Coast Watch UAV

THE UNIVERSITY OF ADELAIDE

SCHOOL OF MECHANICAL ENGINEERING

Aircraft Design Coast Watch UAV

GROUP 7

Kelly Balnaves

Bradley Cook

Alex Horstmann

Ryan Middleton

Christian Rogers

Aircraft Design Project Group 7

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Criteria Mark (Total – 100)

Project Definition /10

Research Activities /15

Technical Calculations /25

Drawings /25

Format of the Report /10

Novelty of the Solution /15

Name Group Mark Peer Assessment Total Mark

Kelly Balnaves

Bradley Cook

Alex Horstmann

Ryan Middleton

Christian Rogers

Name Id Number Signature

Kelly Balnaves 1132985

Bradley Cook 1133395

Alex Horstmann 1131838

Ryan Middleton 1133404

Christian Rogers 1130940


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1 Executive Summary

This project presents the conceptual design and sizing of a Coast-Guard UAV.

The UAV was designed to monitor the coastal waters of Australia without utilising

expensive manned vehicles such as boats and helicopters. The UAV was

designed for use in remote and populated areas alike with a catapult launch and

a hook and cable landing. The aircraft is primarily designed for loiter at an altitude

of 1000ft. This report contains a statistical analysis, preliminary calculations,

configuration design and technical drawings. The final aircraft has a take-off

weight of 172lbs, a range of 600km and a cruise speed of 100km/hr.


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1 EXECUTIVE SUMMARY..................................................................................................................2

2 TABLE OF FIGURES.........................................................................................................................5

3 TABLE OF TABLES...........................................................................................................................6

4 TABLE OF EQUATIONS...................................................................................................................7

5 INTRODUCTION................................................................................................................................8

5.1 AIMS .............................................................................................................................................8

5.2 SCOPE ...........................................................................................................................................8

5.3 BACKGROUND INFORMATION........................................................................................................8

5.4 SIGNIFICANCE ...............................................................................................................................9

6 TECHNICAL TASK .........................................................................................................................10

6.1 STANDARD REQUIREMENTS ........................................................................................................10

6.2 PERFORMANCE PARAMETERS......................................................................................................10

6.3 TECHNICAL LEVEL ......................................................................................................................13

6.4 ECONOMICAL PARAMETERS ........................................................................................................13

6.5 POWER PLANT TYPE AND REQUIREMENTS....................................................................................14

6.6 MAIN SYSTEM PARAMETERS REQUIREMENTS ..............................................................................14

6.7 RELIABILITY AND MAINTAINABILITY..........................................................................................15

7 STATISTICAL ANALYSIS..............................................................................................................16

7.1 WING SPAN/AIRCRAFT LENGTH..................................................................................................18

7.2 TAKE-OFF METHODS ...................................................................................................................19

7.3 LANDING METHODS ....................................................................................................................20

7.4 SUMMARY OF BENCHMARKS.......................................................................................................21

8 CONCEPT SKETCHES....................................................................................................................22

9 WEIGHT ESTIMATION..................................................................................................................29

10 SENSITIVITY ANALYSIS...............................................................................................................33

11 AIRCRAFT SIZING .........................................................................................................................37

11.1 STALL SPEED SIZING ...................................................................................................................37

11.2 TAKE OFF DISTANCE SIZING.........................................................................................................37

11.3 LANDING DISTANCE SIZING .........................................................................................................38

11.4 CLIMB SIZING..............................................................................................................................39

11.4.1 FAR 23.65 Rate of Climb Sizing .......................................................................................41

11.4.2 FAR 23.65 Climb Gradient Sizing ....................................................................................41


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11.4.3 FAR 23.77 Climb Gradient Sizing ....................................................................................42

11.5 CRUISE SIZING ............................................................................................................................43

11.6 OVERALL SIZING CHART.............................................................................................................43

12 OVERALL CONFIGURATION DESIGN......................................................................................46

12.1 FUSELAGE DESIGN ......................................................................................................................46

12.2 AEROFOIL SELECTION .................................................................................................................48

12.3 WING DESIGN AND POSITIONING ................................................................................................52

12.4 TAIL DESIGN ...............................................................................................................................55

12.5 CONTROL SURFACE SIZING .........................................................................................................60

12.6 PROPULSION SYSTEM ..................................................................................................................61

12.7 PROPULSION INTEGRATION .........................................................................................................66

12.7.1 General Configuration......................................................................................................66

12.7.2 Position .............................................................................................................................68

12.8 TAKE OFF METHODS ...................................................................................................................70

12.9 LANDING METHODS ....................................................................................................................71

12.10 DETACHABLE EQUIPMENT BAY ..................................................................................................73

13 WEIGHT AND STABILITY ANALYSIS .......................................................................................74

13.1 WEIGHT ANALYSIS .....................................................................................................................74

13.2 STABILITY ANALYSIS ..................................................................................................................78

14 PERFORMANCE ANALYSIS AND CONCLUSION ...................................................................82

15 REFERENCES...................................................................................................................................84

