AIRCRAFT DESIGN PROJECT 2009 -...

THE UNIVERSITY OF ADELAIDE

SCHOOL OF MECHANICAL ENGINEERING

AIRCRAFT DESIGN PROJECT 2009

GROUP 5

AUSTRALIAN FIRE-FIGHTING AIRCRAFT

Kevin Chan 1132668

Rachel Harch 1132827

Ian Lomas 1132921

Simon Mitchell 1132439

Carlee Stacey 1132235

Aircraft Design Project 2009 Australian Fire-Fighting Aircraft

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Table of Contents

1 Introduction ......................................................................................................................13

1.1 Background................................................................................................................13

1.2 Aim and Objective......................................................................................................14

2 Feasibility Study.................................................................................................................15

2.1 Literature Review.......................................................................................................15

2.2 Market Evaluation .....................................................................................................16

2.2.1 Prototypes .........................................................................................................17

3 Conceptual Design.............................................................................................................20

3.1 Technical Task............................................................................................................20

3.1.1 Standard Requirements......................................................................................20

3.1.2 Performance Requirements................................................................................21

3.1.3 Technical Level ...................................................................................................32

3.1.4 Economical Parameters......................................................................................32

3.1.5 Main System Requirements................................................................................32

3.1.6 Reliability and Maintainability ............................................................................33

3.1.7 Safety.................................................................................................................34

3.1.8 Unification level .................................................................................................34

3.1.9 Ergonomics ........................................................................................................34

3.1.10 Cabin Design ......................................................................................................34

3.2 Statistical Analysis......................................................................................................34

3.2.1 Empty Weight versus Takeoff Weight.................................................................36

3.2.2 Cruise Speed ......................................................................................................37

3.2.3 Stall Speed .........................................................................................................37

3.2.4 Rate of Climb......................................................................................................37

3.2.5 Cruise Altitude ...................................................................................................37

3.2.6 L/D Estimation ...................................................................................................37

3.3 Mission Profile...........................................................................................................39

3.3.1 Mission Profile Diagram .....................................................................................39

3.3.2 Mission Profile Requirements.............................................................................39

3.4 Weight Estimation .....................................................................................................40


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3.4.1 Technical Task Requirements .............................................................................40

3.4.2 Statistical Analysis Requirements .......................................................................40

3.4.3 Remaining Sizing Requirements..........................................................................40

3.4.4 Fuel Fraction Estimates ......................................................................................41

3.4.5 Takeoff Weight Estimation .................................................................................42

3.5 Sensitivity Analysis.....................................................................................................44

3.6 Aircraft Sizing.............................................................................................................45

3.6.1 Sizing to Stall Speed ...........................................................................................45

3.6.2 Sizing to Takeoff Distance...................................................................................45

3.6.3 Landing Distance Sizing ......................................................................................46

3.6.4 Sizing to Climb Requirements .............................................................................46

3.6.5 Corrected Lift Coefficient ...................................................................................47

3.6.6 Drag Polar Estimate............................................................................................47

3.6.7 Sizing to Cruise Speed Requirements..................................................................48

3.6.8 Matching Diagram and Design Point...................................................................49

3.7 Configuration Selection..............................................................................................50

3.7.1 Concept 1...........................................................................................................51

3.7.2 Concept 2...........................................................................................................52

3.7.3 Concept 3...........................................................................................................53

3.7.4 Concept 4...........................................................................................................54

3.7.5 Concept 5...........................................................................................................55

3.7.6 Design Considerations........................................................................................56

3.7.7 Concept Selection ..............................................................................................56

3.8 Fuselage Design .........................................................................................................57

3.8.1 Cockpit Requirements ........................................................................................57

3.8.2 Overall Design of the Fuselage ...........................................................................58

3.8.3 Visibility Diagram ...............................................................................................60

3.8.4 Fire Retardant Tanks and Distribution System ....................................................61

3.8.5 Fuselage Structure .............................................................................................62

3.9 Propulsion System Design ..........................................................................................63

3.9.1 Propulsion System Type Selection ......................................................................63

3.9.2 Number of Engines and the Power Required per Engine.....................................65


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3.9.3 Propeller Sizing ..................................................................................................67

3.9.4 Propulsion System Integration............................................................................69

3.10 Wing Design...............................................................................................................76

3.10.1 Vertical Position .................................................................................................76

3.10.2 Sweep ................................................................................................................76

3.10.3 Aspect Ratio.......................................................................................................77

3.10.4 Thickness Ratio ..................................................................................................78

3.10.5 Taper Ratio ........................................................................................................78

3.10.6 Twist ..................................................................................................................78

3.10.7 Dihedral .............................................................................................................79

3.10.8 Wing Loading .....................................................................................................79

3.10.9 Wing Longitudinal Location ................................................................................79

3.10.10 Aerofoil Selection ...........................................................................................80

3.10.11 Incidence Angle ..............................................................................................84

3.10.12 Flap Sizing ......................................................................................................84

3.10.13 Aileron Sizing..................................................................................................84

3.10.14 Spoiler Selection.............................................................................................85

3.10.15 Flow Control Devices ......................................................................................85

3.10.16 Wing Tips .......................................................................................................85

3.10.17 Centre of Gravity ............................................................................................86

3.10.18 Structure ........................................................................................................86

3.10.19 Wing Design Summary....................................................................................87

3.11 Empennage Design ....................................................................................................88

3.11.1 Empennage sizing...............................................................................................88

3.11.2 Horizontal Stabiliser Geometry...........................................................................89

3.11.3 Vertical Stabiliser Geometry...............................................................................90

3.11.4 Elevator Sizing and Geometry.............................................................................90

3.11.5 Rudder Sizing and Geometry ..............................................................................92

3.11.6 Stabiliser Aerofoils .............................................................................................93

3.12 Landing Gear Design ..................................................................................................95

3.12.1 Landing gear arrangement .................................................................................95

3.12.2 Landing Gear Sizing Nomenclature .....................................................................96


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3.12.3 Landing Gear Placement Criteria ........................................................................97

3.12.4 Nose Weight Criterion........................................................................................97

3.12.5 Height Criterion..................................................................................................98

3.12.6 Landing Gear Position ........................................................................................99

3.12.7 Nose Weight Criterion......................................................................................101

3.12.8 Height Criterion................................................................................................101

3.12.9 Roll-Over Criterion ...........................................................................................101

3.12.10 Over-Turn Angle Criterion ............................................................................102

3.12.11 Tip-Back Angle Criterion ...............................................................................102

3.12.12 Summary......................................................................................................102

3.12.13 Landing Gear Loads ......................................................................................102

3.12.14 Number, Type and Size of Tyres....................................................................103

3.12.15 Tyre Pressure Calculations............................................................................103

3.12.16 Suspension Method and Requirements ........................................................104

3.12.17 Length and Diameter of Landing Gear Struts ................................................105

3.12.18 Nose-Wheel Steering and Castoring Dimensions ..........................................106

3.12.19 Gear Retraction Geometry ...........................................................................107

3.13 Isometric Views .......................................................................................................108

4 Weight and Balance Analysis ...........................................................................................109

5 Stability Analysis..............................................................................................................111

6 Aerodynamic and Performance Analysis ..........................................................................114

6.1 Aerodynamic Analysis ..............................................................................................114

6.1.1 Zero-Lift Drag Coefficient Calculation ...............................................................114

6.1.2 Required Lift Coefficients in Cruise and Loiter Phases.......................................114

6.1.3 Drag Coefficient in Cruise and Loiter Phases.....................................................115

6.1.4 Lift to Drag Ratio Calculation ............................................................................115

6.2 Final Design Weight Estimate...................................................................................115

6.3 Design Point Analysis ...............................................................................................116

7 Conclusion.......................................................................................................................118

8 References ......................................................................................................................119

Appendix A – Fire-fighting Aircraft Statistical Analysis .............................................................122

Appendix B –Statistical Analysis Relevant Aircraft....................................................................123

Appendix C – Calculated Fuel Fractions ...................................................................................124

Appendix D – Sensitivity Calculations ......................................................................................127

Appendix E – MATLAB Code for Takeoff Weight Estimation and Sensitivity Analysis ................132


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Appendix F - Honeywell TPE331-14GR Specifications...............................................................135

Appendix G – Flap Sizing Data .................................................................................................136

Appendix H – Neutral Point Calculations .................................................................................137

Appendix I – Three View Drawings ..........................................................................................139


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List of Figures

Figure 1 - Coordinate System (NASA 2009) ................................................................................12

Figure 2 - Air Tractor 602 (Airliners.net 2009)............................................................................17

Figure 3 - Air Tractor 802 (Airliners.net 2009)............................................................................18

Figure 4 - Canadair CL-215 (Airliners.net 2009)..........................................................................18

Figure 5 - Canadair CL-415 (Airliners.net 2009)..........................................................................19

Figure 6 - Major Australian Airports (Australian Institute of Criminology Website 2004)............23

Figure 7 - Fire Danger Seasons (Australian CSIRO Website 2009) ...............................................23

Figure 8 - Map of the Population Distribution in Australia .........................................................24

Figure 9 - Map of land usage in Australia overlayed with areas covered by the aircraft located at

the selected bases. The solid circles indicate most likely bases, and the dashed circles indicate

other possible aircraft bases (Modified from Australian Natural Resources Atlas Website 2008).

.................................................................................................................................................24

Figure 10 - Probability for the Success of a First Attack Success (Plucinski, Gould, McCarthy,

Holis, 2007)...............................................................................................................................25

Figure 11 - Probability for the Success of a First Attack Success (Plucinski et al. 2007) ...............26

Figure 12 – Figure showing the regions within Australia which can be reached by the fire-fighting

aircraft within different response times (Modified from the Australian Natural Resources Atlas

Website 2008)...........................................................................................................................28

Figure 13 – Figure showing the response time of the fire-fighting aircraft overlayed onto a

population density map (Modified from the Department of Environmental, Water, Heritage and

the Arts Website 2001) .............................................................................................................29

Figure 14 - Australian Runway Lengths ......................................................................................31

Figure 15 - Graph of Takeoff Weight versus Empty Weight for Statistically Analysed Aircraft.....36

Figure 16 - Mission Profile .........................................................................................................39

Figure 17 - Takeoff and Empty Weight Estimate ........................................................................43

Figure 18 - Matching Diagram with Met Area and Design Point Marked ....................................49

Figure 19 - Concept 1 Sketch .....................................................................................................51





Figure 24 - Cockpit Dimensions .................................................................................................58

Figure 25 - Fuselage Sketch .......................................................................................................58

Figure 26 - Front View of Fuselage Sketches ..............................................................................59

Figure 27 – Visibility Diagram ....................................................................................................61

Figure 28 - Tank Location in the Fuselage ..................................................................................61

Figure 29 – Engine Selection: Single Engine versus Twin Engine.................................................66

Figure 30 - Propeller Engine Configurations: Tractor and Pusher (Raymer 2006 p.252) ..............71


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Figure 31 - Engine Mounting Locations: Fuselage, Wings, Tail or as Upper Fuselage Pod...........72

Figure 32 - Honeywell TPE331-14GR Geometry (all dimension in inches) (Honeywell 2006).......72

Figure 33 - Cooling System Configuration (Raymer 2006, p.256)................................................73

Figure 34 - Empennage Configurations (Raymer 2006) ..............................................................88

Figure 35 - Horizontal Stabiliser Arrangement ...........................................................................89

Figure 36 - Vertical Stabiliser Arrangement ...............................................................................90

Figure 37 - Elevator Geometry...................................................................................................91

Figure 38 - Elevator Trim Tab Geometry ....................................................................................92

Figure 39 - Rudder Geometry ....................................................................................................93

Figure 40 - Rudder Trim Tab Geometry......................................................................................93

Figure 41 – Landing Gear Configurations (Raymer 2006) ...........................................................95

Figure 42 - Landing Gear Nomenclature (Roskam 2006) ............................................................97

Figure 43 - Over-turn Angle Criterion (Raymer 2006 p. 232) ......................................................98

Figure 44 - Figure Describing Over-turn Criterion.......................................................................99

Figure 45 - Figure Showing Trail and Rake of the Wheel (Raymer 2006)...................................106

Figure 46 - Sliding Bar Linkage (Raymer 2006) .........................................................................107

Figure 47 - Centre of Gravity Envelope ....................................................................................110

Figure 48 - Longitudinal X-plot for the Operational Empty Weight Configuration.....................112

Figure 49 - CG Envelope, Neutral Point and Static Margin for Each Flight Configuration...........113

Figure 50 - Weight Estimate for the Final Design .....................................................................116

Figure 51 - Final Matching Diagram. ........................................................................................117


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List of Tables

Table 1 – Summary of Response Times for an Aircraft Cruise Velocity of 375 km/h....................27

Table 2: Payload Drop Types. ....................................................................................................30

Table 3 - Aircraft Operating Conditions......................................................................................32

Table 4 - Mission Profile Summary ............................................................................................39

Table 5 - Parameters Estimated from Prototypes and Literature ...............................................40

Table 6: Estimated Fuel Fractions (Roskam 2005) ......................................................................41

Table 7: Sensitivity Analysis Results ...........................................................................................44

Table 8 - Aircraft Sizing Results..................................................................................................49

Table 9 - Design Considerations.................................................................................................56

Table 10 - Fineness Ratio as Specified by Roskam (2004) ...........................................................59

Table 11 - Comparison of the Fineness Coefficient for the Designed Aircraft Compared with the

Recommended Values as Specified by Roskam (2004)...............................................................60

Table 12 - Recommended Frame and Longeron Spacing, and Frame Depth for a Small

Commercial Aircraft as specified by Arjomandi (2009)...............................................................62

Table 13 - Suggested Engine Models (Jackson 2008)..................................................................67

Table 14 - Statistical Analysis of Relevant Engines (Roskam III 2002)..........................................68

Table 15 - Aerofoil Candidates...................................................................................................82

Table 16 - 2D Aerofoil Comparison Table...................................................................................82

Table 17 - 3D Aerofoil Comparison Table...................................................................................82

Table 18 - Wing Tip Table ..........................................................................................................85

Table 19 - Wing Design Summary ..............................................................................................87

Table 20 - Tyre Selection Table................................................................................................103

Table 21 - Suggested Weight Distribution as Percentages (Eger 1983; Arjomandi 2009) ..........109

Table 22 - Aircraft Weight Breakdown and Centre of Gravity Locations ...................................109

Table 23 - Centre of Gravity Locations for Various Payload and Fuel Configurations ................110

Table 24 - Longitudinal Stability in Each Flight Configuration ...................................................112

Table 25 - Comparison of Assumed and Estimated Lift to Drag Ratios......................................115

Table 26 - Fire-fighting Aircraft Statistical Analysis...................................................................122

Table 27 - Honeywell TPE331-14GR Specifications (Jackson 2008) and (Honeywell TPE331-14

2006) ......................................................................................................................................135

Table 28 - Flap Sizing Table......................................................................................................136

Table 29 - Aileron Sizing Table.................................................................................................136


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Nomenclature

Acronyms

2D Two -Dimensional

3D Three-Dimensional

CAD Computer Aided Design

CASA Civil Aviation Safety Authority

CFS Country Fire Service

FAR Federal Aviation Regulations

MAC Mean Aerodynamic Chord of wing

NACA National Advisory Committee for Aeronautics

NAFS National Aerial Firefighting Centre

NASA National Aeronautics and Space Administration

UIUC University of Illinois at Urbana-Champaign

Symbols

a Speed of sound

Vclimb Climb Velocity

(L/D)aircraft Aircraft (L/D)

(L/D)aerofoil Aerofoil (L/D)

(L/D)max Maximum L/D

(L/D)wing Wing (L/D)

(t/c)wing Wing thickness ratio

(W/S) Wing loading

A exhaust Area of Exhaust

Aintake Area of Intake

ARwing Wing aspect ratio

B The distance between the nose and the main landing gears

bwing Wing span

Cd Aerofoil drag coefficient

CGaft The distance from the nose of the aircraft to the most aft CG

CGfore The distance from the nose of the aircraft to the most forward CG


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CGwing Wing centre of gravity

CL Wing lift coefficient

Cl Aerofoil lift coefficient

CL max Maximum wing lift coefficient

Cl max Maximum aerofoil lift coefficient

CL max, L Maximum wing lift coefficient at landing

CL max, TO Maximum wing lift coefficient at takeoff

CLα Lift-curve slope

Cm Aerofoil pitching moment coefficient

CM Wing pitching moment coefficient

Cwing Wing chord

D Drag

Dfuselage The diameter of the fuselage

DP Propeller Diamater

g Acceleration due to gravity (32.2 slugs/ft3)

Hlg The height of the landing gear (from the ground to the bottom of

the fuselage)

Htail The height of the tail above the bottom of the fuselage

HW The half-width of the main landing gear, i.e. the lateral distance

between a main landing gear and the centre-line of the aircraft

L Lift

l characteristic length

Lfuselage The length of the fuselage [ft]

M Mach number

Ma The distance between the main landing gear and the most aft CG

Mf The distance between the main landing gear and the most forward

CG

Na The distance between the nose landing gear and the most aft CG

Nf The distance between the nose landing gear and the most forward

CG

np Number of propeller blades

Pbl Balde Power Loading

Pmax Maxium power output per engine

S Platform area

Sflapped Flapped surface area

SHP Uninstalled Engine Power

Sref Reference surface area


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Swet Wetting surface area

THP Installed Engine Power

Vcr Velocity at cruise altitude

Vfoam Foam Volume

Vfuel Fuel Volume

Vrot Rotational Speed of Engine

Vtip Propeller Tip Speed

VTO Takeoff velocity

WE Engine weight

WE Installed Installed Engine weight

Wlanding The total weight of the aircraft at landing

WTO Takeoff weight

xnosegear The distance between the nose of the aircraft and the nose landing

gear


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Coordinate System Designation

Figure 1 below shows the coordinate system used throughout this report.

Figure 1 - Coordinate System (NASA 2009)


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1 Introduction

The purpose of this report is to detail the design of an Australian fire-fighting aircraft.

1.1 Background

Bushfires present a significant risk to Australia and its people, land and resources. Recently, 210

people died when the 2009 Victorian bushfires destroyed over 400,000 hectares of land (WA

Today 2009). It is imperative that there be systems in place to suppress and control such

bushfires to minimise the risk to human life. One of the most effective methods of containing a

bushfire is through aerial fire-fighting, which is the use of an aircraft for releasing fire fighting

chemicals onto a fire. Both fixed wing and rotary wing aircraft are capable of aerial fire-fighting,

with possible chemicals including water, foams, gels and fire retardants. The key characteristics

of a fire-fighting aircraft include a high useable payload weight and a high cruise speed.

Several aircraft designs have demonstrated excellent aerial fire fighting effectiveness, including

those specially modified for aerial fire-fighting purposes. For large fires, modified commercial

airliners or military transport aircraft have been used with great success. In the past, Australia

has considered using larger aircraft for aerial fire-fighting, but this has proven to be both

expensive and unnecessary. Small companies contracted by state and commonwealth

governments use modified agricultural aircraft, such as the Air Tractor 802, Air Tractor 602 and

M18 Dromader aircraft, for aerial fire-fighting (Dunn Aviation Australia 2009). Agricultural

aircraft often have poor aerodynamic efficiency, but posses improved manoeuvrable over larger

aircraft.

A market analysis was performed to compare existing fire-fighting, agricultural and twin-engine

regional turboprop aircraft. Different configurations were examined, and the most optimal

aircraft were selected. The aerodynamics, stability and performance of the aircraft were

investigated, before a final design was proposed and documented using CAD models and

engineering drawings. A description of manufacturing, maintenance and through-life support is

beyond the scope of the project.


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1.2 Aim and Objective

The aim of this project is to design an Australian fire-fighting aircraft. A design tailored for

unique Australian conditions would give the aircraft an advantage in performance and mission

effectiveness compared with fire-fighting aircraft currently used in Australia. The project will

focus on the conceptual phase of the design process. The primary purpose of the project is to

teach undergraduate students the aircraft design process.


