Human Powered Aircraft Report

MECHENG 4108

Aircraft Design Project

HUMAN POWERED AIRCRAFT

Marion Byrne 1141356Jesse Coombs 1147910Rebecca Mills 1113961Mai-Chi Nguyen 1141502

April, 2010

Executive Summary

This project details the preliminary conceptual design of a human powered aircraft.The aircraft is design to win the Kremer Marathon prize by completing a specified mis-sion profile in the required time frame. This report details the technical task, statisticalanalysis, preliminary designs of the aircraft components, weight balance analysis andtechnical drawings. The final design features a high wing, large aspect ratio tractorpropeller aircraft, driven by the pedalling of a pilot in a faired fuselage, and controlledwith ailerons and all moving tail control surfaces.

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Contents

1 Introduction 2

2 Feasibility Study 32.1 The Product . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.2 Mission Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.3 Competitors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.4 Regulations and Standards . . . . . . . . . . . . . . . . . . . . . . . . . 52.5 Patents and Intellectual property . . . . . . . . . . . . . . . . . . . . . . 62.6 Environmental Issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72.7 Financial Issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72.8 Technology Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82.9 Risk Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82.10 Aircraft type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92.11 Technical Task . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2.11.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122.11.2 Standard requirements . . . . . . . . . . . . . . . . . . . . . . . . 122.11.3 Performance parameters . . . . . . . . . . . . . . . . . . . . . . . 132.11.4 Technical level of the product . . . . . . . . . . . . . . . . . . . . 132.11.5 Economical parameters . . . . . . . . . . . . . . . . . . . . . . . 142.11.6 Power plant type and requirements . . . . . . . . . . . . . . . . . 142.11.7 Main system parameters requirements . . . . . . . . . . . . . . . 142.11.8 Special systems and miscellaneous requirements . . . . . . . . . . 152.11.9 Reliability and maintainability . . . . . . . . . . . . . . . . . . . 152.11.10 Unification level . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3 Literature Review 163.1 Successful Designs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3.1.1 Gossamer Condor . . . . . . . . . . . . . . . . . . . . . . . . . . 173.1.2 Gossamer Albatross . . . . . . . . . . . . . . . . . . . . . . . . . 173.1.3 Musculair I & II . . . . . . . . . . . . . . . . . . . . . . . . . . . 193.1.4 Daedelus 88 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

4 Conceptual Design 214.1 Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

4.1.1 Available power . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214.1.2 Recumbent versus upright pedal cycle . . . . . . . . . . . . . . . 234.1.3 Final power plant design . . . . . . . . . . . . . . . . . . . . . . . 25

4.2 Statistical Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.2.1 Empty weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.2.2 Wing area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.2.3 Wing span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 274.2.4 Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

4.3 Weight Analysis & Sensitivity . . . . . . . . . . . . . . . . . . . . . . . . 274.4 Sizing Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

4.4.1 Size to cruise requirements . . . . . . . . . . . . . . . . . . . . . 30

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4.4.2 Size to manoeuverability requirements . . . . . . . . . . . . . . . 304.4.3 Sizing to climb requirements . . . . . . . . . . . . . . . . . . . . 314.4.4 Sizing for human power . . . . . . . . . . . . . . . . . . . . . . . 324.4.5 Matching diagram and design point . . . . . . . . . . . . . . . . 32

4.5 Concept Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.5.1 Concept 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.5.2 Concept 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.5.3 Concept 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 374.5.4 Concept selection . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

5 Preliminary Design 395.1 Wing design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

5.1.1 Wing planform design . . . . . . . . . . . . . . . . . . . . . . . . 395.1.2 Wing airfoil selection . . . . . . . . . . . . . . . . . . . . . . . . . 405.1.3 Design lift coefficient and power factor . . . . . . . . . . . . . . . 415.1.4 Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 435.1.5 Moment coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . 435.1.6 Airfoil selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . 445.1.7 The DAE series . . . . . . . . . . . . . . . . . . . . . . . . . . . . 455.1.8 Wing Structural Considerations . . . . . . . . . . . . . . . . . . . 47

5.2 Empennage design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 485.2.1 Empennage Planform . . . . . . . . . . . . . . . . . . . . . . . . 485.2.2 Horizontal tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . 495.2.3 Vertical tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 495.2.4 Tail airfoil selection . . . . . . . . . . . . . . . . . . . . . . . . . 50

5.3 Control surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 525.3.1 Design of control surfaces . . . . . . . . . . . . . . . . . . . . . . 525.3.2 Handling of control surfaces . . . . . . . . . . . . . . . . . . . . . 53

5.4 Propeller design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 535.4.1 Propeller location . . . . . . . . . . . . . . . . . . . . . . . . . . 545.4.2 Number of propeller blades . . . . . . . . . . . . . . . . . . . . . 555.4.3 Propeller optimisation . . . . . . . . . . . . . . . . . . . . . . . . 555.4.4 Further Propeller Guidelines . . . . . . . . . . . . . . . . . . . . 585.4.5 Propeller design summary . . . . . . . . . . . . . . . . . . . . . . 60

5.5 Drive train design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 605.5.1 Possible power transmission mechanisms . . . . . . . . . . . . . . 605.5.2 Drive train selection . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.6 Landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 635.6.1 Landing gear type . . . . . . . . . . . . . . . . . . . . . . . . . . 635.6.2 Landing gear arrangement . . . . . . . . . . . . . . . . . . . . . . 645.6.3 Landing gear sizing . . . . . . . . . . . . . . . . . . . . . . . . . . 655.6.4 Shock absorption . . . . . . . . . . . . . . . . . . . . . . . . . . . 665.6.5 Braking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 665.6.6 Driven Main Wheel . . . . . . . . . . . . . . . . . . . . . . . . . 675.6.7 Landing Gear Positioning . . . . . . . . . . . . . . . . . . . . . . 675.6.8 Final landing gear design . . . . . . . . . . . . . . . . . . . . . . 67

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5.7 Fuselage and frame design . . . . . . . . . . . . . . . . . . . . . . . . . . 68

6 Stability analysis 706.1 Aircraft Centre of Gravity . . . . . . . . . . . . . . . . . . . . . . . . . . 706.2 Aircraft Neutral Point . . . . . . . . . . . . . . . . . . . . . . . . . . . . 716.3 Static Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

7 Conclusion 73

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List of Figures

2.1 The Kremer Marathon Competition course (Royal Aeronautical Society1988) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

3.2 The Gossamer Condor in flight (Don Monroe 1977) . . . . . . . . . . . . 183.3 The Gossamer Albatross in flight (Jim Harrison) . . . . . . . . . . . . . 183.4 The Musculair II in flight (Ernst Schoberl 1985) . . . . . . . . . . . . . 193.5 The Daedelus 88 in flight (John McIntyre, 1988) . . . . . . . . . . . . . 204.6 Maximum output of mechanical power vs total duration of exercise for

various types of exercise : x- cycling, + rowing, ∆- running uphill, *-cycling and turning a hand crank (the around a symbol indicates per-formance by a champion athlete) (Wilkie, 1960) . . . . . . . . . . . . . . 22

4.7 Power requirements for touring bicycle, upright and crouched racing bi-cycles and Vector recumbent human powered vehicle (Hennekam, 1990) 24

4.8 Wing area versus empty weight for prototype aircraft . . . . . . . . . . . 264.9 Wing span versus empty weight for prototype aircraft . . . . . . . . . . 274.10 Aircraft length versus empty weight for prototype aircraft . . . . . . . . 284.11 Aircraft height versus empty weight for prototype aircraft . . . . . . . . 284.12 Aircraft power loading versus wing loading for the prototype aircraft . . 294.13 Matching diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.14 Met area and design point . . . . . . . . . . . . . . . . . . . . . . . . . . 334.15 Sketch of concept 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 354.16 Sketch of concept 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 364.17 Sketch of concept 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 385.18 Airfoil shapes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 415.19 Comparison of Airfoil Drag Polar . . . . . . . . . . . . . . . . . . . . . . 425.20 Lift coefficient versus angle of attack for the three airfoils . . . . . . . . 435.21 Coefficient of moments for the three airfoils . . . . . . . . . . . . . . . . 445.22 Lift to drag ratio versus Reynolds number for various DAE airfoils. . . . 465.23 Internal wing structure of Zephyrus (Campbell et al. 2009). . . . . . . . 475.24 Internal wing structure of Musculair (Shoberl 1986). . . . . . . . . . . . 485.25 Vertical tail dimensions. . . . . . . . . . . . . . . . . . . . . . . . . . . . 505.26 Drag polar for tail section airfoils. . . . . . . . . . . . . . . . . . . . . . 515.27 Coefficient of moment vs angle of attack for tail airfoils. . . . . . . . . . 525.28 Thrust and propeller rpm vs. propeller radius for the E193 airfoil with

Cl=0.5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 575.29 Optimal propeller chord along the propeller radius. . . . . . . . . . . . . 595.30 Optimal propeller twist along the propeller radius. . . . . . . . . . . . . 595.31 Conceptual sketch of belt-driven power transmission . . . . . . . . . . . 615.32 Conceptual sketch of gear-driven power transmission . . . . . . . . . . . 625.33 Sketch of the Bionic Bat showing the fairing function of the fuselage on

the landing gear (Lloyd 1985). . . . . . . . . . . . . . . . . . . . . . . . . 647.34 Isometric view of the final design . . . . . . . . . . . . . . . . . . . . . . 73[]

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List of Tables

1 Prototype aircraft statistics . . . . . . . . . . . . . . . . . . . . . . . . . 252 Prototype aircraft wing loading, power loading and power index . . . . . 303 Wing planform. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 404 Airfoil thickness and camber characteristics. . . . . . . . . . . . . . . . . 405 Design and Stall characteristics at Re=600000. . . . . . . . . . . . . . . 456 Empennage Statistical analysis. . . . . . . . . . . . . . . . . . . . . . . . 497 Horizontal Tail Planform design. . . . . . . . . . . . . . . . . . . . . . . 498 Vertical Tail Planform design. . . . . . . . . . . . . . . . . . . . . . . . . 509 Vertical Tail Planform design. . . . . . . . . . . . . . . . . . . . . . . . . 5310 Airfoil characteristics of the E193 and FX60-100 (Barnhart et al. 2004,

p. 30). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5511 Maximum propeller efficiency and corresponding thrust for the FX60-100

and E193 airfoils at Cl 0.5 and 0.7. . . . . . . . . . . . . . . . . . . . . . 5812 Wheel diameters of human powered aircraft (RAeS 2009). . . . . . . . 6513 Horizontal distance between landing gears for recumbent human powered

aircraft (RAeS 2009). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6814 Dimensions of the selected main and nose tyre (MBL 2007, p.3, p. 14) . 6815 Empty weight breakdown by component of prototype aircraft . . . . . . 7016 Empty weight breakdown by component of prototype aircraft . . . . . . 71[]

1

1 INTRODUCTION

1 Introduction

Human powered vehicles have been an important part of human history for thousands

of years. Before animals were trained to provide transport and well before the internal

combustion engine was invented, people had to transport themselves under the power

of their own muscles – by walking or paddling a boat or, later, by pulling or pushing

carts or by using machines such as bicycles. In the last century, human powered vehi-

cles have undergone a revolution. Especially in first-world countries, the use of human

powered vehicles as sporting and recreational equipment has led to technological ad-

vancements in the realm of land, water and air vehicles. In particular, the 20th Century

saw the birth of a new form of travel – by air. The potential for an aircraft powered

solely by human muscle was quickly realised however significant developments in light-

weight materials had to be made before the first successful human powered flight could

be made in 1961. Since then several human powered aircrafts have been made and flown.

In this report, a conceptual design for a human powered aircraft is developed based

on a feasibility study, a review of successful human powered aircraft designs and power

and sizing requirements. More detail of the design is then presented so that a prelimi-

nary design including technical drawings can be found.