16 DRAWINGS .......................................................................................................................................86


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2 Table of Figures FIGURE 6-1: MISSION PROFILE.................................................................................................................12

FIGURE 7-1: EMPTY WEIGHT VERSUS TAKE OFF WEIGHT .........................................................................17

FIGURE 8-1: CONCEPT 1 SKETCH .............................................................................................................22



FIGURE 8-4: CONCEPT 4, SKETCH 1 .........................................................................................................25




FIGURE 9-1: WEIGHT ESTIMATION GRAPH ...............................................................................................32

FIGURE 11-1: SIZING CHART ...................................................................................................................44

FIGURE 11-2 : SIZING CHART WITHOUT LANDING AND TAKEOFF ............................................................45

FIGURE 12-1 FINENESS RATIO TERMS .....................................................................................................47

FIGURE 12-2: FINENESS RATIO FOR SUBSONIC AIRCRAFT .......................................................................47

FIGURE 12-3: AEROFOIL LIFT CURVES (MODELFOIL) .............................................................................50

FIGURE 12-4 – NACA 4415 (MODELFOIL) ..............................................................................................51

FIGURE 12-5 – NACA 0012 PROFILE (MODELFOIL)................................................................................52

FIGURE 12-6: TWIN BOOM TAIL CONFIGURATION (HTTP://AEROWEB.LUCIA.IT/RAP/PARIS97)................55

FIGURE 12-7: VERTICAL STABILISER DIMENSIONS..................................................................................59

FIGURE 12-8: HORIZONTAL STABILISER DIMENSIONS.............................................................................60

FIGURE 12-9: GENERAL TRACTOR AND PUSHER CONFIGURATIONS (RAYMER, 1992) ..............................66

FIGURE 12-10: ENGINE POSITIONS FOR PUSHER CONFIGURATION (RAYMER, 1992).................................69

FIGURE 12-11: CATAPULT LAUNCH SYSTEM FOR SURVEILLANCE UAV.................................................70

FIGURE 12-12: HOOK LANDING SYSTEM.................................................................................................72

FIGURE 12-13: DETACHABLE EQUIPMENT BAY.......................................................................................73

FIGURE 13-1: CENTRE OF GRAVITY ENVELOPE .......................................................................................77


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3 Table of Tables TABLE 7-1: TAKE OFF WEIGHTS FOR SIMILAR AIRCRAFT .........................................................................17

TABLE 7-2: WING SPAN AND AIRCRAFT LENGTH .....................................................................................18

TABLE 7-3: TAKE OFF METHODS .............................................................................................................19

TABLE 7-4: LANDING METHODS ..............................................................................................................20

TABLE 7-5: SUMMARY OF STATISTICAL ANALYSIS ..................................................................................21

TABLE 9-1: TIME, FUEL CONSUMPTION AND POWER FOR EACH STAGE ....................................................30

TABLE 9-2: FUEL WEIGHT AND WEIGHT RATIO FOR EACH STAGE.............................................................31

TABLE 10-1: VALUES FOR SENSITIVITY EQUATIONS (FROM AIRPLANE DESIGN BY J. ROSKAM).............34

TABLE 10-2: VALUES USED FOR CALCULATING ‘C’ AND ‘D’...................................................................35

TABLE 10-3: REQUIRED VALUES FOR SENSITIVITY ANALYSIS..................................................................35

TABLE 10-4: SENSITIVITIES TO THE MAIN PARAMETERS.........................................................................36

TABLE 11-1: TAKE-OFF DISTANCE SOLUTION ..........................................................................................38

TABLE 11-2: LANDING DISTANCE SIZING.................................................................................................39

TABLE 11-3: LIFT COEFFICIENT FOR ALL CONFIGURATIONS ....................................................................39

TABLE 11-4: DRAG POLAR VALUES .........................................................................................................40

TABLE 11-5: ASPECT RATIO AND OSWALD EFFICIENCY FACTOR .............................................................40

TABLE 11-6: DRAG POLAR EQUATIONS FOR FAR23 CLIMB SIZING ..........................................................40

TABLE 11-7: FINAL CLIMB SIZING VALUES ..............................................................................................42

TABLE 11-8: CRUISE SIZING VALUES .......................................................................................................43

TABLE 11-9: FINAL SIZING VALUES .........................................................................................................45

TABLE 12-1: COMMON AIRFOIL SECTIONS FOR UAV AIRCRAFT (LEDNICER, 2007).................................49

TABLE 12-2: AIRCRAFT VOLUME COEFFICIENT DATA (AVALAKKI ET AL, 2007) ....................................57

TABLE 12-3: MAIN WING PROPERTIES ....................................................................................................57

TABLE 12-4: ENGINE SELECTION TABLE..................................................................................................66

TABLE 13-1: WEIGHT BREAKDOWN ........................................................................................................74

TABLE 13-2: AIRFOIL LIFT-CURVE SLOPES (MODELFOIL) ......................................................................78