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2 Feasibility Study

The feasibility study was conducted at the beginning of the project to determine the viability of

the project concept and scope. The feasibility study consists of a literature review of texts

pertaining to aircraft design, and a market evaluation and benchmarking study to investigate

similar aircraft

2.1 Literature Review

The conceptual design of the fire-fighting aircraft required research of current prototypes and

design techniques through a literature review. A comprehensive investigation was carried out,

which yielded a number of useful references, including textbooks, published reports, databases

and websites. These sources will be discussed in the following sections, and include those used

for the design of the aircraft structure, configuration and sizing. During the feasibility study and

statistical analysis, numerous aircraft were referenced for statistical data. Aircraft primarily

designed for aerial fire-fighting did not provide adequate data, so agricultural aircraft were also

considered. Of particular interest were the Air Tractor series of aircraft.

The literature used for the project is based on information and equations contained in a range of

texts pertaining to different aspects of aircraft design. For the general embodiment design,

several textbooks and reference books were used. These were namely the Airplane Design

series (Roskam, 2004) and Aircraft Design: A Conceptual Approach (Raymer, 1992). The Roskam

series provides an incremental approach to the design of an aircraft, which can be adapted to

suit the requirements specific to the fire-fighting aircraft. In contrast, Raymer offers a classical

approach to aircraft design with detailed theory and equations.

Aerofoil selection was aided with the use of the UIUC Aerofoil Coordinate Database (UIUC 2008).

This database provides a considerable selection of aerofoils designed and recommended for

aircraft. In addition, Javafoil aerofoil analysis online software was used to compare and select

the most appropriate and suitable aerofoils for the aircraft. Introduction to Aeronautics: A

Design Perspective (Brandt et al. 2004) was used for stability calculations and determination of


Page 16 of 139

landing gear location. Other references have also been used throughout the project, and are

cited where applicable.

2.2 Market Evaluation

A market evaluation of existing fire-fighting aircraft was undertaken in order to gain knowledge

regarding fire-fighting aircraft. The market evaluation was conducted in parallel with the

literature review, and provided the group with invaluable knowledge regarding fire-fighting

aircraft and valuable benchmarking from which design work could be compared.

Initial investigation focused on fire-fighting aircraft. Properties such as take-off weight, empty

weight, payload capability, cruise speed, range and wing area were determined for over twenty

aircraft that had fire-fighting capabilities. These aircraft included the following:

• Bronco OV-10

• TBM Avenger

• Douglass DC-3

• Grumman F7F-3 Tigercat

• Grumman S2-Tracker

• Grumman CDF S-2A Tracker

• Bombardier Canadair 415

• Bombardier Canadair CL-215

• Consolidated PB4Y-2 Privateer

• Boeing B-17 Flying Fortress

• Alenia C-27J Spartan

• Douglas DC-4

• Fairchild C-119 Boxcar

• Beriev Be-200 Altair

• Shinmaywa US-1A

• P3-Orion

• McDonnell Douglas DC-7

• C-130 Hercules

• JRM Mars

• McDonnell Douglas DC-10-10

• Boeing 747

• Antonov An-2 'Colt'

• ROKS-Aero T-101 Grach


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The capabilities of these aircraft, tabulated in Appendix A, exhibited significant variation. The

investigated aircraft included both amphibious and non-amphibious aircraft, converted jet transport

aircraft and small single piston engine aircraft. The confliction in the data meant that it was not

possible to determine a defining relation between takeoff weight and empty weight. However,

several conclusions could be drawn from this data as outlined below:

� Both amphibious and non-amphibious aircraft are used for fire fighting. Amphibious aircraft

demonstrate great payload capability relative to takeoff weight. However, the design

complexity and limitation of suitable landing locations in Australia meant that the

amphibious aircraft were considered unfavourable.

� Large aircraft with fire-fighting capabilities are often produced as single models. These

appeared to represent heavily modified transport aircraft rather than specially designed

fire-fighting aircraft. Consequently, they exhibit comparatively reduced payload capability

compared to smaller aircraft that are intentionally designed for fire-fighting capacities.

2.2.1 Prototypes

The selection of these prototypes was based on the following:

• Similar physical size to the expected fire-fighting aircraft size

• Similar weight to the expected fire-fighting aircraft size

• Similarity of mission requirements and applications

The Air Tractor 602 is a single engine turboprop agricultural aircraft. It has a maximum takeoff

weight of 12,500 lb and has a payload capacity of 630 gallons (2,380 L). The first flight of the Air

Tractor 602 occurred in 1995, with production currently continuing. (Air Tractor 2009)

Figure 2 - Air Tractor 602 (Airliners.net 2009)


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The Air Tractor 802F is a single engine turboprop aircraft primary designed for fire-fighting

applications. It has a takeoff weight of 1,600lb and a payload capacity of 820 gallons (3,100L). The

Air Tractor 802F is a modified version of the Air Tractor 802 agricultural aircraft. The 802 is the

largest existing agricultural aircraft, and as such, defines the boundaries of agricultural aircraft

design. Both models are popular as they offer high efficiency and similar performance compared

with larger twin-engine aircraft. The first flight of the Air Tractor 802 occurred in 1990, and

production of both the 802 and 802F models is currently continuing. The 802F can also be fitted

with Wakeri Floats to create an amphibious aircraft (Air Tractor 2009).

Figure 3 - Air Tractor 802 (Airliners.net 2009)

The Canadair CL-215 is a twin engine amphibious fire-fighting aircraft. It has a take-off weight from

land of 43,500 lb and a payload capacity of 1,400 gallons (5,455 L). The first flight occurred in 1967

and production ceased in 1998 with 121 aircraft built. The CL-215 has a flying boat configuration,

and hence, offers significant aerodynamic advantages when compared with the Air Tractor 802F

fitted with floats. The CL-215 was designed for Canadian conditions, where large lakes provide still

flat surfaces where rapid water collection can occur. (Airliners.net 2009)

Figure 4 - Canadair CL-215 (Airliners.net 2009)


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The Canadair CL-415 was developed from the CL-215, and first flew in 1993. The CL-415 offers

advantages such as an increased takeoff weight of 43,850 lb and a payload capacity of 1,620 gallon

(6,120 L). Other design improvements include an updated cockpit, improved water release system

and corrosion resistance. The CL-415 has been popular in Canada, France and Italy. However, as the

aircraft refills by scooping water from larger rivers or lakes, it does not meet Australian requirements

(Airliners.net 2009).

Figure 5 - Canadair CL-415 (Airliners.net 2009)


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3 Conceptual Design

The conceptual design process aimed to generate, select and develop the most feasible concepts

that could meet all the design requirements. This process was conducted using a classical approach

involving multiple design iterations. Each iteration led to further development of the concepts until

design decisions were made based on sound knowledge and calculations. The following section

outlines the conceptual design process, from initial configuration design through to planform design,

aerofoil and control surface selection, fuselage sizing and propulsion system selection. The resultant

design is brought together in three view drawings.

3.1 Technical Task

This section outlines design requirements for the aircraft. Requirements due to standards,

performance, technological level, economics, main sub-systems and reliability are used to define the

overall constraints on the aircraft.

3.1.1 Standard Requirements

The aircraft must be compliant with regulations defined by the Civil Aviation Safety Authority (CASA)

and the National Aerial Firefighting Centre (NAFC). CASA regulations outline required performance

parameters, construction, testing and operational procedures. Civil aircraft operating in Australia

must receive CASA certification, and hence, it is necessary that the aircraft satisfies all relevant CASA

requirements. This design will be engineered and constructed in Australia, and hence, must adhere

to the Type Certificate (Australian Manufactured) and be manufactured by a CASA approved

company (CASA 2008). CASA regulations frequently refer to requirements defined by the Federal

Aviation Regulation (FAR). A fire-fighting aircraft will need to satisfy components of Part 25

(Airworthiness Standards: Transport Category Airplanes) and Part 91 (General Operating and Flight

Rules). FAR does not outline specific requirements for fire-fighting aircraft. Consequently, Part 137

(Agricultural Aircraft Operations) will be utilized for additional requirements.

NAFC is an Australian Commonwealth government organisation that coordinates the procurement of

fire-fighting aircraft, and defines the required capabilities of fire-fighting aircraft, specifying the

required payload capabilities and delivery systems.


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3.1.2 Performance Requirements

Aircraft Base Location and Range

The aircraft is being designed to supplement the existing aerial fire-fighting capabilities of Australia.

The location at which the aircraft would potentially be based is an important consideration when

determining the range of the aircraft. Once possible bases are identified, the range can be

determined by identifying distances that the aircraft would be required to travel to the site of a fire.

The fire-fighting aircraft being designed will be larger than the existing aircraft currently used by

Australia, which will enable a greater amount of fire retardant to be released upon arrival. To enable

a more economical usage of these aerial fire-fighting resources, it is intended that these aircraft will

operate out of major Australian airports, where existing maintenance facilities and personnel can be

utilised. By centralising the fleet, it is hoped that placing fire-fighting aircraft on standby during

extreme fire hazard days will be more easily accommodated.

Operational costs of the aircraft will also be significantly less, and allows for the set up of specialised

facilities to assist with the loading, maintenance of the aircraft, and to reduce the number of aircraft

(and, because of this, the cost) of placing aerial fire-fighting aircraft on standby. Although aerial fire-

fighting aircraft cannot be used in populated areas due to the hazard of the falling fire retardant,

generally, populated areas are the most central location about which regional areas, where fixed

wing aerial aircraft are most effective, are located.

By examining Figure 7, Figure 8, and Figure 9, it is likely that aircraft would need be based out of, or

nearby, the following airport:

• Perth

• Adelaide

• Melbourne (Tullamarine)

• Mildura

• Sydney (KSA)

• Canberra (The region surrounding Canberra could be covered by aircraft based out of

Melbourne and Sydney. Due to political reasons and public perception, it is likely that an

aircraft would be based at Canberra regardless).

• Tamworth


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• Hobart (Unlikely to warrant its own aircraft due to the climate. If the range of the aircraft is

sufficient, Tasmania could be covered by an aircraft based out of Melbourne.)

• Mackay (Unlikely, as the population density near Cairns is small. This would not warrant a

first attack aircraft. Since the fire season for the north of Australia is during winter, it is

possible to locate the aircraft stationed in the southern regions during summer and in the

northern regions during winter).

Examining the fire danger seasons from Figure 7, the largest number of populated regions within

Australia are exposed to the fire danger seasons during summer. To enure sufficient coverage of all

fire danger areas, the following minimum aircraft bases are recommended to provide sufficient

coverage throughout the summer:

• Perth

• Adelaide

• Melbourne

• Sydney

• Tamworth

It is also recommended that aircraft be stationed at the locations listed below for additional

coverage, faster response times to all areas, and to ensure that there is a degree of contingency

should aircraft from one location be unable to be deployed to a nearby fire:

• Canberra

• Hobart

• Mildura

During other seasons, it would be possible to relocate aircraft from the above bases to other

locations. Using Figure 8 and Figure 9, the distances between these bases, and the potential regions

requiring aerial fire-fighting assistance, were determined. The selected range was selected to be a

minimum of 500km (one way), as this provides sufficient coverage of most regions within Australia.

Consequently, the aircraft should be capable of flying in cruise configuration for up to 1000km. The

coverage provided by an aircraft with these capabilities is shown in Figure 8 and Figure 9.


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Figure 6 - Major Australian Airports (Australian Institute of Criminology Website 2004)

Figure 7 - Fire Danger Seasons (Australian CSIRO Website 2009)


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Figure 8 - Map of the Population Distribution in Australia

(Modified from the Department of the Environmental, Water, Heritage and the Arts Website 2001)

Figure 9 - Map of land usage in Australia overlayed with areas covered by the aircraft located at the selected

bases. The solid circles indicate most likely bases, and the dashed circles indicate other possible aircraft

bases (Modified from Australian Natural Resources Atlas Website 2008).


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Desired Response Time and Cruise Velocity

As the aircraft is being designed primarily as a first attack aircraft, the response time has a direct

impact upon the success of the first attack. The sooner the aerial fire-fighting aircraft arrives at the

scene, the higher the probability of the first attack being successful. A ‘successful first attack’ refers

to an occasion where the contribution of a first attack aircraft contributed to controlling the fire.

The desired response time of the fire-fighting aircraft can be determined by considering the

probability of success of a first attack by a fixed wing aircraft. This is shown graphically in Figure 10

and Figure 11. From these figures, it can be seen that the probability of success is greater if the time

to first attack is reduced.

Figure 10 - Probability for the Success of a First Attack Success (Plucinski, Gould, McCarthy, Holis, 2007)


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Figure 11 - Probability for the Success of a First Attack Success (Plucinski et al. 2007)

It can be seen from Figure 11, that as the time increases, the probability of a successful first attack is

reduced. For an immediate first attack (i.e. a time of zero), the probability of success for FFDI (Forest

Fire Danger Index) values <24 (i.e. low-high classification) is greater than 85%. After 2 hours, this

probability of success is still greater than 80%. In this case, although the probability of success is

higher with a faster response time, decreasing the response time less than 2 hours does not greatly

increase the probability of a successful first attack.

For very high and extreme FFDI (>50), the effects of the response time on the probability of success

are more pronounced. For the ideal, zero time to first attack, the probability of a successful first

attack is approximately 50%. After one hour, this has dropped to approximately 45%, and after 2

hours, the probability has dropped to approximately 40%. It is fires on high FFDI days such as those

recently experienced in Victoria, which threaten to cause the most harm to people and property.

Any advantage to assist with the control of these fires would be desirable. As a result, it is desirable

to achieve the fastest response time possible.

To design an economical aircraft to meet Australia’s fire-fighting needs, some compromise is

required. Although it would be desirable to have the first attack aircraft reach every possible

location of a fire within 30 minutes, this is not feasible. It was decided that the aircraft have a


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response time of no less than 2 hours, including the time from when the first call is received to when

the aircraft takes off from the runway.

For the purpose of this report, it will be assumed that the time between receiving notification of the

fire and takeoff is 30 minutes. The aircraft is therefore required to travel a minimum of 500km in 1.5

hours. This requires a cruise velocity of 333.33km/h. The aircraft will therefore be designed with a

375km/h (or 189 knots) cruise speed.

Table 1 – Summary of Response Times for an Aircraft Cruise Velocity of 375 km/h

Response Time Distance of Fire

Approximate

probability of

success FFDI 24

(high)

Approximate

probability of

success FFDI 50

(extreme)

0.5 hours 0 km

1.0 hours 175 km 82% 50%

1.5 hours 350 km

1.9 hours 500 km

2.0 hours 525 km 78% 40%

The above coverage is shown graphically in Figure 12 and Figure 13.


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Figure 12 – Figure showing the regions within Australia which can be reached by the fire-fighting aircraft

within different response times (Modified from the Australian Natural Resources Atlas Website 2008)


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Figure 13 – Figure showing the response time of the fire-fighting aircraft overlayed onto a population

density map (Modified from the Department of Environmental, Water, Heritage and the Arts Website 2001)

Although the aircraft is designed to return to base if necessary, for extended aerial suppression

campaigns, it is intended that the fire retardant is transported to a closer regional airport and the

aircraft can use this as a base to reduce the turnaround time and fuel costs. It is hoped that the

larger payload capacity and faster response time of the fire-fighting aircraft will allow increased

suppression of the fire, and hence, a more effective first attack.

Payload Weight


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Aerial fire-fighting aircraft standards require that fixed wing aircraft drop retardant or water

payloads in an effective zone which is no less than 40 m long and 15-20 m wide, and that no more

than 15% of the release falls outside of this effective zone (NAFC 2004). The standards require a

minimum coverage of 0.2 L/m2. However, coverage up to 4.0 L/m

2 is required to suppress the

heaviest bushfires (Plucinski et al. 2007). Standards also require a leakage loss rate of no more than

15 L/hr. To provide 4 L/m2 coverage to an effective zone of 40m by 20m and assuming a total time

between payload delivery and filling of 140 minutes (20 minutes between filling and takeoff, 100

minutes to target and 20 minutes on target), the volume of fire retardant required is calculated as

follows:

Equation 1: Required Payload Volume

Long-term fire retardants, such as Phos-Chek D-75-R, are up to three times more effective in

containing bushfires than water (Plucinski et al. 2007). The payload of the fire-fighting aircraft can be

assumed to have a similar density to Phos-Chek D-75-R of 1.067 kg/L (USDA Forest Service 2006).

The payload mass is then 3,966 kg, which was rounded up to 4,000 kg as a conservative estimate to

allow for possible density variations. A payload of 4,000 kg of Phos-Chek allows the payload drop

types seen in Table 2. A three-drop configuration may be possible, depending on the payload

delivery system, but is not required by aerial fire-fighting aircraft standards.

Table 2: Payload Drop Types.

Drop type Coverage

One drop 4 L/m2

Two drops 2 L/m2

Four drops 1 L/m2


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Crew Weight

NAFC outlines a pilot weight of 190 lb (86kg), with 15kg of baggage. The aircraft should only provide

accommodation for one crew member. No additional crew members are required to operate the

aircraft. Hence, controlling the aircraft and releasing the fire retardant are both performed by the

pilot.

Takeoff and Landing Capabilities

Due to the mission of the aircraft, it is desirable for the aircraft to be operated from all airports in

Australia. Runway lengths for airports are shown graphically in Figure 14.

Figure 14 - Australian Runway Lengths

The presence of several short personal runways significantly skews the data. Consequently, it was

decided that the aircraft should operate from the upper 75th

percentile of Australian runways. This

suggests a take off and landing length of 4000ft.

Operational Conditions

The operating conditions of fire-fighting aircraft were researched. However, no overriding

documents or guidelines were found. Consequently, the meteorological conditions of the ten worst

bushfires in Australia's history were investigated. From investigation, the extreme of the aircraft

operating conditions were determined.


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Table 3 - Aircraft Operating Conditions

Air Temperature (°C) 46 2009 Victorian Bushfires

Maximum Temperature in Fire (°C) 2000 1983 Ash Wednesday Bushfires

Temperature Recommended by

Building Codes for Bushfire Prone

Areas (°C)

1300 2009 Victorian Bushfires

Wind Speed (km/hr) 120 Mount Lubra Bushfires

Humidity 6% 2009 Victorian Bushfires

Air Pollution

1500 µg of small

particles per cubic

meter

2009 Victorian Bushfires

Speed of Burning Front – Forest

(km/h) 11 Otways Bushfires

Speed of Burning Front –

Grassland (km/h) 22 Otways Bushfires

The above conditions outline an extreme bushfire normally classified as a firestorm. The height of

the fire front can be over 15m (50ft). The formation of Pyrocumulus cloud can lead to serve

turbulence.

3.1.3 Technical Level

The aircraft is designed to replace existing aircraft, and hence, should demonstrate improved

technologies. In particular, increased fuel efficiency, improved materials and better manufacturing

processes are desirable. The cockpit should also benefit from superior instrumentation. It is

intended that this aircraft will be flown by a single pilot with high-level skills and appropriate

certification.

3.1.4 Economical Parameters

The aircraft should be affordable by small companies as well as larger organisations and government

bodies. It is intended that the aircraft should be more affordable than competing aircraft, in initial

purchase cost, running costs and maintenance costs.

3.1.5 Main System Requirements

Propulsion System Requirements

Propulsion requirements are outlined in FAR 25 Subpart E. Particular reference should be made to

Section 25.961 (Fuel System Hot Weather Operation). No specifications regarding engine number or

engine type exist.


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Landing Gear Subsystem

Rural operation requires that the aircraft must be able to operate from paved and unpaved runways.

Amphibious landing capabilities are not required. FAR 25 Section 25.473 requires the following:

• Maximum descent velocity of 10ft/s at the design landing weight

• Maximum descent velocity of 6ft/s at the design takeoff weight

• The coefficient of friction between the tires and the ground should be less than 0.8

Fuselage Requirements

The fuselage design is required to accommodate the fire retardant release system.

Fire Retardant Release System

NAFC specifies the following requirements:

• The fire retardant release system must be able to produce a “full dump” with a minimum

flow rate of 1000 litres per second under typical dumping conditions.

• The system must be capable of dropping fire retardants at rates less than the maximum flow

rate. It is recommended that the system is capable of at least four flow rates. Flow rates of

500 litres per second, 1000 litres per second and 1500 litres per second are recommended.

• The systems must be capable of splitting the load into more than one drop. Systems with

capacity greater than 3000L must be able to drop the load in four parts.