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2 FEASIBILITY STUDY

2 Feasibility Study

Human powered vehicles are largely the domain of hobbyists and interest groups. Due

to the low power output of a human, there are many limitations associated with human

powered aircraft, particularly range and reliability. This means that there are almost

always more practicable means of transportation, even other human powered vehicle

solutions, which are more attractive in a variety of ways such as cost, speed, inherent

safety and ease of maintenance. Because of this, there has been almost no interest by

industry in human powered flying vehicles. The upside of this is that there is little

competition, while the downside is that specialised production facilities do not exist.

2.1 The Product

The product to be developed is a human powered vehicle capable of winning the Kre-

mer International Marathon Competition prize, with the extension goal, of the vehicle

being either capable or easily adaptable to also attempt to win the Kremer Human

Powered Aircraft Competition for Sport. The Marathon prize will be awarded to the

first human powered vehicle, conforming to a number of rules, able of completing series

of loops around and figure of eight (Section 2.2). The Sport prize will be awarded to

the first vehicle, also complying with certain rules, completing a course consisting of a

0.5km side-length equilateral triangle, under a mean ambient wind speed of 5m/s. The

product is not intended to be sold, and so factors such as market environment and sales

strategy are moot.

There have been previous human powered vehicles that have been successful in flight,

with some winning previous Kremer prizes. It is recommended that these aircraft be

used for statistical analysis and benchmarking.

2.2 Mission Profile

As set out by the Royal Aeronautical Society (1988), in order to win the Kremer Prize,

the aircraft must complete the circuit shown in Figure 2.1. Starting from the course

datum line the aircraft must takeoff, perform two laps of the outer, oval circuit followed

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2 FEASIBILITY STUDY

Figure 2.1: The Kremer Marathon Competition course (Royal Aeronautical Society 1988)

by one lap of the figure-of-eight circuit and then two more outer circuits – all within

one hour. All points on the aircraft must be at least 5 meters above the ground when

the aircraft is passing the course datum line and the turning point markers. The time

taken to land is not included in the time counting but the aircraft must land safely to

win the Prize.

The radius of the turns at either end of the course is allowed to vary between 25

and 75 meters. This means that the entire distance that the aircraft traverses may vary

from 41295.6m to 42867.6m. The historically documented difficulty in controling hu-

man powered aircraft indicates that attempting the course with the maximum possible

turn radius is a better choice, despite the increased range requirements. To complete

the chosen, 42867.6m course within an hour the aircraft must have an average velocity

of 11.9m/s. The design will assume an average velocity of V = 12m/s for surety.

2.3 Competitors

There are no known groups who are currently attempting to complete the Kremer

Marathon course. However, there are other groups known to be attempting to win the

Kremer Sport prize, including Virginia Tech University and their aircraft Iron Butterfly,

and Pennsylvania State University and their aircraft Zephyrus.

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2 FEASIBILITY STUDY

2.4 Regulations and Standards

Due to the low weight of human powered vehicles, they fall into the microlight/ultralight

classes, for which the regulations differ, between countries. In America such craft would

be qualified under the Federal Aviation Administrations (FAA), Federal Aircraft Regu-

lations (FAR), Part 103 (FAR 103), for which no certification is required. In Australia

such aircraft would fall under the Civil Aviation Safety Authoritys (CASA), Civil Avi-

ation Orders (CAO), Part 95, Section 95.10. It should be noted that the Australian

regulations are currently under review, and it is expected that they will be brought

into closer alignment with FAR103. In the United Kingdom, the aircraft falls under the

Civil Aviation Authoritys (CAA) Microlight category. The limitations to the aircraft

according to the above regulations are summarised below.

FAR103:

• If unpowered, weighs less than 155 pounds empty weight

• If powered, weighs less than 254 pounds empty weight

CASA Section 95.10:

• The vehicle must have a takeoff weight of not more than 300kg.

• The aeroplane is required to have a wing loading not greater than 30 kilograms

per square metre at maximum all-up weight

• The vehicle would be required to be registered with Recreational Aviation Aus-

tralia (RAA).

• The vehicle would be required to display a plaque stating ‘Neither CASA nor the

RAA guarantee the airworthiness of the aeroplane’

• The vehicle would be required to display a plaque stating ‘The pilot operates the

aeroplane at the pilots own risk’

Another relevant piece of regulation from Section 95.10 with regards to the competition

flight is that the aeroplane is generally required to fly 500ft above ground level, except

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2 FEASIBILITY STUDY

for when the aircraft is flying “over land that is owned by, or under the control of, the

pilot or of another person (including the Crown) who, or an agent or employee of whom,

has given permission for the flight over the land at such a height; and at a distance of at

least 100 metres horizontally from any person (other than any person associated with

the operation of the aeroplane) and from any public road”.

Alternatively, written application to CASA to fly a craft otherwise than in accordance

with the flight conditions and the application must be made at least 28 days prior to

the flight.

CAA Microlight

• The aeroplane can have a maximum of 2 seats

• A single seat landplane has a Maximum Total Weight Authorised (MTWA) of

300kg

• A two seat landplane has MTWA of 450kg

• Wing loading at MTWA less than 25 kg/m2 or stalling speed at MTWA less than

35 knots (calibrated speed)

Since the aircraft needs to be flown in the United Kingdom to be eligible for the Kre-

mer prize, the CAA Microlight regulation needs to be met. If the aircraft operates in

Australia, the regulations in CASA Section 95.10 must also be met.

2.5 Patents and Intellectual property

Numerous patents for human powered aircraft exist, which includes complete aircraft

which range from a fixed wing type to ornithopters and helicopters. Individual patents

for parts for a human powered aircraft, such as drivetrains also exist. Some of these

patents are International Patents as well as patents for individual countries.

In terms of intellectual property (IP), according to the Australian Patents Act (1990),

the patent holders are granted exclusive rights to the patent and prevent others from

making, using, importing, selling or otherwise exploiting the patented invention (Hood

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2 FEASIBILITY STUDY

& Dibbens 2010). However, the rights that result from the filing of a patent is only

valid for the country in which the patent was filed. To gain IP protection in other

countries, individual patents for those countries must be applied for. Alternatively, a

single international patent may be filed which will give protection in countries part of

the Patent Cooperation Treaty (PCT) (IP Australia 2010).

Since this project is a university course project, and the aircraft will not be manu-

factured, a patent is not necessary. However, if there is any intention for this design to

be realised, then it is recommended that a patent is filed for Australia and any countries

where competitors are present (such as the United States) to protect this design, or a

PCT application may be filed instead. To be eligible for the any Kremer prize, the flight

must be made within the United Kingdom under officially observed conditions (RAeS

1988). Thus, a patent should be filed for the United Kingdom as well.

2.6 Environmental Issues

Due to the vehicles human powered nature, environmental issues are minimal during

the operation of the aircraft. No emissions will be produced, and due to the low speed

of the aircraft the propeller, the noise produced is considered not to be a pollutant.

The only expected environmental issues are within the construction of the aircraft and

any maintenance materials required over its lifetime, as well as eventual decommission-

ing. Where possible, sustainable or recycled materials should be used for the aircraft,

although it is noted that this will be difficult if the use of carbon fibre materials is

extensively used.

2.7 Financial Issues

Although the aircraft has no commercial viability, money is still required for research,

development, manufacture, testing and transportation of the human powered aircraft.

This is an important consideration as without financial support, the project cannot

proceed. For example the Raven human powered aircraft project was shut down before

completion as an additional $300,000 was required for manufacturing of carbon fibre

components (Illian 2001). It is recommended that sponsorship from Universities, or-

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2 FEASIBILITY STUDY

ganisations, private enthusiasts and manufacturers is found to fund the project, and the

budget to be carefully managed.

2.8 Technology Required

Minimal structural weight is desirable for human powered aircraft. Thus, technologi-

cally advanced materials are often used to ensure that the aircraft is its lightest while

still maintaining structural integrity. As a benchmark, materials used on the successful

human powered aircraft, and more modern craft where more modern materials have

been used.

Materials with high strength to weight ratios are desired throughout the aircraft. Car-

bon fibre should be used for the aircraft main frame and spars for the wing, propeller

and control surfaces. Aluminium may also be used in low fatigue areas. Light plastics

is recommended for the skin of the aircraft, and the use of balsa wood or foam is rec-

ommended for the ribs.

Although new and technologically advanced materials can be chosen, the cost, availabil-

ity, ease of manufacture, maintenance and repair of the aircraft components must be

kept in mind. Since the source of income for the project is due to sponsorship, the cost of

materials must be factored into the choice of materials. In previous university projects

of human powered aircraft (for example the Iron Butterfly), many of the components

have been manufactured by the team members, thus, the technical level required to

manufacture the parts must be suited to the skills of the people available. The repair

and detection of manufacturing faults should also be a relatively easy process to ensure

safety of the aircraft.

2.9 Risk Factors

The unique application and design of human powered aircraft leads to other design

issues, namely the manoeuvrability and structural integrity of the aircraft due to the

low weight required.

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2 FEASIBILITY STUDY

A well noted concern of vehicles with very low speeds and relatively long wings is

that when the vehicle attempts to turn, the resulting difference in velocity distribution

across the inside and outside wings tends to create a moment that causes the vehicle

to overturn. Because of this particular care must be taken when designing the vehicles’

control mechanisms and surfaces.

Another concern is due to the relatively low structural mass and large wing areas that

are required to practicably achieve human powered flight that the structure itself will

not be strongly resilient to the loads caused by gusts. To mitigate this initially, the craft

will only be flow in times and places, where very low winds have been predicted, and

only after on site testing of the present wind speeds before the running of the course re-

veals that it is appropriate, potentially while carrying a data logging and measurement

system to check that the structure is behaving as predicted. Subsequent to analysis of

this, flight testing during more appreciably gusty conditions could proceed, potentially

repeatedly, until the limits of the vehicle are determined.

Power system failure is a significant concern for all aircraft. Being of great concern,

the potential failure methods of mechanical power systems, and the resulting reliability

that they ensure is well understood and quantifiable. Using a human as the power

source introduces variability and uncertainty into the systems performance and relia-

bility. To mitigate this issue, only appropriately trained and fit pilots will be allowed to

operate the vehicle, and they shall be under strict instruction to land the vehicle before

the onset of terminal power system failure due to fatigue or other issues.

With the above mentioned risks, there is a danger to the pilot and people in the vicinity

of the aircraft flight. Even though the aircraft is flown at a very low altitude, serious

injury can result if the aircraft becomes unstable or uncontrollable. Thus, it is recom-

mended that the pilot should wear protective apparel, such as a helmet.

2.10 Aircraft type

There are three main types of aircraft that have been used in attempt for human powered

flight: fixed wing, flapping wing (ornithopters) and rotary wing (helicopters). Each of

9

2 FEASIBILITY STUDY

these designs has advantages and disadvantages and has had varying degrees of success

in human powered flight. In addition to the benefits and limitations of each design, the

major constraint of the low power output and uncertain flight regime requires a high

level of technology as well as research. Although the Kremer Marathon competition only

allows fixed wing aircraft, the ornithopter and helicopter designs are also investigated

for completeness.

Fixed Wing Fixed wing aircraft are the conventional design for human powered

aircraft. The method of propulsion is through a propeller that is usually powered by

a pedalling motion, and is the most efficient means of propulsion in practice. Many

configurations of the fixed wing aircraft exist and choices include propeller placement,

wing considerations (canard or conventional, high or low wing, biplane or monoplane)

and pilot positioning.

There are many successful fixed wing human powered aircraft. The MIT built Daedalus

88 holds the record for the greatest distance and longest duration flight, which is

115.11km (74 miles) and 3 hours, 54 minutes respectively (FAI 2010). The record

for the fastest speed over a closed circuit is 44.32km/h by the Musculair II (FAI 2010).

Other notable human powered aircraft include the Bionic Bat, Velair, Gossamer Series,

Monarch Series and Airglow.

