn

[m/s] [km] [kg/m3] [-]37.49952 3 0.909 n/a

ASSUMEDLambda (P/Mg)eng Op It Fac ApFuel/Mo Vbar Vv bar

[-] [kW/kg] [kg] [kg]0.4 0.065 3 0.01 0.85 0.065

Specific to fuel cellsINPUT DATA

Fus.L Fus.B Fus.H c1 c2 c3 c4 c5 Payload Nbar Neng Overall L Del l w l Fus l Tail l PP l SYS l PAY l OP IT Del lwg fuel l fus fuel l nose g r del l mn gr[m] [m] [m] [-] [-] [-] [-] [-] [kg] [-] [-] [m] [m] [m] [m] [m] [m] [m] [m] [m] [m] [m] [m]9 1.7 2.05 0.0016 0.034 1.8 0.16 1.24 400 4.125 2 9 0.1 4.15 8.5 4 1.75 1.75 1.75 0.33 5 1.05 0

Might be sllightly increasedCALCULATIONS Cruise Climb(S/Mo)^.45 p bar Req(T/Mg) Av(T/Mg) Av./Req (c)des (c)od Z X1 Fact Qv (T/Mg) (Vv)EAS Dist EAS Wf/(Mg)o Desc Dist[m2/kg]^.45 [bar] [-] [-] [-] [N/kWs] [N/kWs] [-] [-] [m/s] [-] [m/s] [km] [km]0.155559 0 0.11033 0.099626 0.902977 0.61 0.61 1 1.001418 25.71984 0.148155 2.369857 48.63035 0.0196726 12.726

Mc1/Mo M fus c1 bar Mpp/Mo Msys/Mo M op it M fixed Net Range Log 10 Mc1/Mc2 Mc2/Mo Mf/Mo Kappa Mo xbar(0.25)/Mc1/Mc2 root chrd

[-] [-] [-] [-] [-] [-] [kg] [km] [-] [-] [-] [-] [-] [-]0.980327 161.3292 0.044168 0.29589 0.16 12 573.3292 738.6436 0.008004 1.0186017 0.962425 0.047575 0.503465128 0.25

ANALYSIS SUMMARY-FINAL RESULTS(Mo)est1 (Mo)est2 Kappa*Mo Mlift. surf. (Mo)calc error M wing 83.81 7.12 Wing Area 27.10 Wg Apex 2.401522.646 1089.978 548.766 103.9192 1226.014 136.0363 M Fus 161.33 13.70 Wing span 15.62 Hor tl area 7.09

M tail 20.11 1.71 Mean chrd 1.74 Vert tl area 4.88l CG l WG APX l TL ARM S Hor Tail S Vert Tail M gear 61.30 5.21 A 9.00 Prop dia. 1.732.860283 2.399909 5.639717 7.08752 4.877881 M power p 245.88 20.88 Del 0.25 0.00

M sys 134.86 11.45 Lambda 0.40 M wng/S 3.09M op.it 12.00 1.02 t/c 0.17 (Mg/S)o 613.00M oew 719.29 61.08 SP 12.49 (T/Mg)o 0.49M pay 400.00 33.97 Power 123.44 Wing fuel 58.33M fuel 58.33 4.95 Fuel Mass 246.02 Fus fuel 0.00Mo 1177.62 100 avail.in wing

%

SUBSONIC FLIGHT -FUEL CELL AIRCRAFT- OPTIMISER

Page 1

design study of a fuel-cell powered aircraft 06

Documents