• The system should be well constructed and include appropriate sealing mechanisms to

prevent leakages. During sixty minutes of static ground testing, losses should be less than

two litres. During a twenty minute turnaround, mission losses should be less than five litres.

• The systems should have the capability to inject the water payload with a measured amount

of foam concentrate.

3.1.6 Reliability and Maintainability

NAFC recommends the following:

• Systems should be simple, robust and reliable

• Systems should have an appropriate level of redundancy. In the event of partial equipment

failure, it must be possible to continue the firebombing mission.

• The use of specialised parts should be avoided

• The aircraft should be field maintainable


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3.1.7 Safety

FAR 91 Section 91.107 states the requirements of one shoulder safety belt as a minimum

requirement for all aircraft. FAR Part 137 requires that agricultural aircraft be fitted with a bird proof

windshield, wire cutters and wire deflectors due to their low altitude operation. The criteria will also

be applied to the aircraft.

3.1.8 Unification level

The vehicle should incorporate both new and existing design components. Inherited design elements

include the wing and empennage aerofoil, the propulsion system, and the flight deck

instrumentation. New designs will occur for the fuselage and fire retardant release system. Iterative

design of the aircraft aerodynamics and the fire retardant release system will be required to reach

the optimal design solution.

3.1.9 Ergonomics

NAFC recommends that the aircraft should be controllable without excessive strength or movement

by the pilot. In particular, fire retardant release should not result in large pitch movements or

excessive trim changes.

3.1.10 Cabin Design

To achieve high accuracy when releasing the fire retardant, the pilot visibility pattern must be

considered. The cockpit should be designed such that the over-nose angle is a minimum of ten

degrees. The pilot should have over-the-side vision of 35 degrees, with 70 degrees of head

movement. The pilot should have completely unobstructed upward vision angles. The cockpit

windscreen should have a minimum angle of 30 degrees to prevent mirroring effect of sunshine

angles.

3.2 Statistical Analysis

Statistical analysis of relevant data is required to produce the technical diagram and suggest base

parameters for design. The technical task outlined a payload capability of 8,820 lb and a range of

584nm. These definitions were used to determine the relevance of aircraft data. Only aircraft

currently in use were considered.

The statistical analysis was limited by relevant fire-fighting aircraft. Consequently, additional data

points were obtained by using agricultural aircraft and small regional turboprops. The investigated

aircraft included the following:


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� Bombardier Canadair 415 (Fire-fighting Aircraft)

� Bombardier Canadair CL-215 (Fire-fighting Aircraft)

� Air Tractor AT602 (Fire-fighting Aircraft)

� Air Tractor AT802 (Fire-fighting Aircraft)

� PZL-Mielec_M-18_Dromader (Agricultural Aircraft)

� Antonov An-2 (Agricultural Aircraft)

� G-164B Super B Turbine (Agricultural Aircraft)

� Pac Cresco (Agricultural Aircraft)

� CASA C-212 (Regional twin turboprop)

� Saab 340B (Regional twin turboprop)

� Sukhoi Su-80 (Regional twin turboprop)

� Convair CV-240 (Regional twin turboprop)

� Embraer EMB 110 Bandeirante (Regional twin turboprop)

� Embraer EMB 120 Brasilia (Regional twin turboprop)

� Handley Page Jetstream (Regional twin turboprop)

� Grumman G-159 Gulfstream I (Regional twin turboprop)

� CASA C-235 (Regional twin turboprop)

� Antonov An-140 (Regional twin turboprop)

� Dornier 328 (Regional twin turboprop)

Properties that were investigated included:

• Weights (takeoff, empty and payload weights)

• Speed (maximum, cruse and stall speed)

• Rate of climb

• Range

• Ceiling

• Geometrical properties (wing area and wing span)

The full data set for these aircraft can be found in Appendix B.


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3.2.1 Empty Weight versus Takeoff Weight

A technology diagram was created to determine the relationship between takeoff weight and empty

weight. The diagram is shown in Figure 15 below.

Figure 15 - Graph of Takeoff Weight versus Empty Weight for Statistically Analysed Aircraft

Three data sets were used to determine a relationship between takeoff weight and empty weight.

The data sets were chosen to match the desired aircraft demographic as closely as possible.

Sufficient data on fire-fighting aircraft were not available, so data on large agricultural aircraft and

regional twin turbo-prop aircraft were used to supplement the statistical analysis. All aircraft used a

turboprop engine for propulsion, and were all designed within the last thirty years. The relationship

between takeoff weight and empty weight is best described using a logarithmic equation. The outlier

(Bombardier Canadair CL-215) was not considered in the analysis. The following resulting

relationship was used as part of the matching diagram:


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3.2.2 Cruise Speed

The technical task outlines a cruise speed of 375km/h (202 knots). Agricultural aircraft exhibit

substantially lower speeds than that required, whilst regional aircraft exhibit speeds higher than the

design requirement. The difference in trends between the three data sets shows that the statistical

analysis is attempting to define an aircraft that is not simply classified. The aircraft required by the

technical task has the roles of a fire-fighting aircraft, and operates similarly to an agricultural aircraft.

The aircraft is heavier than an agricultural aircraft, and lighter than a twin turboprop aircraft.

3.2.3 Stall Speed

The aforementioned statistical analysis was used to determine an appropriate stall speed. For the

aircraft sized in Section 3.4, the stall speed from the statistical analysis was determined to be 82.5

knots.

3.2.4 Rate of Climb

The rate of climb from the statistical analysis was determined to be 850 ft. This was influenced by

the Air Tractor AT-802F fire-fighting aircraft. As discussed in the technical task, FAR 25 requirements

dictate the minimum rate of climb as 300ft, which is much lower than the rate of climb from the

statistical analysis. The difference is due to the agility and manoeuvrability required in order to fight

fires effectively.

3.2.5 Cruise Altitude

The cruise altitude from the statistical analysis was based on the Air Tractor AT-802F, which was

deemed to have the same altitude requirements for fire fighting. The altitude from prototyping in

the statistical analysis was 14,000ft.

3.2.6 L/D Estimation

Data on L/D statistics are not readily available. For the statistical analysis, the L/D was calculated

from other statistics using the Breguet Range equation. Usage of this equation is likely to be

accurate to within 30%, due to the following assumptions:

• The aircraft is cruising for the entire flight

• The aircraft has a constant L/D at all times

• The aircraft has a constant cruise speed at all times

• The aircraft has a constant fuel consumption at all times


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From these assumptions, the L/D was calculated for each aircraft by using the following formula,

where CD is approximated to be 0.137 for each aircraft:

A mathematical model was made from this data, and the relation is as follows:

For the design weight, the L/D for cruise is 12.7. The L/D for loiter is 0.866(L/Dcruise) (Raymer 2006).

Thus,


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3.3 Mission Profile

The following section outlines the mission profile and its associated requirements.

3.3.1 Mission Profile Diagram

Figure 16 below diagrammatically illustrates the mission profile for the fire-fighting aircraft.

Figure 16 - Mission Profile

3.3.2 Mission Profile Requirements

The phases of the mission profile and associated relevant details are given in Table 4.

Table 4 - Mission Profile Summary

Phase Details

1 Engine start and warm-up

2 Taxi

3 Takeoff

4 Climb Climb to 14 000 ft

5 Cruise 540 km (335.54 sm) at 375 km/h

6 Descent To assumed payload drop altitude of 70 ft

7 Loiter and Payload drop 20 minutes (E=0.33 hrs) at 1.3 Vstall

8 Climb Climb to 14 000 ft

9 Cruise 540 km (335.54 sm) at 375 km/h

10 Descent To sea level

11 Landing, taxi and shut down


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3.4 Weight Estimation

The takeoff weight and empty weight of the fire-fighting aircraft can be estimated from the mission

profile, the requirements of the technical task (Section 3.1) and the results of the statistical analysis

(Section 3.2). The requirements from each of these sections are summarised below.

3.4.1 Technical Task Requirements

The technical task requirements are summarised below:

• Payload: 4000 kg (8818.49 lbs)

• Single pilot and baggage design weight: 86kg + 15kg = 101 kg

• Cruise speed: 375 km/h = 341.7542 ft/s

• Radius: 540 km

• Loiter time for payload drop: 20 minutes

3.4.2 Statistical Analysis Requirements

Parameters that were not specified by the technical task were determined using a statistical analysis.

The values of some parameters were weight dependent. Hence, an iterative process was used to

determine the requirements. The results of the statistical analysis are presented below.

Stall speed, Vstall=82.5 knots = 139.2443 ft/s = 94.9393 sm/h

Cruise altitude, hcr = 14,000 ft

Technology diagram: A = -0.8126 and B = 1.2966

3.4.3 Remaining Sizing Requirements

Several parameters were not defined by the stages above, and were estimated from prototypes and

literature. Values for these parameters and the corresponding prototypes are shown in Table 5.

Table 5 - Parameters Estimated from Prototypes and Literature

Parameter Value Source

Rate of Climb 850 fpm = 14.167 ft/s Air Tractor 802F (Air Tractor 2007)

Propeller Efficiency 0.82 (Raymer 2006)

Cruise Power SFC 0.471 lbs/hp/hr (Honeywell 2009)

Loiter Power SFC 0.571 lbs/hp/hr cp(loiter) = 0.1 + cp(cruise) (Raymer 2006)

Reserve Fuel Fraction 0.06 (Roskam 2005)


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Unusable Fuel Fraction 0.005 (Roskam 2005)

3.4.4 Fuel Fraction Estimates

Fuel fractions for phases 1-4, 6, 8, 10 and 11 were estimated using statistics for agricultural aircraft.

Fuel fractions for phases 5, 7 and 9 were calculated based on mission profile requirements. The

mission fuel fraction was then calculated from the individual phase fuel fractions. The results are

shown in Table 6 and the corresponding calculations in Appendix C. Whilst the start and finish

altitudes for the climb and decent of phases 4 and 10 differ from the altitudes for phases 8 and 6, it

is reasonable to assume that these phases have equivalent base fuel fractions as this difference in

small.

Table 6: Estimated Fuel Fractions (Roskam 2005)

Phase Fuel fraction

Engine Start and Warm-Up (Phase 1)

Taxi (Phase 2)

Takeoff (Phase 3)

Climb (Phase 4)

Cruise (Phase 5)

Descent (Phase 6)

Loiter and Payload Drop (Phase 7)

Climb (Phase 8, Corrected for Payload

Drop) Cruise (Phase 9)

Descent (Phase 10)

Landing, Taxi and Shutdown (Phase 11)

Mission Fuel Fraction

*Indicates a base value that must be corrected for payload drop at a later stage.


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3.4.5 Takeoff Weight Estimation

The takeoff weight of the aircraft is estimated from a takeoff weight component breakdown and the

technology diagram. This is achieved by solving Equation 2 and Equation 3 simultaneously for takeoff

weight.

Equation 2 - Takeoff Weight Component Breakdown

Equation 3 - Technology Diagram Equation for Takeoff and Empty Weight

Fuel weight is calculated as a percentage of takeoff weight, and consists of useable and trapped fuel.

Useable fuel consists of mission fuel and reserve fuel. The technical task stated no specific

requirements for trapped fuel or reserve fuel. Hence, conventional fuel fraction estimates of 0.005

and 0.06 respectively, were used. The fuel weight is calculated in Equation 4.

Equation 4 - Fuel Weight

Substituting Equation 4 into Equation 2 and rearranging for WTO gives Equation 5.

Equation 5 - Empty Weight Equation

Equation 5 and Equation 3 were solved graphically using Figure 17, resulting in a takeoff weight of

19,735.3 lbs and an empty weight of 8,697.9 lbs.


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Figure 17 - Takeoff and Empty Weight Estimate


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3.5 Sensitivity Analysis

A sensitivity analysis provides information about the consequences of changing design parameters

on the aircraft takeoff weight. It is a useful tool for determining which parameters have the greatest

effect on the aircraft design. A sensitivity analysis also provides guidance on where to focus weight

reduction efforts. The sensitivity of takeoff weight was calculated to the following:

• Payload weight

• Crew weight

• Empty weight

• Power specific fuel consumption

• Propeller efficiency

• Lift to drag ratio

• Range

• Endurance

• Loiter speed

• Cruise speed

Sensitivity results are shown in Table 7 and the calculations are shown in Appendix D. Takeoff weight

has the greatest sensitivity to power specific fuel consumption, lift to drag ratio and propeller

efficiency during cruise. A reasonable change in power specific fuel consumption or propeller

efficiency of 0.01 can result in changes in takeoff weight of 29 lbs and 17 lbs respectively, whilst a

change in lift to drag ratio of one results in a 108 lbs change in takeoff weight. Large increases in

mission profile requirements (cruise radius and endurance) will also have a significant effect on the

takeoff weight of the aircraft.

Table 7: Sensitivity Analysis Results

Parameter Takeoff Weight Sensitivity

Payload 1.79 lbs/lbs

Crew 1.79 lbs/lbs

Empty weight 2.94 lbs/lbs

Cruise radius 4.10 lbs/sm

2924 lbs/lbs/hp/hr

during cruise -1680 lbs

(L/D)cruise -108 lbs

Endurance 532 lbs/hr

310 lbs

during loiter -216 lbs


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1.44 lbs/sm/hr

(L/D)loiter -12.1 lbs

3.6 Aircraft Sizing

The aircraft has a takeoff weight of 19,735 lbs and must be sized according to FAR25 requirements.

FAR25 includes requirements for takeoff, landing and climb phases of flight. The technical task

specifies a cruise speed requirement and the statistical analysis provides a reasonable stall speed. A

matching diagram method was used to ensure that all requirements were met simultaneously.

3.6.1 Sizing to Stall Speed

The statistical analysis indicated that a stall speed of 82.5 knots ( 139 ft/s) is appropriate for a fire-

fighting aircraft of this size. Stall speed sizing was required for the clean configuration at cruise

altitude as this was the limiting case due to lower lift coefficients and air density. The aircraft was

sized to the stall speed requirement at cruise altitude using Equation 6.

Equation 6 - Stall Speed Equation.

3.6.2 Sizing to Takeoff Distance

Takeoff distance requirements for FAR25 state that the aircraft must clear a 35 ft obstacle at the end

of its takeoff field length. The technical task requires that the takeoff field length be less than or

equal to 4,000 ft. It is assumed that takeoff occurs at 1.1Vstall, and hence, a lower takeoff lift

coefficient is required as shown in Equation 7.

Equation 7 - Takeoff Lift Coefficient

The FAR25 takeoff parameter, shown in Equation 8, is used in to calculate the relationship between

wing loading and thrust loading as suggested by Roskam (2005). The appropriate conversion, seen in

Equation 9, between thrust and static shaft power can then be made to determine the power

loading. The relationship between wing loading and power loading for takeoff requirements is given

in Equation 10, and assumes takeoff at sea level.


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Equation 8 - FAR25 Takeoff Parameter

Equation 9 - Correction between Thrust and Static Power

Equation 10 - Limiting Relationship between Wing Loading and Power Loading for FAR25 Takeoff

3.6.3 Landing Distance Sizing

FAR25 landing requirement state that the aircraft must clear a 50 ft obstacle at the start of the

landing distance. It is desired that the aircraft be able to land with full payload and fuel. Hence, no

weight correction will be necessary to the wing loading or power loading. Statistical data is used to

size aircraft to FAR25 landing requirements. The approach speed (in knots) is related to the landing

field length by Equation 11.

Equation 11 - FAR25 Relationship between Approach Velocity and Landing Field Length

The stall speed in the landing configuration is given by , which gives the

limiting wing loading for landing in Equation 12.

Equation 12 - Limiting Wing Loading for Landing

3.6.4 Sizing to Climb Requirements

The Air Tractor 802F, the prototype aircraft for this analysis, only requires a single turboprop engine.

This aircraft will be initially sized assuming a single engine. However, if the required power is in

excess of what can be provided by a single engine, the aircraft will be resized for two engines. A


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FAR25 aircraft with a single engine must only be sized to the FAR25.119 (AEO) climb gradient

requirement. The drag polar and corrected lift coefficient must be calculated for the FAR25.119

configuration and requirements.

3.6.5 Corrected Lift Coefficient

FAR25.119 (AEO) required a speed of 1.3VSL, and hence, the corrected lift coefficient is given by

Equation 13.

Equation 13 - Corrected Lift Coefficient for FAR25.119 Requirements

3.6.6 Drag Polar Estimate

The drag polar is estimated from the wetted area ratio , equivalent skin friction coefficient

and the estimated effect of landing gear. The wetted area ratio of a fire-fighting aircraft of this size

will be similar to that of a Cessna Skylane. Hence, is a reasonable assumption

(Raymer 2006). The equivalent skin friction coefficient for this fire-fighting aircraft may be assumed

to be similar to a single engine lift aircraft, and hence, (Raymer 2006). Roskam (2005)

suggests that landing gear add an additional 0.015 – 0.025 to the zero-lift drag coefficient. Assuming

well-designed landing gear with fairings, is a reasonable estimate. It

was also assumed that approach flaps were equivalent to landing flaps with

. The zero-lift drag coefficient for the FAR25.119 (AEO) condition is

calculated in Equation 14. The drag polar is then given by Equation 16, where Oswald’s efficiency

factor was calculated for the clean configuration in Equation 15 and landing flaps were assumed to

reduce Oswald’s efficiency factor by 0.05.

Equation 14 - Zero-Lift Drag Coefficient for the FAR25.119 Configuration

Equation 15 - Oswald's Efficiency Factor for the Clean Configuration


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Equation 16 - FAR25.119 (AEO) Drag Polar

The FAR25.119 climb gradient requirement of 3.2% is met by Equation 18, where the climb gradient

parameter (CGRP) is given by Equation 17. The power loading must be corrected for temperature

and humidity effects. Roskam (2005) suggest that a correction factor of 0.85 is appropriate.

Equation 17 - Climb Gradient Parameter

Equation 18 - FAR25 Climb Gradient Limiting Relationship between Power and Wing Loading

3.6.7 Sizing to Cruise Speed Requirements

Cruise speed sizing for propeller aircraft uses the power index, IP. Roskam (2005, p. 163) suggests

that for a cruise speed of 375 km/h (233.0142 mph), a power index of IP=1.32 is required. The

density at cruise altitude, 0.001546 slugs/ft3, gives a density ratio of

. The limiting relationship between power loading and wing

loading is given by Equation 19. A correction factor of 0.7 was required to convert the cruise power

loading at cruise altitude to a takeoff sea level power loading (Roskam 2005).

Equation 19 - Limiting Relationship between Wing Loading and Power Loading for Cruise Requirements


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3.6.8 Matching Diagram and Design Point

The matching diagram for this aircraft shows the stall, takeoff, landing, climb and cruise

requirements in Figure 18. The highlighted area enclosed by the requirement lines represents the

met area, the area within which any combination of power loading and wing loading meets all

requirements. Vertices of this area represent possible design points. Design points that require a

higher wing loading have greater aerodynamic efficiency and stability during turbulence. Design

points that require a lower wing loading have a lower stall speed, takeoff distance and landing

distance. High power loadings require less power and may result in lighter engines, whilst low power

loadings have better performance. A higher wing loading was selected to maximise the aerodynamic

performance of the aircraft (cruise speed and range) as emphasised by the technical task. The design

point wing loading, power loading, wing area and power required are summarised in Table 8.

Figure 18 - Matching Diagram with Met Area and Design Point Marked

Table 8 - Aircraft Sizing Results.

Wing loading 36.85 lbs/ft2

Power loading 11.74 lbs/hp

Wing area 535.56 lbs/ft2

Power 1681 hp


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3.7 Configuration Selection

Fire-fighting aircraft can be classified by their payload capability, propulsion system and landing

system. Payload capacity for the aircraft was specified by the technical task as 8,820 lb. This

payload is heavier than that carried by agricultural or existing single engine turboprop aircraft.

However, the payload is much less than that carried by twin-engine aircraft. Consequently, both

configurations were investigated.

Common propulsion systems include jet, turboprop, piston or radial engine. Aircraft that use a jet

propulsion system are significantly faster than those powered by radial or piston engines. However,

large aircraft have reduced aerobatic capabilities and are hence, rarely used for fire-fighting aircraft.

Turboprop and piston engines are regularly used for fire-fighting aircraft. Both propulsion methods

are further investigated.