Ornithopter An ornithopter aircraft consists of wings which undergo a flapping mo-

tion which produces both lift and thrust, unlike a fixed wing aircraft whose wings only

produce lift. Ornithopters have some advantages over fixed wing design, but also have

many limitations. They have a high theoretical efficiency, more lift and greater manoeu-

vrability compared to fixed wing and helicopter designs (Chronister n.d.). However, the

complexity of creating a flapping wing has limited the success of ornithopters. For the

wing to produce the lift and thrust, they must twist in the right location with a spe-

cific magnitude to produce the required angle of attack. This requires the design of the

wing with material properties that allow the wing to twist. The principal shortcoming of

a human powered ornithopter remains to be the physical capabilities of the human body.

10

2 FEASIBILITY STUDY

Several ornithopter manned flights have occurred with limited success, but were not

truly human powered. More recently, in 2006 the UTIAS Ornithopter No.1 achieved

a 14 second flight at about 2 metres above the ground which covered about 1/3 of

a kilometre (Black 2006). The takeoff phase of this flight was assisted by jet power.

Continuing projects of human powered ornithopters may possible in the future.

Helicopter A helicopter has the ability to take off and land vertically, hover and fly

backwards and laterally unlike a fixed wing aircraft. In terms of the Kremer Marathon

course, a vertical takeoff can result in a shorter take off time and hence the aircraft

will have more time (compared to a fixed wing or ornithopter) to complete the rest of

the course. However, there are many limitations with the human powered helicopter

which outweigh the advantages. Compared to a fixed wing aircraft, helicopters require

a higher amount of power to operate for the same gross weight and cruise speed (Leish-

man 2006, p. 47). Additionally, there is a large lack of knowledge in the theory of

human-powered hovering flight. This includes flow around the rotating blade, espe-

cially for counter-rotating designs, the ground effect and its effect on the airfoil and

stability (Naito 1991, pp. 1-8). Structurally speaking, the loading on a helicopter rotor

is high at the rotor tip, in contrast to the wing loading for a fixed wing. This produces

high bending moments at the blade root which the blades structure has to support.

Despite the many challenges of a human powered helicopter, many design attempts

and a number of successful hovers have been achieved. The world record for the longest

human powered helicopter hover is held by the Yuri-1 helicopter that achieved hover

for 19.46 seconds in 1994 at a height of 0.2 m (Sopher 1997, p. 32). To date, there are

no human powered helicopters that have achieved forward flight and turn.

In conclusion, ornithopter and helicopter designs are not feasible due to limitations

in their development. Even if ornithopters and helicopters were allowed for the Kremer

Marathon prize, the required high endurance and altitude is currently not achievable

with ornithopters and helicopters. The success of fixed-wing aircraft for human powered

flight indicates that this is the most feasible design in order to achieve the project ob-

jective. Therefore, it is recommended that only fixed-wing style aircraft are considered

11

2 FEASIBILITY STUDY

for the final design.

2.11 Technical Task

2.11.1 Introduction

There are currently two prizes available for the development of human powered aircraft

through the Royal Aeronautical Society which were sponsored by Henry Kremer. The

project base is formed on the challenge of designing a human powered aircraft capable

of winning the Kremer International Marathon competition. The value of the prize is

£50,000 and will be awarded to the first human powered aircraft to complete a marathon

distance course in less than one hour, under official competition conditions. The con-

cept of human powered flight is an interesting research and development area, however

it does not have a viable commercial market. Funding for the design, build and testing

of the aircraft would therefore require sponsorship by Universities or private enthusiasts.

Whilst this project is focused on designing an aircraft for the Kremer Marathon Prize,

consideration will also be given to the possibility of then modifying the craft to win the

£100,000 Kremer Human Powered Aircraft competition for Sport.

2.11.2 Standard requirements

The aircraft must follow the specific rules of the Kremer prize in order to be eligible

to win the prize. The Kremer guidelines specify that the aircraft be classified by the

Federation Aeronautique Internationale (FAI) Sporting Code General code and part 11

for class I-C human powered aircraft. The Kremer rules require that the pilot of the

human powered aircraft obtain an FAI sporting licence.

Human powered aircraft may be classified as ultra light and therefore are required

to adhere to the safety regulations and standards of FAR 103 for ultra light aircraft

and the CASA 95.10 regulations. As the Kremer prize is based in the UK, the aircraft

should also be designed to meet the requirements of the CAA microlight category.

12

2 FEASIBILITY STUDY

2.11.3 Performance parameters

The performance parameters of the human powered aircraft are determined by the spe-

cific rules of the Kremer International Marathon competition. The competition requires

the aircraft to complete a 26 mile (42 km) marathon course in less than 1 hour. The

course consists of two markers 4051m apart, and a datum line which is set halfway

between the markers and perpendicular to the line joining the two markers. The course

consists of two laps around the markers, followed by a figure eight and two more laps

around the markers. A sketch of the course is shown in Figure 2.1. The recorded

elapsed time starts when the aircraft is stationary and positioned such that the front

of the aircraft is behind the datum line. When timing starts the aircraft must take off

and complete the course. The elapsed time ends when the aircraft crosses the datum

line after the course has been flown. The landing time is not included in the elapsed

time, however a safe landing must be achieved for the flight to be valid.

The minimum height of the aircraft when crossing the datum line and the end markers

is specified to by 5m. In order to complete the turns the flight altitude is determined,

with the only climb parameter being that the aircraft reach this height in the 2km range

from take off to the first turn marker.

The cruise speed is determined by the requirement for the aircraft to complete the

course in less than 1 hour. Including the turns, the cruise speed is found to be 11.9

m/s.

2.11.4 Technical level of the product

The final design is to be within or exceeding the standard of past competition entrants

so that it may win.

The technical difficulty of constructing the aircraft is not of great importance (as the

design will not reach the stage of commercial manufacture) however the design must be

able to be built given available construction techniques and materials.

13

2 FEASIBILITY STUDY

2.11.5 Economical parameters

The aim of the human powered aircraft is to win the monetary prize of £50,000. In

order for the project to be viable, ideally the monetary outlay on the project should

be less than the value of the prize. The main costs involved with designing the human

powered aircraft for the Kremer International Marathon competition include:

• Research and design development

• Materials and manufacture

• Testing

• Competition entry fees

• Minimum insurance required by the Kremer prize rules

• Pilot training and cost – including FAI Sporting licence

2.11.6 Power plant type and requirements

It is required to design an aircraft that can be solely powered by the human body.

This excludes the use of power from any other source (such as fossil fuels, solar power,

geothermal power, wind or water power) but does not exclude the use of batteries

charged by a hand-cranked electricity generator or any other mechanism with poten-

tial energy supplied by the physical exertion of a human. The competition guidelines,

however, requires that no energy is to be stored in the aircraft. Additional energy is

expected to be provided by thermal updrafts but, as this may be countered by any

downdrafts, this is to be neglected.

It is therefore required that the power plant be a human with physical fitness and

body type such that they are capable of producing a mean output of 225W for one

hour. The power produced by the person will be used to rotate a propeller to provide

the aircraft with the required thrust.

2.11.7 Main system parameters requirements

The aircraft’s main system requirements are summarised below:

14

2 FEASIBILITY STUDY

• The human powered aircraft is designed to have fixed landing gear for ease of

control and reduced manufacturing costs.

• As no electrical power is allowed to be stored on board the actuation of the control

surfaces must be mechanical, through the pilots manipulation of a joystick type

handlebar utilising wires and rods attached to the control surfaces

• The vertical and horizontal tails are designed to be all flying.

• Due to the requirement of the Kremer prize for the aircraft to operate with no

electrical power storage, it is therefore not possible for the pilot to have any

onboard avionics to determine flight speed and altitude.

2.11.8 Special systems and miscellaneous requirements

As this is an endurance craft powered by human effort, the pilot/power plant needs to

be supplied with amenities to ensure he is capable of continued exertion. The aircraft

is to include a space to hold enough liquid to hydrate an athlete competing for at least

one hour – approximately 1.5L.

2.11.9 Reliability and maintainability

The aircraft will only be required to fly for an hour at a time between maintenance

sessions. Although the operation time of the aircraft is low, the aircraft should be built

to be reliable and easily maintained.

2.11.10 Unification level

It is anticipated that only one aircraft will be built and used for the competition. The

aircraft should be design to be easily modified should it is desired to attempt the Kremer

Sport prize. To enable modifications to the initial design, the various components of the

aircraft (fuselage, airfoil, drive mechanism, propeller) should be able to be individually

removed and altered.

15

3 LITERATURE REVIEW

3 Literature Review

Firstly, a review of the literature related to human powered aircraft (HPA) will be

undertaken. So that the conceptual designs take into account the specific qualities that

will ensure a capable human powered aircraft, the details of some of the most successful

HPAs will also be given.

The literature relating to human powered flight, and in particular those previously

successful designs of human powered aircraft, has been reviewed. This will include an

examination of the market for which HPA are designed. The purpose of the review is

to uncover those aspects of aircraft design peculiar to the human powered aircraft and

to gain an understanding of the state of the art in HPA design to ensure the conceptual

designs generated are able to be competitive to the level required by the technical task.

3.1 Successful Designs

In this section, previously successful designs of human powered aircraft are introduced.

Those with performance features relevant to the endurance design considered in this

report are expanded on and described in detail.

Attempts to make successful human powered aircraft have been made since early last

century. Many of the early attempts at human powered flight involved catapult or,

in the case of the 1935 HV-1 Mufti, tensioned cables to assist in launch rather than

takeoff solely under the power of a human being. The first officially verified flight

of a strictly human powered aircraft was that of Southampton Universitys SUMPAC

which flew 650m in 1961. After this, various HPAs (such as the Puffin, Liverpuffin and

Jupiter aircraft) were constructed to fly distances of about 1km but the Kremer Prize

of £500,000 for a flight of 1 mile around a figure-of-eight course was won in August 1997

by AeroVironment Inc’s Gossamer Condor. This success was followed by the variant

aircraft Gossamer Albatross which won the second Kremer Prize for crossing the English

Channel under human power in June 1969. The next Kremer Prize to be awarded was

the speed prize which was given to MITs Monarch HPA in May 1984 for completing

a 1.5km triangular course in under 3 minutes. The Monarch included batteries previ-

16

3 LITERATURE REVIEW

ously charged by the action of the pilot as a source of extra energy during its flight.

The same Kremer sports prize was awarded to the non-US Musculair I in 1984. An

aircraft that was not built with the Kremer Prizes in mind but instead a long-range

flight to pay homage to the Greek human powered flight myth of Icarus whos father

Daedelus is said to have built him wings to fly from the island of Crete to the mainland

was MITs Daedelus. It holds the Federation Aeronautique Internationale (FAI) record

for the longest time and distance flown under human power. Next, some of the design

details of these successful human powered aircraft will be discussed.

3.1.1 Gossamer Condor

This aircraft won the first Kremer prize by taking a novel approach to solving the

problem of the very low power to weight ratio possible when using a human to power

an aircraft. The design was not based on a sailplane but instead took many of its

features from the design of hang-gliders. The configuration can be seen in Figure 3.2.

A very large wing area produced more lift at a very slow speed so that the drag penalty

imposed by the wire bracing needed for structural support was negated (Drela 1990, pg

101). The aircraft was controlled by a canard control surface and the pilot, in a reclined

position, powered the pusher propeller. Roll was controlled by wing warping rather

than ailerons which, at the time, would have used too much weight. The wing spars

were made of aluminium and the rest of the aircraft was mostly lightweight plastics.

3.1.2 Gossamer Albatross

This aircraft won the second Kremer prize, for crossing the English channel – 35.82km

in 2 hours and 49 minutes. Like its predecessor, the Gossamer Condor, the Gossamer

Albatross was controlled by a large horizontal canard stabiliser. Its large, high aspect

ratio wings had backward sweep and their external bracing was also used to warp the

wing – giving them the twist required to control the aircraft’s roll. The wing ribs are

made of polystyrene foam, the rest of the aircraft’s frame is made of carbon fibre which

all is covered in a mylar skin.

17

3 LITERATURE REVIEW

Figure 3.2: The Gossamer Condor in flight (Don Monroe 1977)

Figure 3.3: The Gossamer Albatross in flight (Jim Harrison)

18

3 LITERATURE REVIEW

3.1.3 Musculair I & II

Between them, these two aircraft won three Kremer prizes based on their speed and

agility. The aircraft had a conventional empennage with pusher propellor behind the

tail, monoplane, high wing configuration. To achieve the requried maneouverability

the Musculair I had a very small wing area – only 16m2 (Roper 1995, p. 231). The

Musculair I incorporated stored human energy in batteries to help the pilot achieve

the required speeds however the Musculair II (Figure 3.4) was able to improve on the

previous Kremer prize speed by 15%, winning another prize.

Figure 3.4: The Musculair II in flight (Ernst Schoberl 1985)

3.1.4 Daedelus 88

A benchmark marathon HPA was MIT’s Daedelus 88 – shown in flight in Figure 3.5. in

1988 this aircraft was flown entirely on the power of one pilot from the island of Crete

to mainland Greece, a total distance of 74 miles, in 3 hours and 54 minutes. The design

of this aircraft contributed a great deal of knowledge about the flight requriements of

19

3 LITERATURE REVIEW

human powered aircraft. In particular, Drela (1988), designed a new set of airfoil profiles

(the LE and DAE series) for the Daedelus and its prototype (the Light Eagle) specifically

tuned for maximum efficiency at the low speed and so low Reynolds numbers of human

powered flight. These airfoils were incorporated in the Daedelus’ very high aspect ratio

wings made of very light but strong materials: carbon fibre spars with polystyrene ribs

covered in a Mylar skin (McIntyre 1988, p. 2). The high aspect ratio meant that the

wings required wire bracing.The long, slow, straight flight required from the Daedelus

meant that no ailerons were required for its control – all steering was achieved by the

all-moving rudder and elevator, arranged in conventional tail configuration. Bevel gears

were used to transfer the pilot’s pedalling power to the tractor-configuration propeller.

Figure 3.5: The Daedelus 88 in flight (John McIntyre, 1988)

20

4 CONCEPTUAL DESIGN

4 Conceptual Design

Three conceptual designs for a human powered aircraft are developed. These will be

based on probable pilot power, a statistical analysis of the features of successful previous

(prototype) designs as well as theoretical weight analysis and required aircraft sizing

considerations.

4.1 Power Plant

The requirements of the Kremer International Marathon competition is that the aircraft

is to be propelled entirely by human power, with no use of devices for either storage

or supply of additional power. In order to achieve these requirements, the concept of

converting human work to mechanical power must be investigated to determine the

capability of man to produce sufficient power to complete the marathon course over the

one hour time frame.

4.1.1 Available power

The human muscle converts chemical energy obtained from the oxidation of carbohy-

drates and glycogen and the conversion of fatty acids into CO2 and water, into mechani-

cal energy at an efficiency of approximately 20-25% (Wilkie, 1960). This means that for

a total metabolic work of 100 watts, approximately 76 watts go to body heat, leaving 24

watts of usable mechanical power (Bussolari & Nadel, 1989). Figure 4.6 shows that for

longer durations of exercise, the constant power output levels out. It was found that for

the long duration human powered flight of the Daedalus, the maximum power per kg of

human was approximately 3.5 W/kg. Basing the design on a 68kg pilot, and allowing

for a reduction in total output power due to the required manoeuvrability, it was found

through the matching diagram that the design W/P ratio is 0.44 kg/W, resulting in

225 Watts of power produced.

The pedal cycle configuration is a mechanically simple design which can be adjusted for

different durations, and can achieve the best use of kinetic energy of the moving legs.

Whilst power is only obtained through leg movement, the muscle mass of the legs is

21

4 CONCEPTUAL DESIGN

Figure 4.6: Maximum output of mechanical power vs total duration of exercise for various typesof exercise : x- cycling, + rowing, ∆- running uphill, *- cycling and turning a hand crank (thearound a symbol indicates performance by a champion athlete) (Wilkie, 1960)

22

4 CONCEPTUAL DESIGN

large enough to utilise all of the oxygen which can be absorbed therefore making it more

efficient over longer durations than exercise configurations which use both the arms and

legs such as rowing, or cycling and turning a crank shaft.

The rowing configuration involves a sliding seat, and utilises both arms and legs to

create mechanical work. This is shown to be a good method of creating mechanical

work for durations longer than 3 minutes, however it requires the use of a larger muscle

mass for a similar power output to cycling, and therefore has a lower mechanical con-

version efficiency. In a human powered aircraft it is also required that the pilots have

their arms available for controlling the steering of the aircraft around the turns and

for controlling level flight stability. The Musculair projects required that the pilot be

almost completely immobile above the hips to maintain precise control of the aircraft

(Schoberl), therefore the rowing configuration would be ineffective for the required flight

manoeuvrability.

The chosen configuration, therefore, is the pedal cycle for converting human mechanical

power.

4.1.2 Recumbent versus upright pedal cycle

Design of the Marathon Eagle (Bliesner, 1994) has suggested that the upright position

is more efficient for obtaining peak power levels, however the recumbent cycling position

is desirable for minimum aerodynamic drag and the lower aspect ratio is less prone to

losses due to sideslip. The recumbent position also allows for more upper body freedom

of movement to use the flight controls (Bussolari and Nadel, 1989).

Investigations into the on road cycling configurations of human powered vehicles have

compared the standard upright position to the aerodynamic recumbent position. Figure

4.7 indicates that in order to achieve the same on road speed, the recumbent human

powered vehicle required a much lower power/weight output than the conventional cycle

designs, whilst also providing greater back support and therefore increased ride comfort

for longer duration use.

23

4 CONCEPTUAL DESIGN

Figure 4.7: Power requirements for touring bicycle, upright and crouched racing bicycles andVector recumbent human powered vehicle (Hennekam, 1990)

24

4 CONCEPTUAL DESIGN

4.1.3 Final power plant design

The decision was made to utilise the recumbent bicycle configuration for the power

plant of the human powered aircraft. In order for the pilot to maintain better control

and visibility while still having full lower back support and a smaller, more aerodynamic

frontal area, a semi-recumbent bicycle position was chosen.

The pilot is chosen to be a 70kg athlete, with the requirement to deliver 225 Watts

of power to maintain the desired flight speed of 11.9 m/s over the marathon distance

and one hour duration.

4.2 Statistical Analysis

Several HPA’s are treated as prototypes for the aircraft to be designed here. The details

of the prototypes design specifications, including empty weight, maximum speed, wing

loading and span and height and length dimensions, are graphed to create an easy

reference for the likely success of the conceptual designs presented in Section 4.5. The

data is listed in Table 1.

Table 1: Prototype aircraft statistics

AIRCRAFT empty weight (kg) wing area (m2) span (m) length (m) height (m)

Gossamer Condor 1 31.75 74.32 29.25 9.14 5.47

Gossamer Albatross2 32 45.34 29.77 10.36 4.88

SUMPAC3 58.97 27.9 24.4 7.61 –

Iron Butterfly4 97.5 16.7 18.3 6.0 3.96

Raven5 40.8 – 35 9.1 3.35

Daedelus 886 31 30.8 34 7.92 1.83

Musculair 17 28 16.5 22 7.2 2.2

Musculair 27 25 11.7 19.5 – 1.5

Velair 887 37.9 16.4 21.7 – –

Velair 897 30.5 17 23.2 – –

Monarch8 32.66 16.54 18.75 8.4 3.41 Taylor (1978), 2 Moulton, Cowley & Lloyd (1979), 3 Southampton Hall of Aviation (viewed 2010), 4 Emory et al. (2005)5 Illian et al. (2008), 6 McIntyre (1988), 7 Frank (2005), 8 McIntyre (2007).

25

4 CONCEPTUAL DESIGN

4.2.1 Empty weight

The empty weights of the prototype aircrafts range from 25kg to 66 kg with the majority

around 30kg. For the purpose of the initial design sizing, an empty weight of 30kg will

be assumed.

4.2.2 Wing area

The wing areas of the prototype aircraft are shown in Figure 4.8. Note that the more

maneouverable aircraft such as the Musculair 1 and 2 and the Velair 88 have smaller

wing areas (about 17m2) while the endurance aircraft, in particular the Gossamer and

Daedelus models have much higher wing areas. Although this aircraft is to be designed

for the Kremer marathon prize, it will have to be more manoueverable than the long-

distance prototype aircrafts so should have a wing area between the two extremes –

around 20m2.

Figure 4.8: Wing area versus empty weight for prototype aircraft

26

4 CONCEPTUAL DESIGN

4.2.3 Wing span

The prototype aircraft have spans between 15m and 35m so a preliminary estimate of

30m will be taken for the aircraft’s wing span. Note that the Iron Butterfly, in Figure

4.9, has a high weight but low wing span because it is a box-wing design in order to

provide a large wing surface area and so the requried lift.

Figure 4.9: Wing span versus empty weight for prototype aircraft

4.2.4 Dimensions

The lengths and heights of the prototype aircraft are shown in Figures 4.10 and 4.11,

respectively. These will be used as a guide to the overall size of the aircraft for config-

uration purposes.

4.3 Weight Analysis & Sensitivity

The weight of a human powered aircraft will not change during the flight as there is no

fuel to be burned. Competition rules also state that no payload is to be jettisoned from

the aircraft. The design will assume an initial estimate for empty weight or 30kg based

on the prototype weight values and a pilot weight of 70kg – thus a takeoff and landing

27

4 CONCEPTUAL DESIGN

Figure 4.10: Aircraft length versus empty weight for prototype aircraft

Figure 4.11: Aircraft height versus empty weight for prototype aircraft

28

4 CONCEPTUAL DESIGN

weight of 100kg. The takeoff weight of the aircraft is, therefore, directly sensitive to

increases in the weight of any component. For example, the sensitivity to pilot weight

is∂WTO

∂Wpilot= 1 kg/kg. (4.1)

4.4 Sizing Analysis

Because this aircraft will operate very close to ground level, throughout this analysis

the pressure correction factor (σ = p/pref ) is set to 1. Unless otherwise stated, units

are imperial so (W/S) is in lb/ft2 and W/P is in lb/hp.

Statistics were gathered to determine the general range for the wing loading (W/S),

power loading (W/P) and power index (IP ) of a HPA. These are shown in Figure 4.12

and in Table 2 where the power index is calculated

IP = 3

√(W/S)

(W/P ). (4.2)

Figure 4.12: Aircraft power loading versus wing loading for the prototype aircraft

29

4 CONCEPTUAL DESIGN

Table 2: Prototype aircraft wing loading, power loading and power index

AIRCRAFT W/S (lbs/ft2) W/P (lbs/hp) IP (hp1/3/ft2/3)

Gossamer Condor 0.2375 472 0.0795

Gossamer Albatross 0.3904 474 0.0938

Iron Butterfly 1.863 624 0.1440

Daedelus 88 0.5681 702 0.0932

Musculair 1 1.023 484 0.1283

Velair 88 1.153 675 0.1196

Velair 89 1.023 621 0.1181

Bionic Bat 1.019 478 0.1287

Average 0.9097 566 0.1131

Based on these statistics and the airfoil selection made in Section ??, size limits can be

placed on the wing and power loadings of the aircraft.

4.4.1 Size to cruise requirements

The average value for the power index has been found, using statistical analysis, to be

0.1131. Equation 4.2 can then be rearranged to relate the maximum power loading

possible during cruise for a given wing loading.(W

P

)cruise

= I−3P

(W

S

)= 690.4

(W

S

)lb/hp. (4.3)

4.4.2 Size to manoeuverability requirements

So that the aircraft can maintain an average velocity of V = 12m/s, it must be able to

complete the 180◦end turns with sustained turn rate ψ = 0.05 radians per second. This

turn rate can also be expressed in terms of the wing loading,

ψ =g

V (W/S)

√0.5ρV 2CL −

(W

S

)2

, (4.4)

where g = 32.2ft/s2 is gravitational accelleration, CL = 1.057 is the wing’s lift coeffi-

cient and ρ = 0.076lb/ft3 is standard, sea-level density.