Possible landing configurations include seaplane (water only), amphibious (both water and land) and

normal landing (land only) arrangements. Seaplanes and amphibious aircraft offer significant speed

advantages for water refilling. However, Australia lacks the large still bodies of water required for

the refilling process. Hence, water landing capabilities are not seen as advantages. Furthermore,

both seaplanes and amphibious aircraft have reduced aerodynamic performance.


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3.7.1 Concept 1

The first concept considered was a flying boat configuration, where the fuselage can be used as a

hull so that the aircraft can takeoff and land on water. This configuration allows rapid water

collection. However, Australia lacks large inland bodies of water, which makes this concept

unsuitable. A sketch of concept 1 can be seen in Figure 19.

Figure 19 - Concept 1 Sketch


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3.7.2 Concept 2

The second concept considered was a floatplane configuration, where floats are attached to the

fuselage of the aircraft to allow the aircraft to takeoff and land on water. Australia lacks large inland

bodies of water, which makes this concept unsuitable. A sketch of concept 2 can be seen in Figure

20.



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3.7.3 Concept 3

The third concept considered was a twin-engine aircraft. Two engines increase the reliability of an

aircraft, but the maintenance and running costs are higher than a single engine aircraft. A single

turboprop can produce the required thrust for the aircraft, so a twin-engine aircraft was

disregarded. A sketch of concept 3 can be seen in Figure 21.



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3.7.4 Concept 4

The fourth concept considered was a conventional aircraft with a low wing configuration. Although

most agricultural aircraft have a low wing configuration, the wing location decreases stability and

ground visibility. Hence, a low wing configuration was disregarded. A sketch of concept 4 can be

seen in Figure 22.



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3.7.5 Concept 5

The final concept that was considered by the group was a conventional aircraft with a high wing

configuration. This design has high stability and ground visibility, which are two important

considerations for a fire-fighting aircraft. A sketch of concept 5 can be seen in Figure 23.



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3.7.6 Design Considerations

Table 9 presents the design considerations that were considered in the first step of the aircraft

configuration design.

Table 9 - Design Considerations

Consideration Reasoning

Low aerodynamic efficiency A strong structure is more important than

aerodynamic efficiency

Metallic structure Exposure to high temperatures which can damage

composite materials

Operation in harsh environments Exposure to high temperature, humidity and wind

speeds

High cruise velocity Required to reach the fire quickly

High manoeuvrability Required to avoid obstacles, negotiate undulating

terrain and line up for release of payload

Ability to fly at low altitude Payload is released at low altitude

Retractable landing gear Cruise speed is greater than 150 knots

Single tractor turboprop propulsion

configuration

Ease of maintenance, reduced weight, increased

reliability and reduce cost

Simple wing planform Light weight, and cheap and easy to manufacture

High wing configuration High ground visibility, ease of payload loading,

high lateral stability, good structure

Raised cockpit Increased ground visibility

Long nose Payload placement and engine integration

Conventional empennage configuration Light weight, and cheap and easy to manufacture

3.7.7 Concept Selection

Considering each of the concepts and the design considerations listed in Table 9, Concept 5 was

selected as the most feasible option.


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3.8 Fuselage Design

The purpose of the fuselage is to attach the wings and empennage, as well as the cockpit, motor,

payload, and landing gear. The challenge with designing a fire-fighting aircraft is the requirement for

the payload to be located directly on the centre of gravity to ensure that when the payload is

released, there are no significant changes in the stability of the aircraft. The other design

consideration is to ensure that the required components of the aircraft can all fit within the fuselage.

For the fire-fighting aircraft, the required components include the cockpit, the motor, the front and

rear landing gear, the wing attachment, tail attachments, and the payload and payload distribution

system. As the landing gear, wing location and the payload location are all determined by the

location of the centre of gravity, determining the size and layout of the fuselage is an iterative

process.

3.8.1 Cockpit Requirements

Figure 24 shows the final layout of the cockpit. These dimensions are based upon a combination of

recommended crew cabin dimensions for a light aircraft with a stick control, and for a transport

aircraft with a stick control as specified by Arjomandi (2009).


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Figure 24 - Cockpit Dimensions

3.8.2 Overall Design of the Fuselage

The final layout of the fuselage is shown in Figure 25 below.

Figure 25 - Fuselage Sketch

Using these dimensions, the maximum width of the fuselage was determined to be 90 inches as

shown in Figure 26. This value was selected based upon the required space for the storage of the

retardant, the width of the cockpit required for the comfort of the pilot, and also based upon the

aesthetics of the aircraft.


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Figure 26 - Front View of Fuselage Sketches

The overall length of the fuselage, and the length of the nose and tail sections, is dictated by the

fineness ratio. It is desirable to adhere to these recommended fuselage parameters to reduce

friction drag. The recommended fineness ratios for sub-sonic flight are given by Roskam (2004), and

are shown in Table 10.

Table 10 - Fineness Ratio as Specified by Roskam (2004)

Fineness Ratio Recommended Range

Total Fuselage F

FF

D

L=λ 6-9

Cone

F

FCFC

D

L=λ 2-3

Nose

F

FNFN

D

L=λ

1.2-2

The desired length of the fuselage and fuselage sections is dependent on the diameter of the

aircraft. This implies that an iterative process is required to determine the optimum solution. The

main driving parameter in determining these dimensions is the aircraft nose. The nose section of the

aircraft contains the majority of the aircraft components including the cockpit, the nose landing

gear, the motor (and associated air intake and outlet pipes), and a firewall to separate the cockpit

from the engine. Once this layout was sufficiently established, the height of the aircraft could be

determined, and using this along with a reasonable aircraft width, the fuselage proportions could be

determined.


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The final dimensions were determined to be as follows:

• Fuselage height: 120 inches (10 ft)

• Fuselage width: 90 inches (7.5 ft)

• Fuselage overall length (LF) : 630 inches (52.5ft)

• Nose length (LN) : 144 inches (12 ft)

• Cone length (LC): 315 inches (26.25 ft)

The nose length was dictated by the constraints of the motor, nose landing gear and the cockpit,

whilst the overall length was kept to a minimum and the cone length maximised to minimise the

weight of the aircraft. This was possible as all loads, excluding the structure and the empennage, are

located in the foremost half of the aircraft.

The ‘diameter’ DF used to determine the fineness ratio was taken to be the average of the fuselage

height and the width. This was determined to be 105 inches. It can be seen from the table below

that the aircraft fits within the recommended range for the fineness ratio to reduce the friction drag.

Table 11 - Comparison of the Fineness Coefficient for the Designed Aircraft Compared with the

Recommended Values as Specified by Roskam (2004)

Fineness Ratio Recommended range

Total Fuselage 6105

630 ===F

FF D

Lλ 6-9

Cone 3105

315 ===F

FCFC D

Lλ 2-3

Nose 37.1105

144 ===F

FNFN D

Lλ 1.2-2

3.8.3 Visibility Diagram

With the fuselage and cockpit layout finalised, pilot visibility was determined. As the aircraft is being

designed for aerial fire fighting, visibility is of considerable importance. Although the visibility over

the nose of the aircraft could not be improved upon the standard requirements, the inclusion of

large windows in the doors adds to the pilots’ visibility. The visibility diagram is shown in Figure 27.


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Figure 27 – Visibility Diagram

3.8.4 Fire Retardant Tanks and Distribution System

The retardant tanks are located on the centre of gravity. The system itself is required to drop

3746.5L of retardant, which equates to a total space envelope within the fuselage of 228627 square

inches for the retardant alone. The space envelope within the fuselage allowed is 353248 square

inches distributed abut the centre of gravity, to allow for sufficient room for tank structure and

baffles to prevent the effects of sloshing. To further allow for the distribution system, including the

payload bay doors to release the retardant, additional space has been left around the fuselage tank.

This is shown schematically in Figure 28.

It is intended that the tank can be split into components to allow for the distribution of the retardant

as required. Either an off-the-shelf or custom built distribution system could be accommodated

within the provided space.

Figure 28 - Tank Location in the Fuselage


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3.8.5 Fuselage Structure

Table 12 shows the frame depths, frame spacing and longeron spacing for a small commercial

aircraft as specified by Arjomandi (2009).

Table 12 - Recommended Frame and Longeron Spacing, and Frame Depth for a Small Commercial Aircraft as

specified by Arjomandi (2009)

Frame depth (inches) 1.25-1.75

Frame spacing (inches) 24-30

Longeron spacing (inches) 10-15

By considering each section of the fuselage separately, the appropriate frame spacing could be

determined. The frame spacing in the foremost half of the fuselage are primarily dictated by the

locations of the wing leading and trailing spars, as well as the fire wall. The spacing of the formers

around these components was designed to remain within the range specified above.

A firewall is located on an angle of approximately 35 degrees from the horizontal. This angle is

required to allow sufficient room for the air outlets for the motor, and to accommodate the landing

gear location. The main components in the nose of the fuselage (the motor and the landing gear) are

usually attached to the firewall and supported using truss structures. When designing the nose of

the aircraft, sufficient space was required to ensure the structure would fit inside the fuselage.

The longerons were similarly placed depending on the size and shape of the fuselage. These were

designed to ensure that the maximum spacing of 15 inches was not exceeded at any point in the

fuselage. In the region around the cockpit doors and potential payload bay doors, the longeron

spacing was reduced to reinforce the open structure. The frame depth was selected to be 1.5 inches.


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3.9 Propulsion System Design

Propulsion system design is an essential component of aircraft design. Propulsion system design

involves the decision to manufacture or purchase a pre-existing engine, followed by the selection of

the engine model and design integration. This process may flow systematically, but the conflicting

input from many subsystems often causes the process to be iterative. This iterative process is

amplified by the sensitivity of the propulsion system to weight. Increases in weight may result in the

selection of a different engine model or even an increase in the number of engines at later stages in

the design.

3.9.1 Propulsion System Type Selection

The selection of an optimal engine is fundamental for a successful propulsion system design. Engines

available for selection include piston, Wankel, rotary, radial, electric, turboprop, turbojet, turbofan,

ramjet and scramjet engines. The cruise speed of the aircraft critically affects the selected engine

type and is specified by the technical task. The selected engine type is largely independent of the

design of other systems such as weight, aerodynamics and structures, and consequently these

factors will be neglected when investigating engine type. Hence, engine type can be selected

considering only constraints from the technical task. Constraints due to other systems or aircraft

configurations can be neglected.

The technical task specifies a maximum speed of no less than Vmax = 202.5 knots (341.8 ft/s) and a

cruise altitude of 14,000 ft. At this altitude, the speed of sound a = 1061.4 ft/s.

Therefore, the Mach number can calculated as follows:

M = v/a

M = 341.8 /1061.4

M = 0.322

The primary selection criteria for engine type include the following:

• Suitability to aircraft operating envelope (including technology level, required power,

operating ceiling and cruising speed)

• High thrust to weight ratio at flight mach number

• Low Thrust Specific Fuel Consumption (TSFC) at flight mach number


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These criteria will be addressed in the following sub-sections.

Technical Task Requirements

The technical task does not outline any requirements regarding the propulsion system type or

number of engines.

Suitability to Aircraft Operating Envelope

Some engines listed in Section 3.9.1 can be eliminated, as they do not satisfy the conditions outlined

by the operating envelope of the aircraft. These are listed below:

Rotary Engine: Technology level has been surpassed, and are considered very heavy and

aerodynamically inefficient

Electric Engine: Does not satisfy the power requirement for a fire-fighting aircraft, and are best

suited for UAV or RC aircraft

Ramjet Engine: Requires the aircraft to be travelling at Mach numbers, M > 3 to initiate combustion.

As the maximum speed of the aircraft is orders of magnitude below the initiation speed, a ramjet

engine will not be considered for this application.

Scramjet Engine : Requires the aircraft to be travelling at Mach numbers, M > 5 to initiate

combustion. As the maximum speed of the aircraft is orders of magnitude below the initiation

speed, a ramjet engine will not be considered for this application.

Therefore, the remaining engines to be considered are Wankel, radial, turboprop, turbojet and

turbofan. Figure 5.4 in Brandt (2004, p. 178) shows that for a Mach number, M ≈ 0.3 and altitude h

=14,000 ft, a reciprocating propeller is the preferred engine type followed by turboprop, turbofan

and turbojet engine.

Thrust to weight ratio

The highest thrust to weight ratio is desired. Figure 5.2 in Brandt (2004 p. 176) shows that for a

Mach number M ≈ 0.3, an afterburning turbofan achieves the highest thrust to weight ratio. This is

followed by an afterburning turbojet, turboprop and low bypass ratio turbofan.

Thrust Specific Fuel Constant (TSFC)


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The lowest TSFC is desired. Figure 5.3 in Brandt (2004 p. 177) shows that for a Mach number, M ≈

0.3, a piston engine with propeller gives the lowest TSFC followed by turboprop, high by pass ratio

turbofan and low bypass ratio turbofan.

Recommendations

Initial analysis suggests the use of a piston engine with a propeller. A secondary recommendation

exists for a turboprop engine, followed by a low bypass ratio turbofan. Further investigations of

existing piston engines were conducted. Approximately 350 piston engines are listed by Jackson

(2008). Of these, only six provide a power output greater than 500hp. Initial design suggests that

the required power output would lie between 1250 – 3000 hp. Only one engine, the CRM 18DD/SS

provided a power output greater than 1,250hp. However, the CRM 18DD/SS weighed 3,745 lb,

which was considered prohibitive to use on the aircraft. Consequently, piston engines were not

selected as the engine type for the aircraft. An investigation of available turboprop aircraft was

undertaken. Eighteen of the fifty engines listed by Jackson (2008) provide a power output with the

desired 1500 – 3000 hp range. As such, enough variety existed within the turboprop range to allow

for design optimisation. Consequently, a turboprop engine was selected as the propulsion system

type.

3.9.2 Number of Engines and the Power Required per Engine

Initial Design

Initial estimation suggested a total required power output between 1250 – 3000 hp. The large range

in required power existed to encompass both the agricultural and regional jet prototypes. Early

analysis of current aircraft showed that both single engine and twin-engine aircraft existed within

this range. Engine number has a significant effect on configuration design. Consequently, it was

important to identify the point in regards to both power output and engine weight at which the

optimal design switches from single to twin engine. Data for the uninstalled power output and dry

engine weight data for several engines was obtained from Jackson (2008). Installation effects were

also considered. This required the reduction of output power and increase in engine weight.

Installed power output is defined below:

THP = ηp x SHP (Roskam III 2002)

Roskam III (2002) defines ηp = 0.88 for a turboprop.


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The installed engine weight is defined as follows:

WE Installed = 1.6 x WE (Component Weight 2009)

This data was plotted, and is presented in Figure 29 below. By approximating the data with a trend

line, the preferred selection ranges for one engine (1250 – 1850 eshp) and two engines (1850 – 300

eshp) were established.

Engine Selection

0

500

1000

1500

2000

2500

3000

1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200 3400 3600

Power (hp)

Wei

gh

t (l

b)

1 Engine 2 Engine

Figure 29 – Engine Selection: Single Engine versus Twin Engine

Configuration design gave preference to a single engine design.

Custom versus Existing Engines

The decision to design a new engine or purchase a pre-existing engine is fundamental to the design

of the propulsion system. Designing a new engine offers greater flexibility and delivers a product

that is ideal for the application. However, engine design is a lengthy, expensive process and beyond

the scope of this project. The small market that exists for fire-fighting aircraft does not justify the

expense of a new engine design. Consequently, a pre-existing engine will be selected for the

aircraft.

Maximum Power Requirement

Preliminary sizing from the matching diagram gave a required power of 1681 hp.


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Engine Model Selection

Engine models that provided a power output similar to 1681 hp were further investigated. Selection

was limited to single engines, as configuration design preferred this arrangement. Table 13 below

shoes data for suggested engine models.

Table 13 - Suggested Engine Models (Jackson 2008)

Manufacturer Designation

Number

of

Required

Engines

Installed

Power

(hp)

Installed

Weight

(lb)

Installed

Power to

Weight

Ratio

Specific Fuel

Consumption

(lb/(shp.hr))

General

Electric CT7-5A 1 1526.8 1252.8 1.22 Not available

General

Electric CT7-9 1 1707.2 1288 1.33 0.47

Honeywell TPE331-

14GR 1 1724.8 992 1.74 0.51

P&WC PW121 1 1848 1473.6 1.25 0.48

Klimov

TV3-

117VMA-

SB2

1 2200 2011.2 1.09 Not available

The General Electric CT7-9, the Honeywell TPE331-14GR, P&WC 121, and Klimov TV3-117VMA-SB2

all satisfy the installed power requirements. The Honeywell TPE331-14GR has a significantly higher

power-to-weight ratio than the other engines. From the available data, the General Electric CT7-9

has the lowest specific fuel constant. Although low specific fuel consumption was seen as a

desirable characteristic, it was not considered as critical as power-to-weight ratio. Consequently, one

Honeywell TPE331-14GR will be used to power the aircraft.

3.9.3 Propeller Sizing

The required propeller diameter can be determined from the following equation:

DP = ((4 x Pmax) / (π x np x Pbl))1/2

where DP is the propeller diameter and the maximum power per engine (installed) is Pmax = 1724 hp.

The blade power loading, Pbl, and the required number of blades, np, is determined from statistical

analysis of similar aircraft, and is summarised in Table 14 below.


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Table 14 - Statistical Analysis of Relevant Engines (Roskam III 2002)

Aircraft

Maximum Power

per Engine, Pmax

(hp)

Propeller

Diameter (ft)

Number of

Propeller Blades,

np

Blade power

loading Pbl

(hp/ft2)

Air Tractor AT-

310A 600 9.1 2 4.6

PZL-M18A 1000 10.8 4 2.7

Beech 1900 1100 9.1 4 4.2

EMB-110 Bandar 750 7.8 3 5.2

SF-340 1630 10.5 4 4.7

From the above statistical analysis,

Number of blades, np = 4 and blade power loading, Pbl = 4.5

Therefore,

DP = ((4 x 1724) / (π x 4.5 x 4))1/2

DP = 11.04 ft

Larger propellers are more efficient. However, the propeller tip speed must remain subsonic. The

propeller tip speed can be calculated as the vector sum of the rotational tip speed and the aircraft

forward speed.

Vrot = π x n x D

where n is the rotational speed of the engine, n = 1540 rpm = 25.66 rev/sec and D is the proposed

propeller diameter, D = 11.04 ft.

Vrot = π x n x D

Vrot = π x 25.66 x 11.04

Vrot = 890.2 ft/s

The tip velocity can then be calculated using the following e quation:

Vtip = √(Vrot2 + V

2)

The aircraft cruise velocity is V = 341.7 ft/s and the engine rotational speed is as calculated above.

Therefore,


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Vtip = √(890.22 + 341.7

2)

Vtip = 953.5 ft/s

This speed is below the speed of sound (a = 1061.4 ft/s) at the specified cruise altitude. Therefore,

the propeller tip speed maintains subsonic.

Propeller Material Selection

The maximum propeller tip speed dictates the material selection of the propeller. Metallic propellers

should be used for applications with a maximum propeller tip speed of Vtip = 950 ft/s, whilst wooden

propellers have a maximum propeller tip speed Vtip = 850 ft/s. The aircraft has Vtip = 953.5 ft/s, and

consequently, a metallic propeller will be used.

Propeller Type Selection

There are three main propeller types as outlined below:

• Variable pitch: Blade pitch is varied to maintain an optimal lift-drag ratio with speed,

which results in increased thrust across a range of speeds

• Constant speed: Blade pitch angle is varied to maintain constant speed, which improves

fuel efficiency

• Controllable pitch: Pilot can override constant speed mechanism, which is useful to

reverse the blade pitch angle to slow the aircraft down

The additional drag produced by a controllable pitch propeller is not required for the relatively light

aircraft designed in this project. Increased thrust is considered advantageous over increased fuel

efficiency, as it will improve the manoeuvrability of the aircraft. Consequently, a variable pitch

propeller will be selected for the aircraft.

Specific Propeller Selection

The Dowty Aerospace propeller (c) R.389/4-123-F/25 was selected for this application. This

propeller has a diameter of 11 ft (Dowty Propeller 2007) which meets the requirements.

3.9.4 Propulsion System Integration

The following section of the report focuses on the integration of the propulsion system into the

overall aircraft design. Integration includes the selection of the installation configuration, location

and the mounting of the engine. Finally, checks are performed to ensure complete compatibility

with other aircraft systems.


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Pusher/Tractor Selection

Three options exist for the configuration of propeller engines: tractor, pusher and mixed, as shown

in Figure 30 below.

Figure 30 - Propeller Engine Configurations: Tractor and Pusher (Raymer 2006 p.252)

A mixed installation requires two engines, one located as a pusher and the other as a tractor. This is

not appropriate for this design as it requires at least two engines, and hence, will not be discussed

further. Tractor installations place the inlet in the free airstream, resulting in improved engine

cooling. Furthermore, this layout improves the stability of the aircraft, allowing shortening of the

fuselage and a reduction in tail size. Pusher installations reduce the flow disturbance over the wing,

decreasing the skin friction drag and allowing the wetted area of the aircraft to be reduced. Other

benefits of the pusher configuration include improved visibility for the pilot and reduced cabin noise.

However, in a pusher configuration, the propeller receives disturbed airflow, substantially reducing

its efficiency. Additionally, pusher configurations may require larger tail areas, longer landing gear

and are more likely to suffer from FOD damage. These disadvantages are significant and resulted in

the selection of a tractor configuration for the aircraft.

Engine Mounting Selection

Figure 31 below shows the possible mounting locations for aircraft engines, including the fuselage,

wings, tail or as part of an upper fuselage pod.


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Figure 31 - Engine Mounting Locations: Fuselage, Wings, Tail or as Upper Fuselage Pod

(Raymer 2006, p.252)

Wing mounting is not appropriate for this design as only one engine is used. Mounting engines on

the tail or as part of upper fuselage pods results in a high thrust line that degrades the control

characteristics of the aircraft. Consequently, this engine arrangement is used only for applications

that require significant engine clearance, notably, amphibious aircraft. The aircraft does not require

this level of clearance, and hence, the engines will not be mounted in a tail or upper fuselage pod.

Honeywell TPE331-14GR Specifications

The geometry of the Honeywell TPE331-14GR is shown in Figure 32 below.

Figure 32 - Honeywell TPE331-14GR Geometry (all dimension in inches) (Honeywell 2006)

The Honeywell TPE331-14GR has the following dimensions:


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Length: 52.3 inches

Width: 23.0 inches

Height: 36.5 inches

The engine has five mounting points. The locations of these are shown in Figure 32.

The centre of gravity of the engine was not stated by the manufacturer. Consequently, it was

assumed that the centre of gravity was located at the geometric centre of the engine.

The specification of the Honeywell TPE331-14GR are listed in Appendix F.

Cooling System Configuration

Cooling systems can be configured in an updraft or downdraft arrangement as shown in Figure 33

below.

Figure 33 - Cooling System Configuration (Raymer 2006, p.256)

Updraft arrangements have maximum cooling efficiency but exhaust hot dirty air in front of the

windscreen. This can cause the cabin to heat up, and in the event of an oil leak, can reduce pilot

visibility. Downdraft arrangements do not suffer from these problems, but have reduced cool

efficiency. As ground visibility is considered critical, a downdraft configuration will be utilized.

Air Intake Sizing

Raymer (2006) states that area required for the cooling intake can be determined using the

following equation:

A intake = P / (2.2 x Vclimb)

where P is the installed power, (1724.8 hp). The climb speed Vclimb is assumed to be the average of

the cruise and takeoff speeds. From the technical task, Vclimb =187.5 mph, (275 ft/s).


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Therefore,

A intake = 1724.8 / (2.2 x 275)

A intake = 2.85 ft2

Air Exhaust Sizing

Roskam (2006) suggest that air exhausts should be sized as follows:

A exhaust / A intake = 0.8

Therefore, the recommend exhaust area is A exhaust = 2.28 ft2

Firewall

Firewalls prevent the spread of heat or fire from the engine into the cockpit. Raymer (2006) states

the requirement of a 0.015 inch thick sheet of stainless steel with no cut out to act as a firewall. This

sheet should be attached to the first structural bulkhead of the fuselage. Any wires that pass

through the firewall must have a fireproof sealing.

Fuel Type Selection

The Honeywell TPE331-14GR can be powered by Jet A, Jet B, Jet A-1, JP-4, JP-5, JP-8 JP8+100 fuel.

Jet A fuel is only available in the U.S.A, and hence, is not appropriate for an Australian application.

JP-4, JP-5, JP-8 JP8+100 are military standard fuels, and are not appropriate for a civil application.

Jet A-1 and Jet B have similar properties. Jet B has improved cold weather performance but is

significantly more difficult to handle. As low temperatures (below 0oC) are not expected for fire-

fighting applications, Jet A-1 will be selected as the fuel type.

Fuel Tank Type

The three fuel tanks types are discrete, integral, and bladder. Discrete tanks are fabricated

separately and then mounted to the aircraft. Discrete tanks are used predominantly for general

aviation aircraft. Bladder tanks are a thick rubber bag stuffed into a cavity of the structure. Bladder

tanks are self-sealing, but significantly reduce the available volume for fuel, and hence, are preferred

for military applications, which benefit from the self-sealing capability. Integral tanks are part of the

aircraft structure that has been sealed to form a tank. Consequently, an integral tank will be used

for the aircraft.


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Fuel Tank Sizing

Sizing calculations give a fuel weight of Mf =1996.3 lbs.

Jet A-1 fuel has a density, ρjet A-1 = 6.7 lb/gal

Therefore,

Vfuel = m/ ρ

Vfuel = 1996.3 / 6.7

Vfuel = 298.0 gal

Vfuel = 39.8 ft3

The use of porous foam is recommended to reduce the risk of fire hazard. Foams require additional

volume due to displacement and absorption of fuel. Raymer (2006) suggest that an additional five

percent of volume is required due to the foam.

Vfoam = 1.05 x Vfuel

Vfoam = 1.05 x 39.8 ft3

Vfoam = 41.8 ft3

Raymer (2006) states that 85% of the external wing volume is available for integral wing tanks.

Consequently, a total external volume for both wings External = 49.2 ft3. Assuming losses of up to five

percent gives External = 52 ft3. This will be split over both wings. Hence, individual wings must have

External single wing = 26 ft3.

The wing design suggests an exterior volume greater than 30 ft3 per wing. Consequently, this fuel

tank volume is acceptable.


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3.10 Wing Design

The following section of the report details the wing design. The geometry of the wing, including

vertical position, sweep, aspect ratio, thickness ratio, taper ratio, twist, dihedral, wing loading,

incidence angle and longitudinal position are considered. An aerofoil selection is summarised,

followed by control surface sizing, wing tip selection and a summary of the wing structure.

3.10.1 Vertical Position

An aircraft can have three main vertical positions for the wing. A high wing is mounted above the

fuselage, a low wing is mounted below the fuselage and a mid wing is mounted through the centre

of the fuselage. An important consideration for a fire-fighting aircraft is ground visibility. A high wing

configuration offers the best ground visibility. A further consideration is the loading and unloading of

fire retardant. High wing aircraft are preferred for cargo applications, as no special equipment is

needed for loading and unloading. A high wing configuration has high lateral stability and a lighter

structure, as the internal volume of the fuselage is not cut by wing spars and other structural

elements.

Incorporating landing gear into a high wing aircraft is often difficult as a large bay is required inside

the fuselage for the retractable landing gear. The problem can be overcome by designing an

appropriately sized area within the fuselage for the retracted landing gear. High wing aircraft are

also less survivable during crashes in comparison to low wing aircraft. However, there are no

passengers on board a fire-fighting aircraft and the aircraft is flown by an experienced pilot. Hence,

crashworthiness is considered a minor issue. The aircraft fuselage will be designed to bear the

impact loads generated by the wing in a crash.

The high wing configuration has many advantages over a mid wing configuration and a low wing

configuration. The disadvantages of a high wing configuration were considered reasonable for the

application. Hence, a high wing configuration was chosen for the fire-fighting aircraft.

3.10.2 Sweep

Wing sweep is defined as the angle between the leading edge of the wing and the perpendicular to

the fuselage. Wings can either be swept or unswept, depending on the application. An unswept wing

has low weight, as the wing does not require additional structural supports, and exhibits good stall

behaviour. An unswept wing also has good runway visibility, as sweep reduces the lift-curve slope,

which causes the aircraft to have more pitch attitude. Additionally, unswept wings are cheap and


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easy to manufacture, as all structural components are simple and all wing ribs can be made the

same.

A swept wing reduces compressibility drag. However, the fire-fighting aircraft only has a cruise

velocity of 350 kph, and such, compressibility effects would be marginal. A swept wing has higher

longitudinal stability, as the sweep allows the aerodynamic centre to move faster than the centre of

gravity. Additionally, sweep changes the longitudinal moment arm, which has a beneficial effect on

the inherent longitudinal damping characteristics of the aircraft. Swept wings have increased ride

quality. However, there are no passengers on board the fire-fighting aircraft and the aircraft is flown

by an experienced pilot. Hence, ride quality is considered a minor issue.

The advantages of low weight, good structure, good stall behaviour and ease of manufacture were

seen as significant. Hence, an unswept wing was chosen for the fire-fighting aircraft.

3.10.3 Aspect Ratio

Aspect ratio is defined as the square of the wing span divided by the wing area. A high aspect ratio

wing has low induced drag, a high lift-curve slope, good runway visibility from the cockpit and a

higher span. However, a high aspect ratio wing has decreased ride quality through turbulence. High

aspect ratios lead to steeper lift-curve slopes such that aircraft are more sensitive to changes in

angle of attack. Hence, the ride quality of the aircraft is reduced. However, the aircraft is not a

passenger aircraft. Hence, ride quality is considered a minor issue.

High aspect ratio wing require longer structural supports which corresponds to a higher overall wing

weight, and experience low aeroelastic stability. The aircraft has a cruise velocity such that

aeroelastic stability effects would be low. Additionally, high aspect ratio wings have low lateral

stability. However, the fire-fighting aircraft has a high wing configuration, and as such, has a high

lateral stability.

The advantages of low induced drag and good runway visibility were seen as significant. Hence, a

high aspect ratio was chosen for the fire-fighting aircraft.

An average aspect ratio of eight was calculated from the statistical analysis. As such, an aspect ratio

of eight was chosen for preliminary sizing purposes.


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3.10.4 Thickness Ratio

Thickness ratio is defined as the maximum thickness of the wing divided by the chord length of the

wing. A thick wing is lightweight due to the increased bending and torsional stiffness, and provides

maximum lift coefficients. A thick wing can accommodate more fuel volume but has higher profile

drag in the subsonic flight regime.

The advantages of a lightweight and maximum lift were seen as significant. A high thickness NACA

4415 aerofoil was chosen for the fire-fighting aircraft (see Section 3.10.11). This aerofoil has a

thickness of 15%, which is a suitable value for obtaining maximum lift coefficients.

3.10.5 Taper Ratio

Taper ratio is defined as the ratio of the tip chord to the root chord. Low taper ratio reduces the

weight of the wing as the wing lift distribution tends to zero at the wing tip and the area of the wing

near the wing tip is not fully loaded. A wing with a taper ratio of one is also cheap and easy to

manufacture, as all structural components are simple and all wing ribs are the same. The wing tip of

a low taper ratio wing tends to stall sooner as it flies on lower Reynolds’s number airflows, and has a

lower maximum lift coefficient. Additionally, a high taper ratio increases the amount of fuel that can

be stored in the wings. However, a thick wing was chosen to negate these issues.

The advantages of reduced wing tip stall and ease of manufacture were seen as significant. Hence,

no taper was chosen for the fire-fighting aircraft.

3.10.6 Twist

Wing twist occurs when the tip aerofoil has a lower or higher angle of incidence than the root

aerofoil. Wings that have no twist are easy and cheap to manufacture, as all structural components

are simple and all wing ribs can be the same. Wings that have no twist have decreased induced drag.

However, wings that have no twist experience wing tip stall that can generally occur in an

asymmetric manner and cause serious roll control problems. However, a thick wing was chosen to

provide high maximum lift coefficients to negate this problem.

The advantages of decreased induced drag and ease of manufacture were seen as significant. Hence,

no wing twist was chosen for the fire-fighting aircraft.


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3.10.7 Dihedral

A high wing configuration has an inherent dihedral effect that causes the rolling moment due to the

sideslip derivative to be negative. This means that the aircraft has more spiral stability and less dutch

roll stability. Hence, no dihedral was chosen for the fire-fighting aircraft as it has a high wing

configuration.

3.10.8 Wing Loading

The matching diagram was used to determine the most favourable wing loading. In deriving the wing

planform, the corresponding wing loading of the selected design point was reassessed to predict

whether such design is feasible.

Low wing loading provides a shorter takeoff and landing distance, but requires a larger wing area

that increases the weight of the wing. Short takeoff and landing distance is not considered an

important issue as the aircraft is designed to operate out of paved runways. Low wing loading is

used for aircraft that are required to fly at high altitude, which is not an important parameter in the

design of a fire fighting aircraft. Also, low wing loading results in a higher response to changing angle

of attack which corresponds to poor ride quality. However, there are no passengers on board the

aircraft and the aircraft is flown by an experienced pilot. Hence, ride quality is not considered a

major issue.

High wing loading allows the cruise lift coefficient to be similar to that at (L/D)max. A high wing

loading also requires the aircraft to resist higher accompanying stresses. Hence, high wing loading

increases the cost and complexity of manufacture as it requires materials that are more expensive

and more complex manufacturing methods.

The matching diagram was used to determine the aircraft wing loading of 36.85 lbs/ft2.

3.10.9 Wing Longitudinal Location

From the statistical analysis, the average wing leading edge location as a percentage of the fuselage

length was determined to be 26%. Hence, this value was used for preliminary sizing. The initial

fuselage length was 52.5 ft, which gives the wing leading edge location from the nose of the aircraft

as 13.65 ft. Throughout the design process, this was modified to correspond with the geometry of

the aircraft. Hence, the wing longitudinal located was adjusted to be 12 ft from the nose of the

aircraft.


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3.10.10 Aerofoil Selection

The design of the aerofoil section for the wing is critical for ensuring the aircraft can achieve the

required performance. The shape of the aerofoil affects the lift and performance of the aircraft in all

flight regimes, including cruise, takeoff and descent (Raymer 2006).

Operational Reynolds Number

• L = Cwing = 8.59 ft.

• VTO = 153.17 ft/s.

• ν = 1.57 x 10-4

ft2/s at sea level.

Re = VTO L/ν = (153.17)(8.59)/(1.57 x 10-4

).

→ Re = 8.38 x 106.

The aerofoil must be suitable for operation in airflow with a Reynold’s Number Re = 8.38 x 106.

Maximum Lift Coefficients

For an agricultural aircraft, CL max = 1.3 - 1.9 (Roskam 2005). Hence, an average value of 1.6 will be

chosen as the preliminary CL max.

For an untwisted, constant-aerofoil-section wing, CL max/Cl max = 0.9 (Raymer 1992).

→ Cl max = CL max/0.9 = 1.6/0.9.

→ Cl max = 1.78.

The aerofoil must be selected to provide the desired maximum wing lift coefficient CL max = 1.6 and

the desired maximum aerofoil lift coefficient Cl max = 1.78. There are also some additional

considerations as outlines below.

• The aerofoil must have the highest possible (L/D)wing compared with similar aerofoils to allow

the aircraft to achieve the highest possible (L/D)aircraft.

• The aerofoil must have a low pitching moment coefficient Cm to reduce the torsional loads

and induced drag from trimming.


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Design Lift Coefficient

(W/S) = 36.85 lbs/ft2.

ρcr = 0.00116355 slugs/ft3 at a cruise altitude of 22,500 ft.

Vcr = 341.75 ft/s.

L = W = 0.5ρcrVcr2SCL.

→ CL = (W/S)(1/0.5ρV2).

→ CL = (36.85)(1/(0.5*0.00116355*341.752).

→ CL = 0.54.

The design lift coefficient CL = 0.54.

Aerofoil Selection Process

The aerofoil selection process compared the two-dimensional flow performance of the aerofoil

candidates over the range 0o < α < 20

o. The two dimensional performance of the aerofoils differ from

a three dimensional wing. However, a suitable indication of (L/D)wing can be obtained from two-

dimensional data. For the purpose of aerofoil comparison, it was assumed that the aerofoil with the

highest (L/D)aerofoil would produce the wing with the highest (L/D)wing.

Similarly, the aerofoil with the lowest, most constant section pitching moment coefficient Cm would

produce the wing with the lowest, most constant pitching moment coefficient CM. JavaFoil (2009)

was used to compare the performance and suitability of each of the selected aerofoils. The selected

aerofoil profile was to have the properties as outlined below.

• High (L/D)aerofoil

• Low, constant Cm

• Cl max > 1.78 such that CL max > 1.78 after three dimensional correction

Aerofoil Candidates

Three possible aerofoil profiles were identified from research of the aerofoils used on existing

agricultural aircraft. The aerofoils are presented in Table 15 below.


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Table 15 - Aerofoil Candidates

Aerofoil Aircraft Reference

NACA 4415 Air Tractor AT-301 through AT-802 UIUC 2008

NACA 4416 M-18A Dromader UIUC 2008

NACA 4412 Grumman G-164 Ag-Cat UIUC 2008

NACA 4415 Pacific Aerospace Cresco 08-600 UIUC 2008

2D Analysis

Table 16 shows a comparison between the selected 2D aerofoils.

Table 16 - 2D Aerofoil Comparison Table

Aerofoil Cl max Cd Approximately

constant Cm

Average Cm

NACA 4412 1.904 0.01340 Yes -0.12

NACA 4415 2.184 0.01407 Yes -0.13

NACA 4416 2.306 0.01696 Yes -0.13

From Table 16, all the aerofoils provide the minimum desired Cl max value of 1.78, and they all have

similar values for the pitching moment coefficient Cm. Hence, a 3D analysis is required to determine

the most suitable aerofoil.

3D Analysis

3D flow effects cause wings to have lower lift coefficients than the 2D aerofoil lift coefficients.

Consequently, a correction for 3D flows will be considered.

Table 17 shows the results for the 3D aerofoil analysis.

Table 17 - 3D Aerofoil Comparison Table

Aerofoil Cl max Cd (L/D)max Approximately constant Cm Average Cm

NACA 4412 1.535 0.10715 14.33 Yes -0.12

NACA 4415 1.78 0.14341 12.42 Yes -0.13

NACA 4416 1.859 0.15443 12.04 Yes -0.13

The NACA 4412 does not achieve the desired Cl max value of 1.78, whereas the NACA 4415 and NACA

4416 both achieve the desired Cl max value. Both the aerofoils have similar values for the average Cm,

but the NACA 4415 has a higher value for (L/D)max.


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The NACA 4415 is the most appropriate aerofoil to choose for the fire-fighting aircraft, as it provides

an appropriate value for Cl max, has a small Cm value and has a high L/D at a low angle of attack.

Cl max = 1.78.

For an untwisted, constant-aerofoil-section wing, CL max/Cl max = 0.9 (Raymer 1992).

→ CL max = 0.9*Cl max.

→ CL max = 0.9*1.78.

→ CL max = 1.6.

Flap Selection

Most agricultural aircraft in operation utilise Fowler flaps. Hence, Fowler flaps were selected for the

firefighting aircraft.

Fowler flaps provide ΔCl max = 1.3.

Assume Sflapped/Sref=0.1.

Maximum Lift Coefficient for Takeoff:

ΔCL max, TO = (0.7)(ΔCl max)(Sflapped/Sref)(cosΛhinge).

→ ΔCL max, TO = (0.7)(1.3)(0.1)(cos(1)).

→ ΔCL max, TO = 0.091.

For a high aspect ratio wing:

CL max TO = Cl max (CL max/Cl max) + ΔCL max, TO.

→ CL max TO = (1.78)(0.9) + 0.091 = 1.693.

Maximum Lift Coefficient for Landing:

ΔCL max, L = (ΔCl max)(Sflapped/Sref)(cosΛhinge).

→ ΔCL max, L = (1.3)(0.1)(cos(1)).

→ ΔCL max, L = 0.13.

For a high aspect ratio wing:

CL max, L = Cl max (CL max/Cl max) + ΔCL max, L.