30

4 CONCEPTUAL DESIGN

Equation 4.4 can be rearranged to find the lowest allowable wing loading for manoeu-

verabitily. (W

S

)turnrate

=0.5ρV 2CL√ψV/g + 1

= 0.381 lb/ft2. (4.5)

4.4.3 Sizing to climb requirements

The Kremer Prize requires that all of the aircraft must be above 5m altitude at the

course datum line and the turning point markers. During the R = 75m radius turns a

bank angle of

φ = arctan

(V 2

gR

)= 11◦ (4.6)

is required. Assuming a statistically reasonable wingspan of 30m, the centre of the

wings would have to reach an altitude of h = 5 + (30/2)sin(φ) = 7.9m by the turning

point marker after taking off in order to satisfy the Kremer altitude requirements. It

is also assumed that there is a takeoff distance of around 600m – long because of the

friction between ground and aircraft and the low power available to a human. This

makes the climb gradient that the aircraft must therefore have

CGR =7.9

4051/2 − 600= 5.43 × 10−3. (4.7)

The corresponding rate of climb, for average velocity V = 12m/s is, then,

RC = CGR× V = 0.0652 m/s. (4.8)

Climb gradient sizing Using the values for lift coefficient (CL = 1.057) and lift

to drag ratio (L/D = 175) determined by the airfoil selection, the relationship between

wing loading and the corresponding maximum power loading can be calculated. First

the climb gradient parameter is

CGRP =CGR+ (L/D)−1

C1/2L

= 0.0108 (4.9)

and so, assuming a propeller efficiency of about νP = 0.9,(W

P

)CGR

=18.97νP

CGRP × (W/S)1/2= 1575

(W

S

)−1/2

lb/hp. (4.10)

31

4 CONCEPTUAL DESIGN

Rate of climb sizing A similar expression can be found for the maximum power

loading at a given wing loading based on the rate of climb parameter (with appropriate

unit conversion)

RCP =196.85RC

33000= 3.89 × 10−4. (4.11)

So the power loading can be calculated, knowing the airfoil has (C3/2L /CD0)max = 180.8,

as (W

P

)RC

= νP

(RCP +

(W/S)1/2

19(C3/2L /CD0)

)−1

= 2314 +3435

(W/S)1/2lb/hp. (4.12)

4.4.4 Sizing for human power

Shenstone (1960, p. 473) requires that the power loading of a human powered aircraft

be related to its wing loading by(W

P

)Shenstone

= 550 × (C3/2L /CD0)

√ρ

2

(W

S

)−1

= 19384

(W

S

)−1/2

lb/hp. (4.13)

Bussolari and Nadel (1989, p. 9) give the maximum, sustained power output of a human

as 225 Watts. If we assume that the pilot will weigh around 70kg and that the empty

weight of the aircraft is the statistically standard 30kg, this will place a limit on the

minimum power loading of the aircraft.(W

P

)human

=2.205 × (70 + 30)

0.001341 × 225= 730.8 lb/hp. (4.14)

4.4.5 Matching diagram and design point

The sizing equations (4.3, 4.5, 4.10, 4.12, 4.14 and 4.13) are all plotted on the same axes

in Figure 4.13. The region where all requirements are met is below all the curves except

for the horizontal human power line, which it is above, and the vertical manoeuverability

line, which it is to the right of.

Figure 4.14 shows a closer view of the met area and the selected design point. This

point is at the intersection of the cruise requirement line and the human power limit so

is optimum for the majority of the flight. This point has wing and power loadings in

the range suggested by the prototypes in Figure 4.12.

32

4 CONCEPTUAL DESIGN

Figure 4.13: Matching diagram

Figure 4.14: Met area and design point

33

4 CONCEPTUAL DESIGN

This design point has a wing loading of (W/S) = 1.06lb/ft2 = 5.17kg/m2 and a power

loading of (W/P ) = 730.8lb/hp= 0.444kg/W. Therefore, the aircraft with an approxi-

mate takeoff weight of 100kg will require a wing area of S = 19.35m2 and a power of

P = 225W.

4.5 Concept Selection

Three conceptual sketches are presented, each showing design features that may help

the aircraft achieve the desired objective of winning the Kremer marathon prize.

4.5.1 Concept 1

This design, sketched in Figure 4.15, has a high-wing, monoplane design for increased

stability. The pilot is in an upright position to maximise his ability to see the path of

the aircraft and he controls the craft’s movement with an all-moving, vee canard. The

aircraft is propelled by a pusher propeller, mounted at the back of an aerodynamically

curved fuselage to reduce the drag. Because the propeller is behind the wing and canard

it will act with least efficiency as it is working in disturbed flow however the propeller’s

wake is not impinging on the aircraft so is not contributing to drag. The pilot can also

directly power the front, bicycle-style landing gear wheel to aid in gaining speed for

takeoff.

4.5.2 Concept 2

The second conceptual design, shown in Figure 4.16, is also a high-wing monoplane

configuration however its wing has a particularly high aspect ratio to give the required

lift for least induced drag. The empennage is in conventional configuration, behind

the fuselage. The pilot pedals in a reclined position, so that the frontal area of the

fuselage is minimum in order to reduce the form drag of the aircraft, powering a tractor

propeller. Positioning the propeller close to the pedals will reduce the length of the

drivetrain – so reducing its weight and complexity.

34

4 CONCEPTUAL DESIGN

Figure 4.15: Sketch of concept 135

4 CONCEPTUAL DESIGN


36

4 CONCEPTUAL DESIGN

4.5.3 Concept 3

The third conceptual design, sketched in Figure 4.17, has a more agile, biplane config-

uration. The two wings allow a larger surface area (and so lift) for smaller wingspan

however the extra structural bracing will increase the aircraft’s drag. The propeller is

placed behind the conventional tail setup so that the fuselage and wings are flying in

undisturbed air – reducing drag and increasing lift but reducing the effectiveness of the

propeller.

4.5.4 Concept selection

The weight, drag, complexity and potential lift- and power-generating potential of the

three concepts are compared. The biplane configuration, concept 3, will be the most

heavy design because of its extra bracing. The canard control surfaces in concept 1 will

disturb the air upstream of both the fuselage and the propeller – incresing the drag and

reducing the propeller’s efficiency. The high aspect ratio, tractor propeller configuration

of concept 2 is therefore chosen as the best compromise in terms of weight and drag

and it also has the added benefit of less induced drag on the wings.

The conceptual design following is, therefore, to design based on the ‘Configuration

2’ sketch in Figure 4.16.

37

4 CONCEPTUAL DESIGN


38

5 PRELIMINARY DESIGN

5 Preliminary Design

5.1 Wing design

5.1.1 Wing planform design

The aircraft designed for the Kremer marathon prize is required to be optimised for

both endurance and manoeuvrability in order for the aircraft to complete the required

marathon length course. From statistical analysis of previous human powered aircraft

designs, it is evident that these aircraft have been designed for either endurance, in

the case of the Daedalus, or for speed and manoeuvrability such as the Musculair and

Velair designs. It is therefore required that the main planform design parameters in-

cluding wing loading, aspect ratio and taper ratio for the marathon aircraft are selected

to optimise for both range and turning ability.

The wing loading (W/S) is found from the matching diagram in Figure 4.14. The de-

sign point for the human powered aircraft is found to be W/S= 5.17 kg/m2. Based on

a total takeoff weight of 100kg, the wing reference area is therefore found to be 19.35m2.

The aspect ratio of human powered aircraft is required to be high. The advantages

of a high aspect ratio wing is that it has lower induced drag at low air speeds (Roskam

1986, p. 185), and a high lift curve slope which results in greater pilot visibility. The

disadvantage of a high lift curve slope is that it produces a rough ride if there is any

turbulence in the air, therefore requiring that the aircraft is to be flown on a still day in

minimal wind conditions. The high aspect ratio also increases the weight of the aircraft

due to the large span.

The wing is chosen to have a symmetrical linear taper both front and back in order

to simulate an elliptical wing, which produces lower induced drag on the aircraft. For

an aircraft with a large wing span, the taper is important to reduce the weight of the

wings. The taper ratio chosen is λ=0.46.

The wing planform design dimensions are shown in Table 3.

39


Table 3: Wing planform.

Wing span, b 24.9 m

Wing area, S 19.35 m2

Aspect ratio 32

Taper ratio 0.46

Root chord 1.06 m

Tip chord 0.487 m

MAC 0.809 m

Position MAC 5.6 m

5.1.2 Wing airfoil selection

Human powered aircraft airfoils are optimised for low Reynolds number operation, as

they use a high wing and low flight speed. At low Reynolds number flows, the viscous

effects are prominent, which causes increased drag on standard airfoils (Lissaman 1983).

For this reason the development of human powered aircraft has led to the development

of specific airfoils for the low Reynolds number flight. These airfoils include the Lis-

saman 7769, Wortmans FX76- MP series developed for the Velair and Drela’s DAE

series developed for the Daedalus. From the wing planform dimensions, at an airspeed

of 11.9 m/s, the design Reynolds number at the mean aerodynamic chord (MAC) is

591 316. The key parameters which are assessed when selecting an airfoil are design lift

coefficient and wing power factor, stall and moment coefficients.

A comparison was conducted using the output from Javafoil to analyse the lift, drag

and moment coefficients of the FX76-140MP, DAE-21 and Lissaman 7769. The airfoil

shape and characteristics are outlined in Figure 5.18 and Table 4.

Table 4: Airfoil thickness and camber characteristics.

Lissaman 7769 DAE-21 FX76-140MP

Max t/c (%) 11 11.9 14

camber (%) 4.4 6.6 6.7

40


(a) Lissaman 7769

(b) DAE-21

(c) FX76-140MP

Figure 5.18: Airfoil shapes.

5.1.3 Design lift coefficient and power factor

The airfoil is selected such that the most maximum and therefore most efficient L/D

ratio is chosen for the cruise segment of flight. The design point is found by looking

at the drag polar of the airfoil and taking a tangent to the polar curve through the

origin. This point corresponds to the maximum L/D for the airfoil. If the airfoil is well

designed, the drag at this point should be equal to the skin friction drag (Raymer 2006,

p. 45). Figure 5.19 shows the drag polar for the three airfoils. It can be seen that the

DAE-21 displays greater lift than the Lissaman 7769, and a lower drag coefficient than

the FX76-140MP. Another aspect which is looked at for the human powered aircraft is

the power factor which is given by Kogiso and Tsushima (2000) as

C3/2L

CD. (5.1)

The comparison of these values is shown in Table 5. The maximum L/D is higher for

the DAE-21 than for the FX76-140MP, however they display a similar power factor.

41


Figure 5.19: Comparison of Airfoil Drag Polar

42


5.1.4 Stall

Stall of a wing occurs when the airfoil operates at its maximum lift coefficient. The stall

characteristics of the three airfoils are compared in Table 5. The FX76-140MP displays

the highest lift, however it also stalls sooner than the DAE and Lissaman airfoils.

Figure 5.20: Lift coefficient versus angle of attack for the three airfoils

5.1.5 Moment coefficient

For a very high aspect ratio wing, the magnitude of the moment coefficients are critical,

with lower moments equating to reduced induced drag and lower torsional wing loading.

The comparison of moment coefficients are shown in Figure 5.21. It can be seen that

the Lissaman airfoil displays the lowest moment coefficients. The FX76-140MP has the

43


highest moment coefficients.

Figure 5.21: Coefficient of moments for the three airfoils

5.1.6 Airfoil selection

The airfoil performance characteristics are compared and displayed in Table 5. The

DAE-21 and FX76-140MP display much higher power factor and maximum L/D, as

well as higher maximum lift coefficients at stall than the Lissaman 7769.

While the DAE-21 and FX76-140MP have similar power factors, the DAE-21 has lower

coefficients of pitching moment, which is important for a high aspect ratio wing (Drela,

1990), therefore the DAE airfoil series is selected for this design.

44


Table 5: Design and Stall characteristics at Re=600000.