→ CL max, L = (1.78)(0.9) + 0.13 = 1.732.


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3.10.11 Incidence Angle

The wing incidence angle is calculated based on the two factors: the cruise drag and the floor

attitude at cruise. The incidence angle should be chosen so that during the main part of cruise, the

fuselage has no angle relative to the oncoming airstream. If the fuselage cruises nose up or nose

down, the total drag of the fuselage is increased. The floor attitude in cruise is also influenced by the

choice of incidence angle.

The following calculations show the process used to determine the wing angle of incidence.

WTO = 19,735.35 lbs.

S = 535.56 ft2.

ρcr = 0.0015455 slugs/ft3.

Vcr = 341.75 ft/s.

L = WTO = 0.5ρV2SCL.

CL = WTO/(0.5ρcrVcr2S).

→ CL = (19,735.34)/(0.5*0.0015455*(341.752)*535.56).

→ CL = 0.41.

From the CLα curve, CL = 0.41 corresponds to 0 degrees angle of attack. Hence, the wing angle of

incidence is 0 degrees.

3.10.12 Flap Sizing

Table 28 in Appendix G shows the flap chord ratio, the flap location from the fuselage (inboard) and

the flap location from the fuselage (outboard) for some common agricultural aircraft. From the

table, the flap chord ratio used for preliminary sizing was determined to be 21.11%. Similarly, the

flap location from the fuselage (inboard) and flap location from fuselage (outboard) for preliminary

sizing were determined to be 6% and 56% respectively.

3.10.13 Aileron Sizing

Table 29 in Appendix G shows the aileron chord ratio, the aileron location from the fuselage

(inboard) and the aileron location from the fuselage (outboard) for some common agricultural

aircraft. From the table, the aileron chord ratio used for preliminary sizing was determined to be


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23.30%. Similarly, the aileron location from the fuselage (inboard) and aileron location from fuselage

(outboard) for preliminary sizing were determined to be 57% and 94% respectively.

3.10.14 Spoiler Selection

Spoilers are plates located forward of the flaps on the top of the wing and aft of the maximum

thickness point. Spoilers are deflected upwards into the oncoming air stream and are used to spoil

the air over the wing surface immediately behind the spoiler, which causes a reduction in lift.

Spoilers are commonly used on large transport aircraft to augment roll control at low speeds.

However, spoilers have a non-linear response that makes them difficult to use for roll control when

using a manual flight control system (Raymer 1992). Due to the size and performance characteristics

of the fire-fighting aircraft, spoilers are not required.

3.10.15 Flow Control Devices

The aircraft wing has no sweep, so no loss of stability occurs at the wing tips due to the thickening of

the boundary layer and airflow separation. Hence, no overall lift is lost at the wing tips, and ailerons

are not affected. Hence, flow control devices are not required.

3.10.16 Wing Tips

Table 18 summarises the advantages and disadvantages of different wing tips.

Table 18 - Wing Tip Table

Wing Tip Advantages Disadvantages

Rounded Aesthetic High induced drag

Sharp Low induced drag Difficult to manufacture

Cutoff Low induced drag, simple and cheap to

manufacture None

Hoerner Low induced drag Difficult to manufacture

Dropped Increases effective span without

increasing actual span Difficult to manufacture

Upswept Increases effective span without

increasing actual span Difficult to manufacture

Aft-swept Low drag Increases wing torsional loads

End plate Prevents air flowing beneath the wing

escaping around wingtip High drag

Winglet High drag reduction Flutter, twist and camber must

be optimised for one velocity

Cutoff wing tips are the simplest, cheapest and easiest wing tips to manufacture, and do not

increase induced drag. Hence, the fire-fighting aircraft shall be designed with cutoff wing tips.


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3.10.17 Centre of Gravity

The centre of gravity of the wing was determined by calculating the centroid of the NACA 4415

aerofoil, and was determined to be 42% of the wing chord.

Cwing = 8.18 ft.

CGwing = (0.42)(8.18).

→ CGwing = 3.4356 ft from the leading edge of the wing.

3.10.18 Structure

The structure of the wing is relatively simple, as the wing has no taper, twist, sweep, dihedral or

angle of incidence. The wing structure consists of a 0.13 inch (3.3mm) thick skin, 0.13 inch (3.3mm)

thick wing ribs and a 1 inch thick main spar located 25% back from the leading edge of the wing. The

wing ribs are placed 32.73 inches from each other, and have two holes cut out to reduce the weight

of the wing. Ailerons and flaps are mounted to an auxiliary spar at the rear of the wing, which is

located 77% back from the leading edge of the wing. Fuel is placed in tanks that sit in between the

main and auxiliary wing spars. The skin, spars and solid ribs at 130.92 inches from the fuselage

centre line provide an enclosed volume of 50 ft3, which is the fuel volume required. The wing ribs

within the fuel tanks act as structural support for the wing and baffles to prevent fuel frothing. Refer

to the isometric views or three view drawings for additional information.


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3.10.19 Wing Design Summary

Table 19 summarises the above wing design.

Table 19 - Wing Design Summary

Parameter Value

Vertical position High

Wing loading 36.85 lbs/ft2

Area 535.56 ft2

Span 65.46 ft

Chord 8.18 ft

Sweep 0 degrees

Aspect ratio 8

Thickness ratio 15%

Taper ratio 1

Twist None

Dihedral angle 0 degrees

Incidence angle 0 degrees

MAC 8.18 ft

Aerofoil NACA 4415


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3.11 Empennage Design

3.11.1 Empennage sizing

Figure 34 below shows some possible configurations for the empennage design. Raymer (2006)

recommends the use of the conventional arrangement for conventional aircraft as the configuration

will provide adequate stability and control at the lightest weight. Other configurations considered

were the T-tail, the cruciform, the V-tail and the H-tail. The T-tail and the H-tail were not chosen, as

they are heavier than the conventional configuration for an unnecessary gain in stability. The

cruciform was not chosen, as it was not as stable as the conventional configuration. A conventional

tail configuration was chosen for the fire-fighting aircraft application.

Figure 34 - Empennage Configurations (Raymer 2006)

The horizontal stabiliser is the component of the empennage that lies in the horizontal plane. A

statistical approach was used to calculate the area of the horizontal stabiliser. The statistical

approach involves the use of a tail volume coefficient and Raymer (2006) provides data for the

parameters used. The aircraft is modelled as an agricultural aircraft and the volume coefficient VH

was determined to be 0.5. The formula involves the reference area S, which is calculated from the

aspect ratio. The distance between the MAC of the tail and the MAC of the aircraft was calculated

from the stability analysis as 35ft and the chord of the wing as 8.18ft. The horizontal area was

calculated as follows:

The vertical stabiliser is the component of the empennage that lies in the vertical plane. The vertical

stabiliser was calculated in similar way. The tail volume coefficient was determined for an

agricultural aircraft from Raymer (2006) to be 0.04. The parameter


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3.11.2 Horizontal Stabiliser Geometry

The calculation of the horizontal stabiliser dimensions incorporates the tail aspect ratio and taper

ratio. Raymer (2006) recommends that a horizontal stabiliser have an aspect ratio of 4.0 and a taper

ratio of 0.4. The horizontal stabiliser will be configured as shown in Figure 35 below.

Figure 35 - Horizontal Stabiliser Arrangement

The total area of the horizontal stabiliser is calculated as follows:

The aspect ratio is defined as the square of the wing span divided by its area:


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3.11.3 Vertical Stabiliser Geometry

The calculation of the vertical stabiliser dimensions incorporates the tail aspect ratio and taper ratio.

Raymer (2006) recommends that a vertical stabiliser have an aspect ratio of 1.2 and a taper ratio of

0.4. The vertical stabiliser will be configured as shown in Figure 36 below.

Figure 36 - Vertical Stabiliser Arrangement

The total area of the vertical stabiliser is calculated as follows:

The aspect ratio is defined as the square of the span of the wing divided by its area:

3.11.4 Elevator Sizing and Geometry

Raymer (2006) states that the ratio of the area of the elevators to the area of the horizontal tail is

between 0.25 (for a jet transport) and 0.45 (for a general aviation aircraft). The ratio for this aircraft

is 0.3, as fire-fighting aircraft require somewhat more control authority than a jet transport but less

than a general aviation aircraft. The area of the elevators can now be calculated.


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A trim tab will be placed in the elevator arrangement, and will be sized by a similar volume

coefficient method. The volume coefficient for the elevator trim tab is 0.09 (Raymer 2006).

Due to the position of the vertical stabiliser, there is a spanwise area on the horizontal where an

elevator cannot be placed. The thickness of the vertical stabiliser is chosen later in this section and

the thickness to chord ratio is 13%. The chord of the vertical stabiliser at the root was found to be

10.80 ft. This results in a width of 1.40 ft where no elevator can be placed. A gap of 6 inches is placed

between the vertical tail infringement and the start of the elevator. The chord at this location is 5.21

ft.

The geometry of the elevator is shown in Figure 37 below:

Figure 37 - Elevator Geometry

The elevator is chosen to be 40% of the chord of the horizontal stabiliser. Using a similar approach to

the stabiliser sizing, the elevator dimensions are now calculated.

The trim tab will be located at the outboard section of the elevator. The trim tab configuration is

shown in Figure 38 below.


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Figure 38 - Elevator Trim Tab Geometry

From the elevator sizing, . We calculate the span of the trim tab using the following

formula:

3.11.5 Rudder Sizing and Geometry

Similar to the elevator sizing, Raymer (2006) states that the ratio of the area of the rudders to the

area of the vertical tail is between 0.35 and 0.45. The ratio for this aircraft is chosen to be 0.4. The

area of the rudders can now be calculated.

A trim tab will be placed in the rudder arrangement, and will be sized by a similar volume coefficient

method. The volume coefficient for the rudder trim tab is 0.09 (Arjomandi 2009).

The geometry of the rudder is shown in Figure 39 below.


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Figure 39 - Rudder Geometry

The rudder is calculated to be 40% of the chord of the vertical stabiliser. Using a similar approach to

the stabiliser sizing, the rudder dimensions are now calculated.

The trim tab will be located at the topmost section of the rudder. The trim tab configuration is

shown in Figure 40 below.

Figure 40 - Rudder Trim Tab Geometry

From the rudder sizing, . We calculate the span of the trim tab using the following

formula:

3.11.6 Stabiliser Aerofoils

Raymer (2006) recommends that the vertical and horizontal stabilisers have a thickness of 1%-2%

less than the wing, and are symmetric. The wing has a thickness of 14%. For this reason, a NACA


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0012 aerofoil was chosen for both stabilisers.


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3.12 Landing Gear Design

Landing gear placement is essential for ground stability and controllability. A good landing gear

position must provide superior handling characteristics and must not allow over-balancing during

takeoff or landing. The following landing gear characteristics will be determined in this section:

• Landing gear arrangement

• Gear placement criteria

• Landing gear position

• Landing gear loads

• Number, type and size of tyres

• Tyre pressure calculations

• Suspension method and requirements

• Length and diameter of landing gear struts

• Nose-wheel steering and castoring dimensions

• Gear retraction geometry

3.12.1 Landing gear arrangement

Landing gear arrangements are included in Figure 41 below. The two most common landing gear

arrangements for high-wing designs are the tail-dragger and tricycle arrangements (Raymer 2006).

Figure 41 – Landing Gear Configurations (Raymer 2006)


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Bicycle and single main landing gear arrangements are less preferable due to the inherent instability

on the ground. Outrigger wheels are required on the extremes of the aircraft, and the high-wing

configuration makes the placement of these difficult (Raymer 2006). The outrigger wheels would

need to be long to reach from the wing to the ground. The weight of these outrigger wheels would

be significant, and the storage of them difficult. The quadricycle arrangement would involve a

significant increase in weight in comparison to the tricycle and tail-dragger arrangements. The

stability is increased significantly due to the wheel locations and the loads on each wheel are

reduced due to the added wheel (Raymer 2009). The quadricycle arrangement is not considered due

to the width required in storing the landing gear in the fuselage when the gear is retracted. The

fuselage design is not of sufficient width to house all four landing gear.

Both the tricycle and the tail-dragger arrangements are used for high wing aircraft. The tricycle gear

arrangement provides good steering and ground stability characteristics. The advantage of a flat

cabin floor allows for good visibility take-off and during approach as well as the ability to store and

load cargo horizontally. The advantages of flat storage and loading of cargo are not applicable to the

fire-fighting application. The tail-dragger allows an increased angle of attack at take-off and landing

(Torenbeek 1982). This decreases the take-off and landing distances for the aircraft in comparison to

a tricycle gear. Tail-dragger gears are typically smaller, are thus lighter, and require less storage

space in the fuselage (Raymer 2006). Tail-dragger arrangements are unstable during turning

manoeuvres on the ground, due to the centre-of-gravity being located behind the main landing gear.

This significant decrease in stability was considered prohibitive to this design.

A tricycle arrangement was chosen for this configuration due to its good stability and steering, as

well as good visibility.

3.12.2 Landing Gear Sizing Nomenclature

Figure 42 below shows the nomenclature used throughout the landing gear sizing section of this

report. All symbols are defined in the nomenclature list at the beginning of this report.


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Figure 42 - Landing Gear Nomenclature (Roskam 2006)

3.12.3 Landing Gear Placement Criteria

Raymer (2006) gives five criteria for locating the landing gear on the aircraft. These criteria are

outlined below:

• The nose weight criterion

• The height criterion

• The roll-over criterion

• The over-turn angle criterion

• The tip-back angle criterion

3.12.4 Nose Weight Criterion

The nose weight criterion ensures that the correct proportion of weight is carried by the nose gear.

The nose wheel is required to carry more than 5% of the aircraft weight at take-off and after landing.

This allows enough traction on the tyre of the nose-wheel to permit nose-wheel steering (Raymer

2006). The proportion of loads on the nose wheel should be less than 20%. An increased proportion

of weight on the nose wheel results in a more difficult take-off as a larger speed is required to create

the lift required for takeoff rotation (Torenbeek 1982). The upper limit of 20% on this criterion

allows for a reasonable takeoff speed (Raymer 2006). The nose wheel criterion can be expressed as

follows:


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3.12.5 Height Criterion

The height criterion ensures that there is sufficient clearance for the fuselage and propeller including

required safety clearances. The landing gear calculations can determine the vertical height of each

gear. This height is measured from the ground to the centre of gravity of the fuselage. The height of

the landing gear must be greater than the vertical distance between the centre of gravity of the

fuselage and the bottom of the fuselage at the landing gear attachment point. Further, it is required

that the height of the nose landing gear allow enough height for proper rotation of the propeller.

The propeller diameter will be 11ft but is not located at the vertical centre of the fuselage. A

propeller clearance of at least 7” is required for safety purposes (Arjomandi 2009). From preliminary

drawings, the distance from the ground to the centre of the propeller disc is 2.85 ft. For our fuselage

and propeller dimensions, the height criterion is expressed as follows:

The over-turn angle criterion regards ground stability during taxiing. According to Raymer (2006, pg

232), the over-turn angle is “measured as the angle from the [centre of gravity] to the main wheel,

seen from the rear at a location where the main wheel is aligned with the nose wheel”. This

dimension is illustrated in Figure 43 below.

Figure 43 - Over-turn Angle Criterion (Raymer 2006 p. 232)

From geometry, and using the previously defined nomenclature, the equation for the over-turn

angle can be derived from the following diagram


Page 99 of 139

Figure 44 - Figure Describing Over-turn Criterion

Interim dimensions are calculated as follows:

Raymer (2006) states that the over-turn angle shall be no greater than 63o. The over-turn angle

criterion thus follows:

3.12.6 Landing Gear Position

The correct landing gear position is found by beginning with an initial configuration, and iterating the

calculations until all four criteria are met. The following design parameters are used in verifying the

initial configuration. These parameters are determined in the weight estimation section and the

centre of gravity section.

• Most forward CG 15.63 ft

• Most aft CG 15.9 ft

• Fuselage Diameter 8.3 ft


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• Fuselage Length 52.5 ft

• Landing Weight 19735 lb

• Height of tail above bottom of fuselage 8.3 ft

An iterative approach was used to find a landing gear position that meets all four criteria. The

following design point was proposed:

• Distance from nose to nose gear 4.1 ft

• Distance from nose to landing gear 18 ft

The following calculations verify that this design point meets all four criteria.


Page 101 of 139

3.12.7 Nose Weight Criterion

Thus, the nose weight criterion is satisfied.

3.12.8 Height Criterion

The height of the aircraft is determined by solving a quadratic equation, which can be derived by

simultaneously solving two equations that are a result of the geometry. The quadratic is as follows:

Using our parameters, and taking the only positive root:

The height is less than the height required by the height criterion, so the minimum height is used.

3.12.9 Roll-Over Criterion

Setting the angle to 55o, we can calculate the length of the base of the triangle. We denote this

dimension .

By using the properties of similar triangles, the half-width of the main landing gear can be

determined.

The roll-over criterion is automatically satisfied at this half-width (or greater) by the assumption.


Page 102 of 139

3.12.10 Over-Turn Angle Criterion

The over-turn angle criterion is satisfied (

3.12.11 Tip-Back Angle Criterion

The tip-back criterion is satisfied.

3.12.12 Summary

All four criteria are met, and the design point is verified as permissible

3.12.13 Landing Gear Loads

The following four loads are calculated from Roskam (2006). Each load has a 7% safety factor

included in accordance to FAR 25 regulations.


Page 103 of 139

3.12.14 Number, Type and Size of Tyres

The tricycle configuration has three contact points on the ground. Two wheels will be used at each

contact point to minimise the effect of a flat tyre. It is common to use two wheels at each point for

this reason (Torenbeek 1982). For a fire-fighting application, the heat involved will reduce the life of

the tyres, and the instances of flat tyres may be more numerous.

The weight that each tyre will need to support can be determined from the following equation:

For the nose wheel, static and dynamic loads need to be considered. The total load is divided by 1.4

as the nose wheel is permitted to carry more dynamic load than the rated static load (Raymer 2006).

Raymer (2006) recommends the use of Type III or Type VII tyres for traditional aircraft. Type III tyres

are used on aircraft with piston engines and Type VII tyres are used on aircraft with jet engines. Type

VII tyres will be used for this application and are selected from Raymer (pg 235, 2006).

Table 20 - Tyre Selection Table

Nose wheel tyres (2 of)

Size Speed

(knots)

Max

load (lb)

Inflation

(psi)

Max

Width

(in)

Max

Diameter

(in)

Rolling

Radius

(in)

Wheel

Diameter

(in)

Number

of plies

18x4.4 174 2100 100 4.45 17.90 7.9 10.0 12

Main gear tyres (4 of)

Size Speed

(knots)

Max

load (lb)

Inflation

(psi)

Max

Width

(in)

Max

Diameter

(in)

Rolling

Radius

(in)

Wheel

Diameter

(in)

Number

of plies

24x5.5 174 11500 355 5.75 24.15 10.6 14.0 16

3.12.15 Tyre Pressure Calculations

In order to calculate the tyre pressure, the contact area needs to be determined. The contact area

equation comes from Raymer (2006):


Page 104 of 139

The inflation pressure is given by the following formula (Raymer 2006):

In both cases, the inflation pressure is less than the maximum inflation pressure for the rated tyre.

Reducing the inflation pressures can increase the life of the tyres significantly. Raymer (2006) states

that by halving the internal pressure of the tyre, the life of the tyre improves six-fold. Increasing the

life of the tyres is beneficial because it reduces maintenance costs of the aircraft. Having a lower

internal pressure on the tyres allows the aircraft to take off and land at softer runways, which may

be required if landing in rural Australia.

3.12.16 Suspension Method and Requirements

Raymer (2009) states that the oleo-pneumatic shock absorber is the most common type of shock

absorbing mechanism. It is more efficient, more reliable, and has more energy damping compared

with less weight than the other shock absorbing devices. Oleo-pneumatic shock absorbers will be

used in the landing gear assembly on this aircraft. To calculate the required kinetic energy, a vertical

landing speed needs to be assumed. Raymer (2006) states that most aircraft require 10 ft/sec. A

landing speed of 10 ft/sec is required from the technical task. The required kinetic energy that the

shock absorbers are required to dissipate is calculated from the following formula (Torenbeek 1982):

The gear load factor is used in determining how much force passes from the gear to the airframe.