Lissaman 7769 DAE-21 FX76-140MP

Design point

CL 0.842 1.057 1.46

CD 0.008 0.006 0.009

L/Dmax 109.63 175.87 150.67

Power factor 100.6 180.8 182.06

α at L/Dmax 5 3 3

Stall characteristics

CLmax 1.546 1.928 2.178

α stall 15 15 13

Moment coefficient at α = 0

Cm -0.024 -0.124 -0.221

5.1.7 The DAE series

The DAE series of airfoils was developed by Drela for use in the low Reynolds number

flight range of the Daedalus. The Daedalus utilised the thicker DAE-11 airfoil at the

root of the wing, the DAE-21 in the midsection, and the 11.1% thickness DAE-31 at the

tip of the wings. The aim of the variable airfoils was to reduce the thickness to chord

ratio as the Reynolds number became lower toward the wing tip, in order to reduce

the induced drag on the wing. A comparison was made for our aircraft wing to deter-

mine if the airfoil was required to change along the wing span, and if so, the location

of the transition. Figure 5.22 shows the lift to drag ratio of the aircraft vs Reynolds

number for flow over the airfoil. The plot indicates that the DAE-21 has greater L/D

for Reynolds numbers higher than 440000. Given the taper on the wing planform, this

Reynolds number occurs at a distance of 9.6m from the aircraft centre line, with the

DAE-31 performing better for the 2.9m of wing nearer to the tip.

The DAE-31 has approximately 5% improvement in L/D for Reynolds numbers less

than 440000, however the length of the segment which receives this benefit is likely

to be smaller than the ailerons which occur in this region of the wing. For simplicity

of aileron design and manufacture and therefore reduced cost, the DAE-21 airfoil is

selected for use along the entire wing span.

45


Figure 5.22: Lift to drag ratio versus Reynolds number for various DAE airfoils.

46


5.1.8 Wing Structural Considerations

The wing structure of aircraft is usually dependant on strength requirements, however

due to the very high aspect ratio of human powered aircraft, stiffness constraints become

more prevalent (Drela 1990). The load profile on high aspect ratio wings cause deflec-

tions which alter the aerodynamic properties of the wing requiring structural stiffness

criterion which are based on maximum safe speed and load factors. The wing structure

of early human powered aircraft was constructed mostly from balsa wood or aluminium,

the more recent designs have embraced the development of composite materials includ-

ing carbon fibre, plastics and Styrofoam. The wing structure of the Zephyrus is shown

in Figure 5.23. This consists of a tubular spar, whereas the Musculair in Figure 5.24

uses an I beam spar. Campbell et al. (2009) found that the tubular construction has

a lighter structural weight than the I-beam. The tubular spar is designed to take all of

the torsion in the wing, therefore the D shaped tubes formed by the airfoils either side

of the spar do not take any load or torsion, allowing them to be constructed out of this

foam or lightweight fibreglass layers.

Figure 5.23: Internal wing structure of Zephyrus (Campbell et al. 2009).

The ribs are placed perpendicular to the spar and take loads long the wing. From

Roskam part III, the rib spacing in the wing is estimated to be approximately 0.9m.

47


Figure 5.24: Internal wing structure of Musculair (Shoberl 1986).

5.2 Empennage design

5.2.1 Empennage Planform

The tail sections of the aircraft are important for trim, aircraft stability and control.

The control surfaces on the horizontal tail are used to control pitch, and on the ver-

tical tail the rudder controls the yaw of the aircraft. The trim is a lift force which is

generated in the horizontal and vertical tails to balance moments around the centre of

gravity (Raymer 2006, p. 73).

The preliminary design of the horizontal and vertical tail sections is achieved through

statistical analysis of previous human powered aircraft design. The main equations for

the sizing of the tail sections are

SH =VHSC

xH, (5.2)

SV =VV Sb

xV, (5.3)

where SH and SV are horizontal and vertical tail areas respectively, VH and VV are the

horizontal and vertical tail volume coefficients, x is the tail arm from the 25% chord of

the wing to the 25% chord of the horizontal (or vertical) tail, S is the wing reference

area, b is wing span, C is the wing MAC.

48


From statistical analysis of existing human powered aircraft designs, Table 6 gives the

volume coefficients and areas for both the vertical and horizontal tail sections.

Table 6: Empennage Statistical analysis.

S (m2) C (m) b (m) xH (m) VH SH (m2) xV (m2) VV SV (m2)

Daedelus 30.8 0.905 34 5.5 0.52 2.66 6.2 0.01 2.07

Musculair I 16.5 0.75 22 3 0.509 2.1 3.6 0.011 1.14

Velair 88 16.4 0.756 21.7 3.75 0.45 1.5 3.6 0.01 1.05

Monarch 16.54 0.88 18.75 5.4 0.66 1.8 6 0.034 1.78

Zephyrus 11.8 0.61 17.5 4.1 0.5 0.88 4.8 0.02 0.86

5.2.2 Horizontal tail

A suggested range for the volume coefficient of the horizontal tail is between 0.5 and

0.65 (Campbell et al, 2009). By analysing the statistical data the horizontal tail volume

coefficient is chosen to be 0.44. The horizontal tail arm, xH is chosen to be shorter than

the Daedalus for improved manoeuvrability and larger than the Musculair and Velair

sport designs. Therefore xH is chosen to be 3.5m. Table 7 provides the horizontal tail

planform dimensions.

Table 7: Horizontal Tail Planform design.

VH 0.44

xH (m) 3.5

SH (m2) 1.998

mean chord (m) 0.647

taper ratio 0.7

horizontail tail span (m) 3.1

5.2.3 Vertical tail

Campbell et al. (2009) suggest a range for the volume coefficient of the vertical tail

to be between 0.02 and 0.05. By analysing the statistical data the vertical tail volume

49


coefficient is chosen to be 0.02. The vertical tail arm, xV is chosen to be 5.15m. Table

8 provides the vertical tail planform dimensions.

Table 8: Vertical Tail Planform design.

VH 0.02

xH (m) 5.15

SH (m2) 1.80

mean chord (m) 0.893

vertical tail span (m) 1.96

The span of the vertical tail is found to be 1.96m. The tail is shaped as shown in

Figure 5.25 in order to reduce the vertical tail span for the calculated tail area required.

A conventional tail configuration is selected with the vertical tail positioned entirely

above the empennage rod. This is due to the requirement for the angle between the

rear landing gear and the lower most point of the empennage being less than the stall

angle of the aircraft, to ensure the tail does not touch the ground during takeoff. As the

aircraft takes off with a low climb gradient of 0.3 degrees, any interference is unlikely.

Figure 5.25: Vertical tail dimensions.

5.2.4 Tail airfoil selection

The airfoil used for the tail sections is selected by analysing tail airfoils for previous

human powered aircraft designs. The airfoil used for the tail sections should be symmet-

50


rical in order to provide lift in either direction (Roskam pt III 1986, p. 272), and have a

small thickness to chord ratio of approximately 10%. From the tail planform design, the

tail airfoils are required to operate at in the Reynolds number range of 4×105 to 7×105.

The airfoils which were chosen for comparison are the Eppler 182 used on the Velair,

NACA 0009 used on the Zephyrus, NACA 0010 which was used on the MIT Monarch

and FX76-100MP which was designed for the Musculair 1 and 2. The drag polar in

Figure 5.26 shows that the FX76-100MP displays the most desirable aerodynamic L/D

characteristics, with low drag coefficient and high maximum lift coefficient. The FX

series were developed by Wortmann for specific use in human powered aircraft. The

FX76-100MP is a symmetric airfoil with 10% thickness ratio. The moment coefficients

for the NACA airfoils and the FX76-100MP are very similar as shown in Figure 5.27,

therefore the airfoil FX76-100MP is selected for both the vertical and horizontal tail

sections.

Figure 5.26: Drag polar for tail section airfoils.

51


Figure 5.27: Coefficient of moment vs angle of attack for tail airfoils.

5.3 Control surfaces

5.3.1 Design of control surfaces

Aileron The ailerons are control surfaces on the main wings which are used to control

the roll motion of the aircraft during turning. As the marathon aircraft is required to

turn around the end markers, the ailerons need to be sized for efficiency in manoeuvra-

bility. From statistics, the aileron size/wing size ratio is determined from the previous

designs for the speed and manoeuvrability Kremer prizes, as the endurance prize did

not require the aircraft to undertake any turns. By convention the chord of the ailerons

is 25% of the wing chord. The length of the aileron is determined by statistical analysis

as displayed in Table 9. From this it is determined that the aileron length to wing span

ratio is 0.25, therefore the length of the ailerons is 6.2m. The aileron tip is positioned

1.2m from the wing tip.

Elevator and rudder The horizontal and vertical tails are designed to be all

moving, to allow for ‘ease of construction, reduced weight and increased performance’

(Campbell et al. 2009). The all moving control surfaces have been utilised in the past

52


Table 9: Vertical Tail Planform design.

Aircraft Aileron Length/wingspan ratio

Iron Butterfly 0.3

Musculair I 0.23

MIT Monarch 0.18

Zephyrus 0.29

average 0.25

on the Marathon Eagle, Velair 88 and Zephyrus human powered aircraft designs. The

horizontal tail is the elevator and is used to control the pitch movements of the plane,

whereas the rudder is the vertical tail section which controls the yaw motion.

5.3.2 Handling of control surfaces

The control of the control surfaces in a human powered aircraft is very important for

the stability of the aircraft, especially during turning. It is required that the pilot is

able to make precise movements of the ailerons, rudder and elevator whilst pedalling

and maintaining aircraft stability. The Musculair and Marathon eagle both adopted

a three axis control system mounted on the handle bars of the cycle. The ailerons

are controlled by the sideways tilt, rudder by rotation about the vertical axis and the

elevator by rotation of the hand grips. The control surfaces are actuated by cables and

rods connected to the control column (Bliesner 1994).

5.4 Propeller design

As a fixed wing aircraft was chosen for the overall configuration, a propeller is chosen

for thrust generation. The main requirement of the propeller is to produce thrust that

is greater than the drag of the aircraft. Due to the low amount of power available to the

propeller, it should be designed to be as efficient as possible. However, the propeller is

limited by the diameter, weight and manufacturing constraints.

53


5.4.1 Propeller location

The propeller locations that have been used in previous designs are the tractor and

pusher configuration. There are benefits and limitations with these designs from an

aerodynamic and mechanical (drivetrain) point of view.

The tractor configuration allows the aircraft to fly in undisturbed air, so the propeller

efficiency is high, and the drag is lower. However, the rest of the aircraft flies in

turbulence from the propeller wake. The drivetrain for a tractor configuration has the

potential to be mechanically simple as it can be designed to be vertical, which also

minimises weight. A need for fewer components may also increase the efficiency of the

drivetrain.

A propeller in the pusher configuration has a reduced efficiency as it works with the

disturbed airflow from the fuselage, wing and tails. However, the skin friction drag of

the aircraft is reduced as it flies in undisturbed air. Due to the location of the propeller,

this configuration requires larger tails and a larger rear weight. The propeller shaft is

also longer (and therefore heavier) than the tractor propeller shaft. Additionally, there

is a risk of the propeller touching the runway during takeoff as the aircraft noses up.

This is an important consideration as the propeller radius is very close to the fuselage

height. It therefore needs a greater clearance and longer landing gear compared to a

tractor and the additional structure for this may increase the weight and drag of the

aircraft.

A modified pusher design has been used where the propeller is located behind the fuse-

lage but before the tail. This was seen in the Bionic Bat where the propeller shaft was

in line with the tail boom. This seems to be a trade off between the aerodynamic advan-

tages and disadvantages of the tractor and pusher design. The efficiency of the propeller

and skin friction drag of the aircraft will lie between the values of the tractor and pusher

configuration. The tail heavy weight distribution problem with a pusher propeller is

not as pronounce in this design, however the propeller clearance from the ground during

takeoff still needs to be considered. Despite these advantages, the drivetrain for this

aircraft has the potential to be heavier and less efficient than a tractor configuration, as

it needs to cover a larger distance (horizontal to the rear of the fuselage, then vertical)

to have a shorter propeller shaft.

54


The propeller of this aircraft will be a tractor propeller. This will achieve the high-

est propeller and drivetrain efficiency, lower weight and the danger with the propeller

striking the ground during takeoff is eliminated.