For a FAR 25 aircraft, the landing gear load factor is 2.0 (Arjomandi 2009). The stroke of the landing

gear can be calculated from the following formula (Raymer 2006). In this formula, refers to the

efficiency of the oleo-pneumatic shock absorber and refers to the efficiency of the tyre assembly.

S’ is the stroke of the shock absorber and is the stroke of the tyre. The stroke of the tyre is


Page 105 of 139

assumed to be the difference between its un-laden radius and its rolling radius. Efficiencies are

found from Raymer (2006):

A safety factor of one inch is added to this stroke by recommendation of Raymer (2006):

3.12.17 Length and Diameter of Landing Gear Struts

The external diameter of the oleo-pneumatic strut can be approximated by the following formula

(Raymer 2006). In this equation, the load on each oleo is the load on the main gears at touchdown

divided by the two oleo struts. In the main gear equation, the load is multiplied by the

aforementioned gear load factor. The pressure of the cylinder is 1800 psi.

The length of the struts is approximated as 2.5 multiplied by the required stroke.


Page 106 of 139

3.12.18 Nose-Wheel Steering and Castoring Dimensions

Nose wheel steering is accomplished by a separate mechanical link to a nose wheel steering control

in the cockpit. The nose wheel steering will not be linked to the rudder pedals to increase the

controllability of landing in windy or otherwise adverse conditions, such as those associated with

airports near a fire-affected area. For nose wheel steering to be made possible, a trail and rake need

to be introduced in the wheel design. The trail and the rake of the wheel are defined in accordance

with Figure 45 below:

Figure 45 - Figure Showing Trail and Rake of the Wheel (Raymer 2006)

The trail can be calculated from the following equation (Raymer 2006):

The rake for aircraft of this size should be 7o positive (Raymer 2006).


Page 107 of 139

3.12.19 Gear Retraction Geometry

The main gear will be mounted on a swivel mechanism in the fuselage. Due to the high wing design,

it is not practical for the main gear to retract into the wings. When the main gear is retracted, it

swivels rearwards towards the fuselage, and into a flush-mounted bay. Doors to this bay will be

pneumatically operated and mechanically linked to the retraction mechanism to ensure correct and

repeatable employment and deployment. The swivel retraction mechanism will be similar to that

seen on the Cutlass 172RG (a retractable landing gear version of the Cessna 172. A niche can be seen

aft of the main landing gear, painted maroon, where the main landing gear retracts during flight. The

nose wheel retracts into the fuselage.

A three bar linkage will be used to retract the wheels, as recommended by Raymer (2006). This will

ensure a compromise between good mechanical advantage and minimal space requirements. The

main gear will use a sliding pivot three bar linkage as shown in Figure 46 below.

Figure 46 - Sliding Bar Linkage (Raymer 2006)


Page 108 of 139

3.13 Isometric Views


Page 109 of 139

4 Weight and Balance Analysis

The aircraft takeoff weight of 19,735.34 lbs can be distributed to different groups and components

within the aircraft using statistics, except when the weight of actual components or systems is

available in which case actual weights are used. Weight distribution percentages, shown in Table 21,

suggested by Arjomandi (2009), were used as a guide due to the absence of more specific data in

Roskam (1985). System weight was distributed evenly between cockpit systems and payload

systems. Landing gear weight was distributed with 25% at the nose gear and 75% at the main

landing gear.

The location of the centre of gravity of each group is obtained from estimates provided by Roskam

(2005) and actual design locations. The weight breakdown of individual groups and the group

centres of gravity are given in Table 22. Four main weight configurations exist which involve various

combinations of fuel and payload. The centre of gravity of the aircraft in each mission configuration

is given in Table 23. The resulting centre of gravity envelope, plotted in Figure 47, shows that the aft

most centre of gravity possible during flight is in the operational empty weight configuration (i.e.

when the aircraft has dropped its payload, run out of fuel and is gliding).

Table 21 - Suggested Weight Distribution as Percentages (Eger 1983; Arjomandi 2009)

Component Percentage Reference weight

System 12-15% Takeoff weight

Fuselage 30-40% Structural weight

Wings 30-40% Structural weight

Empennage 5-10% Structural weight

Landing gear 10-15% Structural weight

Table 22 - Aircraft Weight Breakdown and Centre of Gravity Locations

Item Item CG position (ft from nose) Weight (lbs)

Fuselage 20.48 2292.9

Wing 15.03 2292.9

Empennage 49.29 286.6

Nose Gear 4.09 215.0

Main Gear 18 644.9

Wet engine components 4.08 363.0

Engine 4.08 629.0

Fixed cockpit equipment 10.8 986.8

Pilot & baggage 10.8 222.7

Trapped fuel 15.03 98.7


Page 110 of 139

Fuel 15.03 1897.6

Payload 15.70 8818.5

Fixed payload equipment 15.70 986.8

Table 23 - Centre of Gravity Locations for Various Payload and Fuel Configurations

Parameter Aircraft weight (lbs) Centre of gravity (%MAC)

Empty weight 8335 37.6%

Empty operational weight 9019 29.7%

Operational weight 10917 28.2%

Takeoff weight 19735 28.5%

Takeoff weight less fuel (i.e. all fuel

consumed without payload drop)

17838 29.3%

Figure 47 - Centre of Gravity Envelope


Page 111 of 139

5 Stability Analysis

Aircraft stability consists of static and dynamic stability in the longitudinal, lateral and directional

axis. Only longitudinal static stability was considered due to time constraints and course scope.

Longitudinal static stability is measured by the static margin, which is calculated from the

longitudinal centre of gravity and the neutral point according to Equation 20. The longitudinal centre

of gravity of the aircraft is given in the centre of gravity envelope in Figure 47.

The neutral point of the aircraft is the position at which the sum of all aerodynamic moments is zero.

The neutral point is dependent on the aerodynamic centre of the wings and fuselage as well as the

effect of the tail as described by Equation 21. A desired minimum static margin of 10% was selected

to provide longitudinal stability characteristics between that of early fighter aircraft (5%) and a

business jet (Learjet 35 - 13%) (Raymer 2006; Brandt, Stiles, Bertin & Whitford 2004). This static

margin provides a balance between the manoeuvrability needed for aircraft position and the

stability needed in the proximity of bushfire-generated turbulence.

Equation 20 - Static Margin

Equation 21 - Aircraft Neutral Point

The final horizontal tail area required, based upon the aft most flight centre of gravity, was

determined to be 69 ft2 in the longitudinal X-plot seen in Figure 48. This X-plot was generated by

considering tail areas in the region of the statistically sized horizontal tail area. Due to the small

changes in tail area, it was assumed that the centre of gravity location remained constant. The

calculations for the aircraft neutral point with the final horizontal tail area are seen in Appendix H.


Page 112 of 139

Figure 48 - Longitudinal X-plot for the Operational Empty Weight Configuration

The longitudinal static stability of the aircraft in each flight configuration is shown in Table 24 and

graphically represented in a combined centre of gravity envelope and neutral point diagram in

Figure 49. In-flight static margins vary between 10% (empty operational weight) and 11.6%

(operational weight). This minor change in static margin should provide consistent stability

characteristics throughout the in-flight centre of gravity envelope.

Table 24 - Longitudinal Stability in Each Flight Configuration

Configuration Static Margin

Empty operational weight 10%

Operational weight 11.6%

Takeoff weight 11.1%

Takeoff weight less fuel 10.4%


Page 113 of 139

Figure 49 - CG Envelope, Neutral Point and Static Margin for Each Flight Configuration


Page 114 of 139

6 Aerodynamic and Performance Analysis

The final conceptual fire fighting aircraft design involved a wing area of 535.56 ft2 and an engine

power of 1724.8 hp. An aerodynamic analysis was performed on the design to determine the lift to

drag ratios for the main mission phases. These new aerodynamic properties and engine data were

used to calculate a final estimated aircraft weight. This aircraft weight, in combination with the

known wing area and engine power, was used to determine whether the design point remained

within the met area of the matching diagram.

6.1 Aerodynamic Analysis

The lift to drag ratio of the aircraft in cruise and loiter phases can be calculated from the ratio of the

respective lift and drag coefficients. These values can then be used to perform a new weight

estimate.

6.1.1 Zero-Lift Drag Coefficient Calculation

The zero-lift drag coefficient can be recalculated using the actual wetted area ratio of the aircraft,

which was calculated from the CAD model. The wetted area ratio is given by Equation 22 to be 3.76.

Equation 22 - Final Design Wetted Area Ratio

The zero-lift drag coefficient was determined, using the original equivalent skin friction coefficient of

0.0055, to be 0.02060.

6.1.2 Required Lift Coefficients in Cruise and Loiter Phases

The mission profile requires a cruise speed of 341.75 ft/s (375 km/h) and a loiter speed of 181 ft/s.

At these speeds, the required wing lift coefficient was calculated, using Equation 23, to be 0.408 and

0.946 respectively.

Equation 23 - Lift Coefficient Required for Cruise


Page 115 of 139

6.1.3 Drag Coefficient in Cruise and Loiter Phases

During cruise and loiter, the aircraft is in the clean configuration. Hence, it has a zero-lift drag

coefficient of 0.0206. The drag coefficients for cruise and loiter were calculated, using Equation 24,

to be 0.0288 and 0.0646.

Equation 24: Drag coefficient

6.1.4 Lift to Drag Ratio Calculation

The lift to drag ratio for each phase was calculated by dividing the phase lift coefficient by the phase

drag coefficient. The lift to drag ratios were calculated to be 14.17 and 14.64 for cruise and loiter

respectively. These lift to drag ratios are compared to the assumed lift to drag ratios in Table 25. This

comparison shows that the aerodynamic performance of the aircraft in cruise has improved

significantly upon the assumed performance. The aerodynamic performance of the aircraft in loiter

has decreased slightly from the assumed performance. As the sensitivity analysis indicated that the

lift to drag ratio in cruise was more critical than the loiter lift to drag ratio, the increase in cruise

aerodynamic performance should result in a decreased aircraft weight.

Table 25 - Comparison of Assumed and Estimated Lift to Drag Ratios

Phase Assumed L/D Estimated L/D

Cruise 12.7 14.17

Loiter 14.67 14.64

6.2 Final Design Weight Estimate

A weight estimate of the final design, using the new values for lift to drag ratios ( 14.17 in cruise and

14.64 in loiter) and the actual engine specific fuel consumption (0.519 lbs/hp/hr in cruise and an

assumed value of 0.619 lbs/hp/hr in loiter), was performed to establish whether the existing aircraft

design could perform the required mission. Using the previously discussed MATLAB code, the weight

of the final design was estimated to be 19,721 lbs as seen in Figure 50. The expected reduction in

weight from the improvement in cruise aerodynamic performance has been offset by the increase in

the specific fuel consumption. The estimated weight of the final design is 14 lbs less than initial

weight estimate. It can be concluded, from this slight weight reduction, that the final design will be

able to perform the required mission profile.


Page 116 of 139

Figure 50 - Weight Estimate for the Final Design

6.3 Design Point Analysis

The design point of the final fire-fighting aircraft design is given by a wing loading of 36.8 and a

power loading of 11.4. A matching diagram for the final design, incorporating the new zero-lift drag

coefficient, is shown in Figure 51. The design point of the final conceptual design still lies within the

met area, and hence, it can be concluded that the conceptual design will meet the takeoff, landing,

climb, stall and cruise requirements of FAR25 and the technical task.


Page 117 of 139

Figure 51 - Final Matching Diagram.

The aerodynamic performance of the aircraft during cruise has improved compared to initial

estimates, whilst the loiter aerodynamic performance has essentially remained the same and the

specific fuel consumption of the final design has increased from the initial estimates. These

performance changes have produced a slightly lighter aircraft, indicating that the final aircraft design

can complete the required mission, and a design point that lies within the met area of the matching

diagram, indicating that the final aircraft design meets all sizing requirements. The design

successfully meets all required performance parameters.


Page 118 of 139

7 Conclusion

The conceptual design process for the Australian fire-fighting aircraft has resulted in an aircraft that

meets or exceeds the requirements specified by the technical task. All goals were considered during

each phase of the design process, resulting in an aircraft that is capable of effectively and efficiently

performing its intended mission. It was further found from the sizing and matching diagram that the

aircraft is adequately designed for its purpose. The aircraft is capable of landing at the majority of

civilian paved airfields. The simple wing planform design allows the aircraft to be stable, while

providing excellent ground visibility and low induced drag. Additionally, the turboprop propulsion

system offers greater economy than alternative systems for the given flight profile. Finally, the fire

retardant release system allows the aircraft to release retardant in four separate drops, increasing

the probability of successful fire suppression.

As a result of the success of the design, it has been determined that the fire-fighting aircraft can

meet Australian aerial fire-fighting requirements, and would be an attractive aircraft for the target

audience.


Page 119 of 139

8 References

AirTractor, 2009, Air Tractor 802F, Onley, U.S.A, viewed 10 May 2009,

<http://www.airtractor.com/Default.aspx?p=5967>

Airliner.net, 2009, The Canadair CL-215 & 415, U.S.A, viewed 10 May,

http://www.airliners.net/aircraft-data/stats.main?id=119

Australian Institute of Criminology, 2004, Crime prevention series, Canberra, viewed 1 April 2009, <

http://www.aic.gov.au/publications/crimprev/transport/air-t.html>.

Australian Natural Resources Atlas, 2008, Map Maker, Canberra, viewed 1 April

<http://www.anra.gov.au/mapmaker/mapservlet?app=anra>

Arjomandi, M 2009, Aircraft Design Lecture Notes, The University of Adelaide, Adelaide.

Aviation Jet Fuel Information, 2009, CSG, Computer Support Group, Inc. and CSGNetwork viewed 9

May 2009 http://www.csgnetwork.com/jetfuel.html.

Baxter, J 1984 Who Burned Australia?: The Ash Wednesday Fires. New English Library, Kent.

Billing, P 1983, Otways Fire No 22-1982/83 Aspects of Fire Behaviour, Fire Research Branch,

Melbourne.

Brandt, S, Stiles, R, Bertin, J & Whitford, R 2004, Introduction to Aeronautics: A Design Perspective,

American Institute of Aeronautics and Astronautics, USA.

Civil Aviation Safety Authority (CASA), 2008 Regulations and Policies, viewed 23 May 2009,

http://www.casa.gov.au/scripts/nc.dll?WCMS:HOMEPAGE:283465047:pc=PC_90001.

Component Weights, Stanford Education, 2009, viewed 9 May 2009,

http://adg.stanford.edu/aa241/structures/componentweight.html.

CSIRO 2008, The months of a fire season, ACT, viewed 1 April 2009, <

http://www.csiro.au/resources/FireSeasonMonths.html> .

Department of the Environmental, Water, Heritage and the arts 2001, Human Settlements Theme

Report, Canberra, viewed 1 April 2009,

http://www.environment.gov.au/soe/2001/publications/theme-

reports/settlements/settlements01.html>

Dowty Propellers, 2007, Federal Aviation Regulation Certification, USA Department of

Transportation, Washington D.C.

Dunn Aviation Australia, 2009, Our Aircraft, Australia, viewed 25 May 2009,

http://www.dunnav.com.au.

Honeywell TPE331-14, 2006, Honeywell, viewed 9 May 2009,

http://www.compairaviation.com/download/TPE331-14%20brochure.pdf


Page 120 of 139

Jackson, P 2008, Jane’s All the World’s Aircraft 2008-2009, Marston & Co publishing. London.

Javafoil 2009, Javafoil – Analysis of Aerofoils, USA, viewed 25 May 2009, http://www.mh-

aerotools.de/aerofoils/javafoil.htm .

Mahoney M 2009, Pointing an Instrument on an Airborne Platform, USA, viewed 24 May 2009,

http://mtp.jpl.nasa.gov/notes/pointing/Aircraft_Attitude2.png.

Operational Guidelines 2008, Approval of firebombing delivery subsystems – Fixed wing firebombing

aircraft (OPS-001), NAFC Standard, Melbourne.

Plucinski, M, Gould, J, McCarty, G, Hollis, J 2007, ‘The Effectiveness and Efficiency of Aerial

Firefighting in Australia’, Bushfire Cooperative Research Centre, Melbourne.

Raymer, D 2006, Aircraft Design: A Conceptual Approach, American Institute of Aeronautics and

Astronautics, Virginia.

RisingUp Aviation, 2006, Federal Aviation Regulations, viewed 23 May 2009,

http://www.risingup.com/fars/

Roskam, J 2004, Airplane Design Part I: Preliminary Sizing of Airplanes, DAR Corporation, USA.

Roskam, J 2005, Airplane Design Part II: Preliminary Configuration Designs and Integration of the

Propulsion System, DAR Corporation, USA.

Roskam, J 2005, Airplane Design Part III: Layout of Cockpit Fuselage, Wing and Empennage:

Cutaways and Inboard Profiles, DARCorporation, USA.

The Age, 2006 ‘Rain eases bushfire fears’ 29 January 2006, viewed 1 May 2009, 2006

http://www.theage.com.au/articles/2006/01/29/1138469595149.html

Torenbeek, E 1982, Synthesis of Subsonic Airplane Design, Kluwer Academic Publishers, London

Townsend, H 2009 ‘City swelters, records tumble in heat’, The Age, 7 February 2009, viewed 1 May

2009, http://www.theage.com.au/national/city-swelters-records-tumble-in-heat-20090207-

80ai.html .

UIUC 2008, UIUC Aerofoil Coordinates Database, USA, viewed 25 May 2009,

http://www.ae.uiuc.edu/m-selig/ads/coord_database.html.

WA Today, 2009, ‘Victorian bushfire statistics’, Australia, viewed 25 May 2009,

http://www.watoday.com.au .


Page 121 of 139


Page 122 of 139

Appendix A – Fire-fighting Aircraft Statistical

Analysis

Table 26 - Fire-fighting Aircraft Statistical Analysis

Aircraft Take-off

weight (lb)

Empty Weight

(lb) Payload (lb)

Maximum

Speed (knots) Range (nm)

Bronco OV-10 14444 6893 3000 281 576

TBM Avenger 17893 10545 2000 240 869

Douglass DC-3 25200 18300 5000 206 890

Grumman F7F-3

Tigercat 25720 16270 1000 400 1000

Grumen S2-

Tracker 26147 23435 - 260 1390

Grumman CDF S-

2A Tracker 27000 18315 6664 243 869

Bombardier

Canadair 415 37850 28400 13500 203 1310

Bombardier

Canadair CL-215 43500 26900 12000 160 1310

Consolidated

PB4Y-2 Privateer 65000 27485 8000 206 2450

Boeing B-17

Flying Fortress 65500 36135 6000 249 1738

Alenia C-27J

Spartan 70106 37479 19841 315 1160

Douglas DC-4 73000 43300 - 244 1897

Fairchild C-119 Boxcar 74000 40000 10000 257 1980

Beriev Be-200 Altair 83550 60850 655600 388 1000

Shinmaywa US-1A 94800 56505 30000 276 2060

P3-Orion 142000 77200 - 411 2070 McDonnell Douglas DC-7 143000 72763 24990 353 4001

C-130 Hercules 155000 83000 45000 348 2835 JRM Mars 165000 75573 32000 192 4300 McDonnell Douglas DC-10-10

430000 240171 99960 530 3302

Boeing 747 833000 392800 200287 510 6700 Antonov An-2 'Colt' - 7300 3307 139 485


Page 123 of 139

ROKS-Aero T-101 Grach - - 3086 162 685

Appendix B –Statistical Analysis Relevant

Aircraft

Aircraft

Empty

Weight

(lb)

Take-off

weight

(lb)

Cruise

Speed

(knots)

Stall

Speed

(knots)

Range

(nm)

G-164B Super B Turbine 3150.00 7020.00 113.00 53.85 172.00

Pac Cresco 2950.00 8250.00 140.00 45.00 364.00

PZL-Mielec_M-18_Dromader 5975.00 11700.00 100.00 59.00 540.00

Antonov An-2 7300.00 12000.00 100.00 26.00 456.00

Air Tractor AT602 5600.00 12500.00 126.00 86.00 538.77

Embraer EMB 110

Bandeirante 7837.00 12500.00 184.00 76.00 1060.00

Handley Page Jetstream 9613.00 15332.00 230.00 86.00 680.00

Air Tractor AT802 6400.00 16000.00 169.00 93.00 695.00

CASA C-212 9680.00 17600.00 170.00 75.00 237.00

Embraer EMB 120 Brasilia 15655.00 26378.00 300.00 55.00 850.00

Dornier 328 19670.00 30840.00 335.00 93.08 1000.00

CASA C-235 21605.00 33290.00 245.00 107.00 1549.25

Grumman G-159 Gulfstream I 21900.00 35100.00 250.00 90.00 2206.00

Saab 340B 17945.00 35245.00 250.00 115.00 935.00

Sukhoi Su-80 34241.00 38045.00 232.00 95.00 702.00

Antonov An-140 28240.00 42220.00 250.00 95.38 745.00

Convair CV-240 25445.00 42500.00 243.00 86.92 1042.00

Bombardier Canadair CL-215 26900.00 43500.00 156.00 92.00 1310.00

Bombardier Canadair 415 28400.00 43850.00 180.00 68.00 1319.11


Page 124 of 139

Appendix C – Calculated Fuel Fractions

The fuel fractions for cruise and loiter phases are more sensitive to design requirements and must be

calculated. Additionally the fuel fraction for phase 8 must be corrected to account for the payload

drop.