5.4.2 Number of propeller blades

Theory states that propeller efficiency increases with blade number. However, this does

not account for interference between the blades as the blade number increases, so two

bladed propellers are more efficient due to the less disturbed flow. Additionally, a two

bladed propeller will achieve minimum weight, which is desirable. These reasons are

probably why all previous successful human powered aircraft have used two bladed

propellers. Therefore, a two bladed propeller will be used for this design as well.

5.4.3 Propeller optimisation

The propeller should be optimised to produce the greatest thrust (and hence efficiency

as they are directionally proportional) for the given input power of 225 watts at the

design flying speed of 11.9m/s. The final propeller design requires the determination of

a radius, rpm, twist and airfoil profile for the design point.

Two potential airfoils are investigated: the E193 and the FX60-100. The E193 air-

foil has been used for the propeller for the Monarch and the Gossamer Albatross. The

FX60-100 was designed specifically for human power and was used on the Velair. Thus,

it can be assumed that the properties of these airfoils at low rotational and forward

speeds are best suited to human powered aircraft. The characteristics of the airfoil is

summarised in Table 10.

Table 10: Airfoil characteristics of the E193 and FX60-100 (Barnhart et al. 2004, p. 30).

Parameter E193 FX60-100

Thickness 0.1023 0.0999

Camber 0.0354 0.0356

LE radius 0.0087 0.0069

TE angle 5.5406 5.2198

55


The minimum induced loss propeller design for human powered vehicles by E. Eugene

Larrabee (Larrabee 1984, p. 9-11) is used to determine these optimum parameters. The

theory has been implemented in a design excel sheet provided by the Royal Aeronautical

Society. This spreadsheet requires the input of propeller radius, blade number, flying

speed, rpm, power and airfoil data (Cl, Cd and Cl0), and give the performance, opti-

mum chord along the radius and propeller twist. Since there are three unknown inputs

(propeller radius, rpm and airfoil data) an iterative approach was taken to determine

the best propeller blade.

The airfoil section data was determined from Javafoil, for a range of Reynolds numbers.

In propellers, the design lift coefficient is usually around 0.5 (Raymer 2006, p. 379) so

this value is initially chosen. The Monarch, Goassamer Albatross and Velair aircraft

had a design lift coefficient of 0.7, and the Daedalus had a lift coefficient of 0.8, so an

optimisation for a lift coeffieicnt of 0.7 was also performed.

For each airfoil, the optimum rpm for a fixed radius was determined, and the corre-

sponding efficiency and thrust was recorded. It was ensured that the appropriate airfoil

section data was used to be similar to the reference Reynolds number (the Reynolds

number at 75% of the radius of the propeller). Figure 5.28 shows the dependence of

thrust and propeller rpm on the propeller radius for the E193 airfoil with Cl=0.5. The

optimum rpm decreases with increasing propeller radius. This trend is also observed

with the FX 60-100 airfoil. It can be seen that the greatest thrust occurs at a radius

of about 3 metres which is impractical for the final design. The thrust produced in-

creases with increasing radius prior to the optimum 3 metre mark, so the largest radius

allowable should be chosen. Therefore, a propeller radius of 1.5 metres was chosen for

analysis for the E193 with Cl=0.7 and for the FX60-100.

For the fixed radius of 1.5m, the airfoil with lift coefficient of 0.7 had better performance

than the airfoil with lift coefficient of 0.5 for both airfoils. The desired lift coefficient

is achieved by placing the blade at the corresponding angle of attack. The E193 airfoil

with Cl=0.7 exhibited the best efficiency as seen in Table 11. Thus, this airfoil is chosen

for the propeller.

The corresponding chord and twist angle is shown in Figures 5.29 and 5.30. The chord

56


Figure 5.28: Thrust and propeller rpm vs. propeller radius for the E193 airfoil with Cl=0.5

57


Table 11: Maximum propeller efficiency and corresponding thrust for the FX60-100 and E193airfoils at Cl 0.5 and 0.7.

FX60-100 E193

Cl=0.5efficiency (%) 0.1023 0.0999thrust (N) 0.0354 0.0356

Cl=0.5efficiency (%) 0.1023 0.0999thrust (N) 0.0354 0.0356

of the propeller is quite narrow, about 9cm at the widest point. This gives a maximum

thickness of about 9mm (the thickness of the E193 airfoil is about 10% of the chord as

shown in Table 10), which is quite small. This seems too narrow to produce the required

thrust so the shape should be verified with a different propeller analysis program. Since

this project is the first iteration of design, this is beyond the scope of the project and

the propeller will be left as is. From a structural point of view, a spar is required to pass

through the centre of the blade for stiffness and torsional rigidity. This is not possible

for the first 0.2m of the radius. However, the inner part of the propeller to about 25%

of the radius has a very small contribution to the thrust (Raymer 2006, p. 251), so

the propeller chord from r=0 to 0.375m can be ignored, and tailored to be suited to

the structural requirements of the blade. In its place, a spinner can be placed over this

area to push the air out to where the propeller is more efficient. In piston engines,

the spinner is not this size due to the need for an air intake for engine cooling. For a

human powered aircraft there is no engine, so the spinner can cover the full 25% of the

propeller radius. However, having a spinner with diameter 0.75m is impractical as the

placement of the propeller at the top of the fuselage, in front of the wing will produce

a lot of drag. The greater efficiency gained by such spinner is outweighed by the drag

it will produce. Hence, a more reasonably sized spinner will be used.

5.4.4 Further Propeller Guidelines

The tip speed of the propeller should be below sonic speed. For the design point of

180 rpm, the tip speed is Vtip = π × n × d/60 = 28.27m/s, which is clearly less than

the speed of sound (340 m/s). Thus, the propeller tip will always be kept below sonic

speed.

58


Figure 5.29: Optimal propeller chord along the propeller radius.

Figure 5.30: Optimal propeller twist along the propeller radius.

59


The helical tip speed of a propeller should be limited to about 700fps to reduce noise.

For the human powered aircraft, Vtip helical =√

(V 2tip + V 2) = 30.68 m/s = 100.6 fps

which is within the specified limits

5.4.5 Propeller design summary

The selected propeller is in a tractor configuration and consists of two blades with a

radius of 1.5 metres and an E193 airfoil section. Its twist and chord is determined

from minimum induced loss propeller theory which states that this propeller has 93.9%

efficiency and produces 17.74N of thrust at 180rpm. The blades are installed at an

angle of attack of 2.5 degrees for a design lift coefficient of 0.7. A spinner will be placed

over the propeller for increased efficiency, without increasing the drag penalty.

5.5 Drive train design

The predalling power of the pilot must be transmitted to rotation of the propeller. In

this section a drive train is designed to provide maximum mechanical efficiency for low-

est weight. The drive train must fit inside the fuselage to reduce drag and not impede

the forwards view of the pilot. First the known power transmission requriements are

described, then possible transmission mechanisms are evaluated. The chosen drive train

design is described in detail.

The pilot is assumed to pedal at a rate of 90rpm and the propeller is designed for

optimal thrust at 180rpm. A speed increase of 2:1 is therefore required. The axes of

rotation of the driving and driven actions are at right angles to each other and offset.

In order to reduce the weight and mechanical complexity of the drive train design, no

changes in speed ratio will be incorporated. This will require a higher torque output

from the pilot at takeoff as they build up the propeller’s speed.

5.5.1 Possible power transmission mechanisms

There are three main methods of transmitting power between rotating bodies: belt

drives, chain drives and gear systems. The benefits and disadvantages of each system

60


are described by Blazewicz (2010).

Belt drives Belt drives can transmit power across the greatest distance and some

angle of twist can be introduced into their motion, allowing power transmission between

non-parallel axes. They typically have a mechanical efficiency of around 97%. A belt

drive driven by the pedals could extend to the top of the fuselage, twisting by 90◦to

an orientation where it can then turn the propeller shaft (as sketched in Figure 5.31).

This design has the advantage of simplicity and, therefore, light weight.

Figure 5.31: Conceptual sketch of belt-driven power transmission

Chain drives Chain drives are used over a shorter distance than belt drives and

must generally transmit power in the same plane – hence, no twisting is possible. Al-

though chain drives typically have a mechanical efficiency greater than 97%, they must

not only be oriented in the same plane, the driven and driving sprockets should not be

vertically aligned. These limitations mean that it is not practical to fit a chain drive to

the pilot’s pedals.

61


Gear drives Gear drives have the highest mechanical efficiency (around 99%) but

driven and driver gear must be touching. This means that to transmit the pilot’s power

from the pedals to the propeller shaft, for example by an intermediate hypoid bevel

gear as shown in Figure 5.32, the intervening space must be bridged by a shaft. A gear

drive will, therefore, be heavier than either a belt or chain drive, as will the fastenings

required to hold the shafts in position.

Figure 5.32: Conceptual sketch of gear-driven power transmission

5.5.2 Drive train selection

A twisted belt drive is chosen to transmit the pilot’s power to the propeller. The lower

weight of a belt drive is considered to be a greater benefit than the slightly higher me-

chanical efficiency of a gear driven system. A flat belt is extended from a driver pulley

of radius R (attached directly to the pedal shaft) and twisted through 90◦to turn the

driven pulley (attached directly to the propeller shaft). In order to achieve the desired

62


speed increase from 90rpm to 180rpm, the driven pulley must have radius 0.5R. From

Funk (1996, Table 31.3) the smallest standard pulley sizes for flat-belt drives are chosen

to meet this radius relation. The propeller shaft pulley is to have a diameter of 80mm

and the driver pulley a diameter of 160mm. To keep the weight of the system down,

wooden pulleys will be used and a mineral-tanned leather belt will be used to maximise

the coefficient of friction (Blazewicz, p. 50).

The pedals will rotate at a radius of 180mm (based on a survey of bicycle pedals)

so the given pilot power of 225W implies a torque of 23.9Nm. A mechanical efficiency

of 97% combined with the pulley ratio of 1:2 means that the propeller shaft will be

turned wtih torque of 11.6Nm. This is sufficient to drive the chosen propeller.

5.6 Landing gear

The landing gear is required to absorb the impact energy during landing, provide ma-

noeuvrability on the ground and to allow the aircraft to achieve the required speed for

takeoff, with minimal weight.

5.6.1 Landing gear type

The landing gear types under consideration are fixed landing gears and retractable land-

ing gears. Fixed landing gears are cheaper, lighter and more reliable than retractable

landing gears, but generate considerable drag. Aircraft with cruise speeds less than

120-140 knots generally have a fixed landing gear, and fairings are then used to reduce

the landing gear drag (Arjomandi 2010, p. 218). The designed human powered aircraft

has a cruise speed of 23 knots so a retractable landing gear is not required. This is the

case for the majority, if not all, operating human powered aircraft. However, in some

human powered aircraft designs, retractable landing gears have been proposed but the

aircraft did not eventually fly. The Marathon Eagle, which was designed for the Kremer

Marathon prize featured a driven retractable landing gear (Bliesner 1994, p. 6). The

Raven human powered aircraft (2001) featured a retractable landing gear due to the

4.5lb drag restriction on the aircraft. Thus, a fixed landing gear will be used for weight

saving, ease of manufacture and reliability.

63


5.6.2 Landing gear arrangement

The landing gear arrangement should be chosen to have the lowest weight, friction and

drag. Since the aircraft is light, a quadricycle and multi-bogey arrangement will not

be considered. A bicycle, taildragger and tricycle landing gear will be considered as

potential landing gear arrangements.

Bicycle gears are have two main wheels, with small outrigger wheels at the wings.

They are suited to aircraft with high lift at low angles of attack, and aircraft with a

narrow fuselage and a wide wing span (Raymer 2006, p. 261).Taildraggers allow for

a greater propeller clearance for tractor propellers and have less drag and weight, but

are inherently unstable. Additionally, this configuration causes the aircraft to tile up,

which impairs the pilots vision of the ground. Tricycle landing gears are commonly used

as they provide good steering and ground stability and good forward visibility.

The most suitable arrangement for a human powered aircraft is a bicycle gear, as the

aircraft is has a narrow fuselage and wide wing span. This configuration is found on the

majority of human powered aircraft. The main wheel is at the rear and is placed close

to the pilot and the smaller wheel is placed at the front of the fuselage. There are some

aircraft which have small wheels at the wingtips in case of a possible wing tip strike

or at the base of the vertical tail in case of tail scrape to protect the structure. Many

aircraft also use the fuselage to partially cover the wheel, acting like a fairing (Figure

5.33).

Figure 5.33: Sketch of the Bionic Bat showing the fairing function of the fuselage on the landinggear (Lloyd 1985).

64


The aircraft will have a bicycle configuration with no ’outrigger wheels on the wings.

It is assumed that the handling of the aircraft is similar to a bicycle due to its narrow

body. For this aircraft, a high wing configuration is used. Additionally, competition

rules state that a wing tip handler can help stabilise the aircraft so wingtip wheels are

not required. Part of the landing gear wheels will be sunken into the fuselage, such that

the fuselage acts as a fairing.

5.6.3 Landing gear sizing

For the conceptual design of the landing gear, it is recommended to use the tyre sizes of

similar designs or a statistical approach (Raymer 2006, p. 266). Limited data exists for

the wheel sizes of human powered aircraft and is summarised in Table 12. The Velair and

Musculair 1 have a pusher configuration, while the others have a pusher configuration

which may explain the larger nose wheel diameter. From this small subset of data, the

main wheel should be 8 to 10 inches (20-25cm) in diameter, and the nose wheel should

be 3 to 4 inches (7.5-10cm) in diameter. The shock absorption and braking functionality

of the main wheel is also considered to determine the wheel sizing.

Table 12: Wheel diameters of human powered aircraft (RAeS 2009).

AIRCRAFT Main wheel Nose wheeldiameter (in) diameter (in)

Daedelus 8 3.5

Monarch B 10 3

Velair 10 8

Musculair I 8 7

Landing gear sizes from aircraft manufacturers are not suited for the application on

the human powered aircraft. Firstly, the rated load and speeds are much greater than

the aircraft, which is a point of overdesign. Secondly, the weight of the smallest tyres

are around 5kg each (Goodyear 2002, p. 22) which is much too heavy for the human

powered aircraft. Thus, another source of wheels is required. The wheels need to be

able to withstand the 100kg load and be lightweight. Therefore, tyres manufactured for

scooters and wheelchairs are deemed suitable for this application.

65


5.6.4 Shock absorption

Shock absorption is required during landing and taxiing. Since the landing weight is

not high (design point is 100kg), the shock absorption can be provided by the tyre

deflection. This reduces the aircraft weight, increases simplicity and reduces the size

of the fuselage. To ensure this choice is reasonable, the deflection of the tyre is during

landing calculated. The deflection is given by

S =v2vert

2gνNgear, (5.4)

where Ngear the vertical deceleration rate and is the shock absorber efficiency. For

a general aviation aircraft, Ngear = 3, and for a tire, ν = 0.47 (Raymer 2006, pp.

275-276).

Assuming that the aircraft lands at a distance of 2025m (the distance between the

course datum line and turning point marker) from an altitude of 5 metres, the vertical

velocity is 0.03m/s. Therefore the deflection of the tyre from shock absorption is

S =0.032

2 × 9.81 × 0.47 × 3= 0.03 mm. (5.5)

It is assumed that the wheel chosen can deflect by this amount during landing.

5.6.5 Braking

The final design of the human powered aircraft should have a braking mechanism.

Although no requirement is set for the landing distance in the Kremer Marathon prize,

braking may be required for landing and during tax. The main wheel diameter should

therefore be large enough to provide a brake that can absorb the braking kinetic energy.

This is given in Raymer (2006, p.271) as

KEbraking =1

2

Wlanding

gV 2stall =

1

2

100

9.814.442 = 100.5 Nm/s. (5.6)

Assuming all this braking is provided by the main wheel, this kinetic energy gives a

wheel diameter range of 7-10 inches (Raymer 2006, p. 271). This lies in the range of

the main wheel dimensions from statistical analysis.

66


5.6.6 Driven Main Wheel

The main wheel of the landing gear may be geared so that the rolling resistance of the

wheels can be overcome. This is seen on the Monarch B where the main wheel was

driven for the first 25 feet of the ground roll (Roper 1995, p. 233). However, a driven

wheel will increase the aircraft weight. A mechanism to transfer the power from the

driven wheel to the propeller shaft should also be incorporated to maximise the power

to the propeller. To determine if a driven gear is required, the rolling resistance of the

wheels is calculated.

The force of rolling resistance is determined by F = CrrNf , where Crr is the rolling

resistance coefficient and Nf is the normal force. The rolling resistance of a bicycle tyre

is 0.0022 to 0.005 (Morrison 2010). Taking the upper value of Crr, and the normal force

equal to the takeoff weight of the aircraft, the rolling resistance is 4.9N, which is less

than the thrust produce by the propeller (17 N). Thus, a driven wheel is not necessary.

5.6.7 Landing Gear Positioning

The positioning of the landing gear is important, particularly in take off and for stability

considerations. For the bicycle landing gear arrangement, the aircraft centre of gravity

should be aft of the midpoint between the two wheels (Raymer 2006, p. 264). Roskam

(1997) states that the aircraft centre of gravity should lie between 50% and 60% of

the distance between the nose and main landing gear throughout the entire mission.

This requires an iterative process as the aircraft centre of gravity is unknown. Thus, a

statistical approach is used to determine a suitable landing gear distance. From Table

13, the mean distance between landing gears is 1.13m thus the distance between the

landing gear is selected to be 1.13 m as the preliminary value.

5.6.8 Final landing gear design

A bicycle landing gear configuration will be used for this aircraft, with a fixed landing

gear that is partially shrouded by the fuselage to reduce drag. The distance between

the main and nose tyre is set to be 1.13m. From tyre catalogues, the nose and main

67


Table 13: Horizontal distance between landing gears for recumbent human powered aircraft(RAeS 2009).

AIRCRAFT Distance between nose andmain landing gear (m)

Daedelus 1.18

Monarch b 1.02

Velair 1.20

Musculair I 1.12

wheel is chosen and is summarised in Table 14.

Table 14: Dimensions of the selected main and nose tyre (MBL 2007, p.3, p. 14)

Parameter Main tyre Nose tyre

Diameter x Width (mm) 200x50 100x22

Axle diameter (mm) 15 8

Axle width (mm) 45 45

Weight (g) 480 220

Rated Load (kg) 75 75

5.7 Fuselage and frame design

Fairings have the potential to greatly reduce the drag developed by the vehicle. Given

the low level and nature of the thrust source, as well as the ambition cruise speed and

endurance required, it is imperative to minimise the drag that the pilot must overcome.

Hence a fairing was developed to enclose the pilot, the frame, and as much of the wheels

as possible. To reduce form drag, the body should be kept long and thin, with a smooth

surface and a shape that delays transition in the boundary layer. Although a thin fair-

ing is desirable, it is necessary for the pilot to be able to fit within the fairing and so it

must have at least its widest part of slightly greater than shoulder width. Investigation

into a symmetrical airfoil of the desired thickness to chord ratio – approximately 0.25

to minimize total (form and friction) drag, revealed that the NACA 0024 would be

suitable as well as relatively simple with regards to fabrication simplification.

As with several previous HPA designs, the fairing around the pilot is to be made of

68


clear plastic such as Mylar to maximise the pilot’s field of view.

The frame was designed to as to be most suitable for transferring loads between the

wheels, the seat, and the top of the frame, while minimizing the space required and

still allowing for the cyclist to fit comfortably, and allowing for sufficient height in the

structure that the propeller would not be in danger of coming into contact with the

ground.

69

6 STABILITY ANALYSIS

6 Stability analysis

6.1 Aircraft Centre of Gravity

The percentage of the total, 30kg empty weight of the aircraft made up by each com-

ponent is to be determined statistically. The standard aircraft weight breakdown es-

timations are not valid for a human powered aircraft as there is no engine or fuel to

be considered. Thus, a statistical analysis of the aircraft weight breakdown of human

powered aircraft was conducted. The weight percentage of each structural component

is given for three aircraft in Table 15. From these an approximately average value for

each component is taken for this design.

Table 15: Empty weight breakdown by component of prototype aircraft

AIRCRAFT Daedelus 886 Puffin2 Velair 893 average

wing (%) 50 62 52 55

fuselage (incl. gear) (%) 25 15.4 28.8 23

propulsion & drivetrain (%) 7 12.1 9.8 10

empennage (%) 3 6.1 4.7 5

miscellaneous (%) 15 4.4 4.7 71 Wilson (1989, p. 5), 2 Wimpenny (1975), 3 Frank (1994).

This aircraft will only operate in one condition, that is with the pilot in his seat –

the structural components being fixed. The centre of gravity of the aircraft, therefore,

can only occur at one position. This is calculated in Table ?? with the weights of the

various components assumed to be the statistically average fraction of the total 30kg

weight. The centre of mass of each component is assumed to be at their geometrical

centre and the miscellaneous components, expected to be the pilot’s water etc, will be

located at the same position as the pilot. These positions are found, or estimated (in

the case of the pilot), from the technical drawings included at the end of this report

and presented in Table 16. The propulsion system weight is assumed to be located at

the medium position between propeller and pedals. All distances are with reference to

the front-most tip of the plane.

70


Table 16: Empty weight breakdown by component of prototype aircraft

Component Weight, Wi (kg) xi (m) Wixi (kg.m) zi (kg) Wizi

wing 16.5 0.81 0.049 0.02 0.0012

fuselage (incl. gear) 6.9 1.74 0.252 -0.455 -0.066

propulsion & drivetrain 3 0.405 0.135 -0.455 -0.152

empennage 1.5 5.135 3.423 0.98 0.653

miscellaneous 2.1 1.21 0.576 -0.455 -0.217

pilot 2.1 1.21 0.019 -0.455 -0.0065

SUM 100 – 4.453 – 0.214

The centre of gravity of the aircraft can then be found to be

xCG =

∑Wixi∑Wi

= 0.0445 m, (6.1)

xCG =

∑Wizi∑Wi

= 0.002 m. (6.2)

In terms of percentage of total aircraft length (7m) the centre of gravity is at

xGC = 0.174 (6.3)

6.2 Aircraft Neutral Point

The aircraft neutral point is calculated using equations given by Roskam (1997) for a

conventional configuration,

xAC =xACwf +

xACH

(CLαH

(a− ∂εH

∂α

)(SHS

))CLαW+F

1+(CLαH

(1− ∂εH

∂α

)(SHS

))CLαW+F

(6.4)

where, for subsonic flight as in this case,

∂εH∂α

=2CLαHπAH

. (6.5)

71


For this aircraft we have already found that

CLαH = 0.11, (6.6)

CLαW+F = 0.1098, (6.7)

xACH = 25.1, (6.8)

xACW+F= 25.1, (6.9)

SH = 1.998 m2, (6.10)

S = 19.35 m2, and (6.11)

AH =b2

SH=

3.12

1.998= 4.8. (6.12)

(6.13)

Therefore∂εH∂α

=2 × 0.11

π × 4.8= 0.015, (6.14)

and so

xAC =0.25 +

0.25(0.11(1−0.015)( 1.99819.35 ))

0.1098

1 +0.11(1−0.015)( 1.998

19.35 )0.1098

=0.27547

1.27183= 0.217. (6.15)

6.3 Static Margin

The static margin of the plane is, therefore,

SM = xAC − xCG = 0.217 − 0.174 = 0.043, (6.16)

i.e. 4.3%. This means that, for the only flight configuration possible, the aircraft is

longditudinally stable. The static margin is quite small, however, so for safety it could

be increased by moving the wings towards the rear of the plane or by seating the pilot

more forwards.

72

7 CONCLUSION

7 Conclusion

A plane has been designed to satisfy the specifications given in the technical task.

Technical drawings of the aircrafts preliminary design are presented at the end of this

report. Specifically a three-view, a detailed view and a manufacturing drawing show

the features discussed and calculated in this report.

Figure 7.34: Isometric view of the final design

73

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