Phase 5: Cruise

During the previous phase (phase 4: climb) part of the radius was covered. This range credit is

calculated from the time to climb and the horizontal velocity during climb. Time to climb is

calculated from Equation 25.

Equation 25: Time to climb for phase 4

The horizontal velocity during climb is estimated to be the average of the cruise velocity and the

takeoff velocity in Equation 26. Takeoff velocity is approximated as 1.1Vstall.

Equation 26: Horizontal velocity during climb

The horizontal distance covered during climb is therefore given by Equation 27.

Equation 27: Horizontal distance covered during phase 4

The distance to be covered during cruise is given in Equation 28.

Equation 28: Cruise distance for phase 5.

The fuel fraction for phase 9 is given by Equation 29.

Equation 29: Fuel fraction for phase 5 (Roskam 2005).

Phase 7: Loiter and payload drop


Page 125 of 139

The aircraft drops its entire payload during a 20 minute loiter period. It is assumed for the purposes

of calculation that the entire payload is dropped instantaneously at the end of this loiter phase. This

assumption provides a conservative estimate as the aircraft carries a greater weight for the later

stages of the loiter period. The fuel fraction for loiter is given in Equation 30.

Equation 30: Loiter fuel fraction (Roskam 2005).

Payload drop

The payload weight ratio must be calculated to determine the appropriate correction factor to be

applied to the fuel fraction for phase 8. The fuel fraction to this point of the mission is given by

Equation 31.

Equation 31: Mission fuel fraction for phases 1 to 7.

The weight at this point in the mission can then be calculated from an assumed takeoff weight. This

results in an iterative. This process was started with an assumed takeoff weight of 16000 lbs

(maximum takeoff weight for the Air Tractor 802F (Air Tractor 2007) and continued until the

calculated takeoff weight was within 10 lbs of the assumed value. The converged takeoff weight was

19735 lbs. The MATLAB code used for this iterative process can be found in Appendix E. Using the

mission fuel fraction in Equation 31 the aircraft weight immediately before payload deployment is

calculated in Equation 32.

Equation 32: Aircraft weight before payload drop.

The aircraft weight after payload deployment and the payload weight ratio (PWR) is given by

Equation 34, which is used to correct the fuel fraction for climb in phase 8.

Equation 33: Aircraft weight after payload drop.

Equation 34: Payload weight ratio.

Phase 8: Climb fuel fraction correction

The fuel fraction for phase 8: climb given in Table 6 must be corrected for the payload drop using the

payload weight ratio.


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Equation 35: Correction fuel fraction for phase 8: climb (Roskam 2005).

Phase 9: Cruise

During the previous phase (phase 8: climb) part of the radius was covered. This range credit is

calculated from the time to climb and the horizontal velocity during climb. Time to climb is

calculated from Equation 36.

Equation 36: Time to climb for phase 8

The horizontal velocity during climb is estimated to be the average of the cruise velocity and the

takeoff velocity in Equation 37. Takeoff velocity is approximated as 1.1Vstall.

Equation 37: Horizontal velocity during climb

The horizontal distance covered during climb is therefore given by Equation 38.

Equation 38: Horizontal distance covered during phase 8

The distance to be covered during cruise is given in Equation 39.

Equation 39: Cruise distance for phase 9.

The fuel fraction for phase 9 is given by Equation 40.

Equation 40: Fuel fraction for phase 9 (Roskam 2005).

Mission Fuel Fraction

The mission fuel fraction for this aircraft is calculated by multiplying the fuel fractions for each

individual phase as shown in Equation 41.


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Equation 41: Mission fuel fraction.

Appendix D – Sensitivity Calculations

Preliminary calculations

The sensitivity of takeoff weight to other parameters is given by Equation 42.

Equation 42: Generalised takeoff weight sensitivity equation.

Where B is the value from the technology diagram equation, C is defined by Equation 43 and D is

defined by Equation 44. In addition to these parameters, the variable F, defined in Equation 45, is

required to calculate the majority of sensitivities.

Equation 43: Definition of the variable C.

Equation 44: Definition of the variable D.

Equation 45: Definition of the variable F.


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Sensitivity to payload weight and crew weight

The sensitivity to payload weight and to crew weight is calculated with and .

Therefore, the sensitivity is given in Equation 46. Similarly, the sensitivity to crew weight is also

1.789 lbs/lbs.

Equation 46: Takeoff weight sensitivity to payload weight.

Sensitivity to empty weight

Sensitivity to empty weight is calculated from the technology diagram equation. Differentiating with

respect to empty weight gives Equation 47.

Equation 47: Takeoff weight sensitivity to empty weight. Equation is identical for sensitivity to crew weight.

Sensitivity to radius (i.e. the range of one cruise leg)

Sensitivity to mission radius is calculated from a Breguet partial, Equation 48, and Equation 49.

Equation 48: Breguet partial for cruise range sensitivity.

Equation 49: Takeoff weight sensitivity to cruise range sensitivity.

Sensitivity to cruise power specific fuel consumption

Sensitivity to cruise power specific fuel consumption is calculated from a Breguet partial, Equation

50, and Equation 51.

Equation 50: Breguet partial for cruise power specific fuel consumption.


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lbs/lbs/hp/hr

Equation 51: Takeoff weight sensitivity to cruise power specific fuel consumption.

Sensitivity to cruise propeller efficiency

Sensitivity to cruise propeller efficiency is calculated from a Breguet partial, Equation 52, and

Equation 53.

Equation 52: Breguet partial for cruise propeller efficiency.

Equation 53: Takeoff weight sensitivity to cruise propeller efficiency.

Sensitivity to cruise lift to drag ratio

Sensitivity to cruise lift to drag ratio is calculated from a Breguet partial, Equation 54, and Equation

55.

Equation 54: Breguet partial for cruise lift to drag ratio.

Equation 55: Takeoff weight sensitivity to cruise lift to drag ratio.

Sensitivity to endurance

Sensitivity to endurance is calculated from a Breguet partial, Equation 56, and Equation 57.

Equation 56: Breguet partial for loiter endurance.

Equation 57: Takeoff weight sensitivity to loiter endurance.

Sensitivity to loiter power specific fuel consumption


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Sensitivity to loiter power specific fuel consumption is calculated from a Breguet partial, Equation

58, and Equation 59.

Equation 58: Breguet partial for loiter power specific fuel consumption.

Equation 59: Takeoff weight sensitivity to loiter power specific fuel consumption.

Sensitivity to loiter propeller efficiency

Sensitivity to loiter propeller efficiency is calculated from a Breguet partial, Equation 60, and

Equation 61.

Equation 60: Breguet partial for loiter propeller efficiency.

Equation 61: Takeoff weight sensitivity to loiter propeller efficiency.

Sensitivity to loiter velocity

Sensitivity to loiter velocity is calculated from a Breguet partial, Equation 62, and Equation 63.

Equation 62: Breguet partial for loiter velocity.

Equation 63: Takeoff weight sensitivity to loiter velocity.

Sensitivity to loiter lift to drag ratio

Sensitivity to loiter lift to drag ratio is calculated from a Breguet partial, Equation 64, and Equation

65.


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Equation 64: Breguet partial for loiter lift to drag ratio.

Equation 65: Takeoff weight sensitivity to loiter lift to drag ratio.


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Appendix E – MATLAB Code for Takeoff

Weight Estimation and Sensitivity Analysis

%Aircraft Design Project

%Mission Fuel Weight Calculations

close all

clear all

clc

%Aircraft Requirements

Range=335.5404; %sm this is actually the radius i.e. a 1080km total range is broken up into 2 540km flights

E=20/60; %time for a single bombing run min->hrs

Wp=4000*2.204623; %payload weight in pounds

Wc=(86+15)*2.2046226218; %crew weight in pounds

M_reserve=0.06; %reserve fuel fraction

M_unuseable=0.005; %unuseable fuel fraction

%Statistical Constants

V_stall=82.5*1.68781; %fps 82.5 knots

V_cr=375*0.9113444; %fps 350 km/h

h_cr=14000; %ft

h_bomb=70; %ft

RC=850/60; %fpm->fps

eta_p_bomb=0.82;

eta_p_cr=0.82;

Cp_bomb=0.571;

Cp_cr=0.471;

LD_cr=12.7;

A=-0.8126; %values from technology diagram

B=1.2966; %values from technology diagram

%Derived Values

LD_bomb=LD_cr/0.866;

V_takeoff=1.1*V_stall;

V_cl=(V_cr+V_takeoff)/2;

V_bomb=1.3*V_stall*0.6818181818;

V_cr=V_cr*0.6818181818;

%Takeoff Weight Calculations

%Mission Fuel Fraction

%Fuel Fractions from Roskam 2005

W1W0= 0.996; %Phase 1 (Engine start and warm-up) for twin engine aircraft [Roskam 2005]

W2W1=0.995; %Phase 2 (Taxi) for twin engine aircraft [Roskam 2005]

W3W2=0.996; %Phase 3 (Takeoff) for twin engine aircraft [Roskam 2005]

W4W3=0.998; %Phase 4 (Climb) for twin engine aircraft [Roskam 2005]


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W6W5=0.999; %Phase 6 (Descend)

W10W9= 0.999;%Phase 10 (Descent)

W11W10=0.998; %Phase 11 (Landing, taxi and shut down)

%Phase 4 Climb: Range Credit

t_cl4=h_cr/RC; %Time to climb to cruise altitude

R_cl4=t_cl4*V_cl*0.00018939393939; %Range covered during climb in ft->sm

%Phase 5 Cruise

R_cr=Range-R_cl4; %Range to be covered during cruise

W5W4=exp((-R_cr*Cp_cr)/(375*eta_p_cr*LD_cr)); %Fuel fraction for phase 5

%Estimated Aircraft Weight

Wto=20000; %estimated aircraft weight

weight_est=0; %initialising this variable

while abs(weight_est-Wto)>10 %loop to calculate the actual takeoff weight from an estimated value

weight_est=Wto;

%Phase 7: Bombing Phase

W7W6= exp((-E*Cp_bomb*V_bomb)/(375*eta_p_bomb*LD_bomb)); %loiter weight fraction

Mff1_7=W1W0*W2W1*W3W2*W4W3*W5W4*W6W5*W7W6; %fuel fraction for phases 1-7

weight7_fuel=Mff1_7*weight_est; %aircraft weight at the end of the bombing run due to fuel consumption

weight_bombing=weight7_fuel-Wp;

PWR=weight_bombing/weight7_fuel;

%Phase 8 Climb Range Credit

W8W7_base = 0.998;%Uncorrected Phase 8 (Climb)

W8W7= (1-(1-W8W7_base)*PWR); %Fuel fraction corrected for water drop

t_cl8=(h_cr-h_bomb)/RC; %Time to climb to cruise altitude

R_cl8=t_cl8*V_cl*0.00018939393939; %Range covered during climb in ft->sm

%Phase 9 (Cruise)

R_cr=Range-R_cl8; %Range to be covered during cruise

W9W8=exp((-R_cr*Cp_cr)/(375*eta_p_cr*LD_cr)); %Fuel fraction for phase 5

%Mission Fuel Fraction

Mff=W1W0*W2W1*W3W2*W4W3*W5W4*W6W5*W7W6*W8W7*W9W8*W10W9*W11W10; %mission

fuel fraction

%Takeoff Weight

if 1

We=1:0.1:30000;

Wto_tech=10^A*We.^B;

Wto_fuel=(We+Wc+Wp)/(1-(1+M_reserve)*(1-Mff)-M_unuseable);

plot(We,Wto_tech,We,Wto_fuel)

xlabel('W_e')

ylabel('W_t_o')

if 1

difference=abs(Wto_tech-Wto_fuel);

intersect=min(difference);

for i=1:length(We)

if difference(i)==intersect

Wto=Wto_tech(i);

We=We(i);

end

end

end


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end

end

Wf=(1+M_reserve)*(1-Mff)*Wto+M_unuseable*Wto;

fprintf('The takeoff weight is: %f and the empty weight is %f',Wto,We)

if 1

%Sensitivity Analysis

C=1-(1+M_reserve)*(1-Mff)-M_unuseable;

D=Wp+Wc;

F=-B*Wto^2*(C*Wto*(1-B)-D)^-1*(1+M_reserve)*Mff;

%Takeoff weight sensitivity to payload weight

dWto_dWp=B*Wto*(D-C*(1-B)*Wto)^-1

%Takeoff weight sensitivity to crew weight

dWto_dWc=B*Wto*(D-C*(1-B)*Wto)^-1

%Takeoff weight sensitivity to empty weight

dWto_dWe=(B*Wto)/(10^((log10(Wto)-A)/B))

%Sensitivities for the cruise phases

%Takeoff weight sensitivity to range

d_R_dR=Cp_cr*(375*eta_p_cr*LD_cr)^-1;

dWto_dR=F*d_R_dR

%Takeoff weight sensitivity to Cp_cr

d_R_dCp_cr=Range*(375*eta_p_cr*LD_cr)^-1;

dWto_dCp_cr=F*d_R_dCp_cr

%Takeoff weight sensitivity to eta_p_cr

d_R_deta_p_cr=-Range*Cp_cr*(375*eta_p_cr^2*LD_cr)^-1;

dWto_deta_p_cr=d_R_deta_p_cr*F

%Takeoff weight sensitivity to L/D_cr

d_R_dLD_cr=-Range*Cp_cr*(375*eta_p_cr*LD_cr^2)^-1;

dWto_dLD_cr=d_R_dLD_cr*F

%Sensitivities for the bombing phase

%Takeoff weight sensitivity to E

d_E_dE=V_bomb*Cp_bomb*(375*eta_p_bomb*LD_bomb)^-1;

dWto_dE=d_E_dE*F

%Takeoff weight sensitivity to V_bomb

d_E_dV_bomb=E*Cp_bomb*(375*eta_p_bomb*LD_bomb)^-1;

dWto_dV_bomb=d_E_dV_bomb*F

%Takeoff weight sensitivity to Cp_bomb

d_R_dCp_bomb=Range*(375*eta_p_bomb*LD_bomb)^-1;

dWto_dCp_bomb=F*d_R_dCp_bomb

%Takeoff weight sensitivity to eta_p_bomb

d_R_deta_p_bomb=-Range*Cp_cr*(375*eta_p_bomb^2*LD_bomb)^-1;

dWto_deta_p_bomb=d_R_deta_p_bomb*F

%Takeoff weight sensitivity to L/D_bomb

d_R_dLD_bomb=-Range*Cp_bomb*(375*eta_p_bomb*LD_bomb^2)^-1;

dWto_dLD_bomb=d_R_dLD_bomb*F

end


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Appendix F - Honeywell TPE331-14GR

Specifications

Table 27 - Honeywell TPE331-14GR Specifications (Jackson 2008) and (Honeywell TPE331-14 2006)

Power output (shp) 1960

Dry Weight (lb) 620

SFC 0.519

ESFC 0.497

RPM gas gen 35645

Rotation direction Clockwise from rear

PWR/WT ration 2.66

Pressure Ratio 11.2

Airflow (lb/sec) 12.0

Fuel Jet A, Jet B, Jet A-1, JP-4, JP-5, JP-8 JP8+100

Oil Type Mil-L-23699C, Type II Mil-L-7808D, F & G, Type I

Start capability SL–20,000 ft

-65°F + 130°F Operational limits SL–31,000 ft

-85°F + 130°F Specification no 21-8355

*to 75°F (to 97°F APR) Propeller drive Single shaft

Arrangement Two centrifugal stages on the same shaft.


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Appendix G – Flap Sizing Data

Table 28 - Flap Sizing Table

Aircraft Flap Chord Ratio

(%)

Flap Location from

Fuselage (Inboard) (%)

Flap Location from

Fuselage (Outboard) (%)

Air Tractor AT-502B 20.00 8.30 51.90

Zlin Z 37T Agro Turbo 20.50 0.00 64.30

PZL Mielec M-15 23.30 0.00 56.00

ICA IAR-827A 18.30 12.70 61.80

Aero Boero 260 Ag 22.40 9.90 47.60

WTA (Piper)PA-36 New

Brave 20.00 8.70 51.75

Sukhoi Su-38 21.80 0.00 55.20

M-18 Dromader 28.50 13.00 58.50

EMB 202 18.70 0.00 58.20

PZL - 106B Kruk 23.00 8.29 50.69

FU-24-954 15.70 0.00 62.50

Average 21.11 5.54 56.22

Table 29 - Aileron Sizing Table

Aircraft Aileron Chord Ratio

(%)

Aileron Location

from Fuselage

(Inboard) (%)

Aileron Location

from Fuselage

(Outboard) (%)

Air Tractor AT-502B 25.00 50.30 89.00

Zlin Z 37T Agro Turbo 29.00 63.50 100.00

PZL Mielec M-15 25.00 56.00 95.70

ICA IAR-827A 21.30 61.80 83.60

Aero Boero 260 Ag 15.80 51.20 94.00

WTA (Piper)PA-36

New Brave 30.00 52.72 92.26

Sukhoi Su-38 21.80 55.20 100.00

M-18 Dromader 27.00 57.80 92.40

EMB 202 18.70 58.60 89.70

PZL - 106B Kruk 23.00 52.80 95.00

FU-24-954 19.70 63.00 100.00

Average 23.30 56.63 93.79


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Appendix H – Neutral Point Calculations

The neutral point of the aircraft is calculated in Equation 66.

Equation 66: Neutral point equation.

Where:

• is the normalised aerodynamic centre of the wing and fuselage combination

relative to the wing MAC leading edge. calculated in Equation 67 to be 0.125.

• is the normalised aerodynamic centre of the horizontal tail relative to the wing MAC

leading edge (4.53).

• is the tail lift curve slope (0.0699).

• is the wing lift curve slope (0.0954).

• is the wing reference area (535.56 ft2).

• is the tail horizontal area (69 ft2 determined by longitudinal X-plot).

• is the rate of change in downwash angle with angle of attack. Calculated in Equation 68

to be 0.3036.

The aerodynamic centre of the wing is at 25% of the mean aerodynamic chord. The aerodynamic

centre of the combined wing and fuselage may be calculated from (Brandt et al. 2004).

Equation 67: Aerodynamic centre of the wing and fuselage.

Where:

• is the wing aerodynamic centre (0.25*8.11 ft)


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• is the length of the fuselage (52.5 ft)

• is the width of the fuselage (8.75 ft)

• is the distance from the nose of the aircraft to the wing aerodynamic centre (14.05 ft)

• is the wing reference area (535.56 ft2)

• is the wing lift curve slope (0.954)

• is the mean aerodynamic chord (8.11 ft)

The rate of change in downwash angle with angle of attack is calculated from Equation 68.

Equation 68: Rate of change in downwash angle with angle of attack.

Where:

• is the wing lift curve slope (0.954)

• is the average geometric chord (8.11 ft)

• is the aspect ratio (8).

• is the distance between the quarter chord of the wing and the quarter chord of the tail

(35 ft).

• is the taper ratio (1).

• is the vertical distance between the plane of the wings and the horizontal tail (1 ft).

• is the wing span (65.46 ft).


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Appendix I – Three View Drawings

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