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4 Table of Equations EQUATION 7-1: STATISTICAL ANALYSIS EQUATION (ROSKAM, 1994) .....................................................16

EQUATION 9-1: TAKE OFF WEIGHT EQUATION USING STATISTICS ............................................................29

EQUATION 9-2: GENERAL TAKE OFF WEIGHT EQUATION .........................................................................29

EQUATION 9-3: FUEL WEIGHT EQUATION ................................................................................................30

EQUATION 9-4: MASS FUEL FRACTION ....................................................................................................30

EQUATION 9-5: MISSION STAGE FUEL WEIGHT ........................................................................................31

EQUATION 9-6: FINAL TAKE OFF WEIGHT EQUATION ...............................................................................31

EQUATION 10-1: TAKE-OFF WEIGHT SENSITIVITIES FOR RANGE AND ENDURANCE CASES .....................33

EQUATION 10-2: EQUATION TO CALCULATE F ........................................................................................33

EQUATION 10-3: EQUATION FOR ‘C’ CALCULATION ...............................................................................34

EQUATION 10-4: EQUATION FOR ‘D’ CALCULATION ...............................................................................34

EQUATION 11-1: STALL SPEED EQUATION ...............................................................................................37

EQUATION 11-2 TAKE-OFF SIZING EQUATION..........................................................................................38

EQUATION 11-3: ZERO LIFT DRAG COEFFICIENT ESTIMATION..................................................................39

EQUATION 11-4: DRAG POLAR EQUATION ...............................................................................................40

EQUATION 11-5: FAR23 ROC SIZING EQUATION.....................................................................................41

EQUATION 11-6: FAR23.65 CG SIZING EQUATION..................................................................................41

EQUATION 11-7: FINAL EQUATION FOR FAR23.67 SIZING ......................................................................42

EQUATION 12-1: FINENESS RATIO ...........................................................................................................46

EQUATION 12-2: VERTICAL TAIL VOLUME COEFFICIENT .........................................................................56

EQUATION 12-3: HORIZONTAL TAIL VOLUME COEFFICIENT.....................................................................56

EQUATION 12-4: VERTICAL TAIL AREA ...................................................................................................58

EQUATION 12-5: HORIZONTAL TAIL AREA...............................................................................................58

EQUATION 12-6: PROPELLER TIP VELOCITY EQUATION (RAYMER, 2006) ................................................62

EQUATION 12-7: REARRANGED PROPELLER TIP VELOCITY EQUATION (RAYMER, 2006)..........................62

EQUATION 13-1: LIFT CURVE EQUATION CONVERSION (NELSON, 1989)..................................................78

EQUATION 13-2: DOWNWASH ANGLE......................................................................................................79

EQUATION 13-3: DOWNWASH ANGLE WITH VARIATION IN ANGLE OF ATTACK (NELSON, 1989)..............79

EQUATION 13-4: NEUTRAL POINT (NELSON, 1989) .................................................................................79

EQUATION 13-5: STATIC MARGIN ...........................................................................................................80


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5 Introduction

5.1 Aims

The aim of this project is to design a small Unmanned Aerial Vehicle which can be used

to assist in coast guard applications and the monitoring of Australian Coastlines.

5.2 Scope

This project is limited to the sizing and conceptual design of an unmanned surveillance

aircraft. Therefore, the report will not cover the complete detailed design of the aircraft.

The drawings associated with this project are also limited to conceptual sizing and design

and so, a complete set of engineering drawings of the aircraft will not be included.

5.3 Background Information

Unmanned Aerial Vehicles (UAV’s) have traditionally been used in military applications

for defence and security. However, in recent years there has been increased use of UAV’s

in the civil and domestic sectors due to new technology and a reduction in the costs

associated with this area of the aerospace industry. Tasks that were once performed by

large manned vehicles can now be performed by smaller unmanned aircraft, offering

economic benefits as well as a decreased risk of pilot casualties. This surge in popularity

has led to UAV’s being used for agricultural use, emergency response and natural

resource management. Recently, a number of contracts have been awarded to companies

to develop surveillance technology for UAV’s, which the Australian government hope

will lead to increased automation of coastline surveillance.


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5.4 Significance

Australia is a large country and is completely surrounded by water and as such, we have

some very long and vast coastlines, which also make up our country’s border. Although

many of our beaches and coasts are monitored by local authorities, much of our coastline

is very remote and activity on these coastlines can easily go unnoticed. Monitoring

coastlines is a difficult task and is made worse by the fact that Australia is so sparsely

populated. Coastwatch is the section of the Australian customs agency that is responsible

for monitoring Australian coastlines. Coastwatch covers more than 37,000 km of

coastline, plus an offshore maritime area of almost 15 million square kilometers

(www.defenseindustrydaily.com). Currently, coast watch is performed by a number of

fixed wing aircraft, helicopters and large transport aircraft such as the AP-3C Orion,

supplied by the RAAF.

Monitoring Australian coastlines can be made easier and more convenient through the

use of smaller, cheaper aircraft and in particular, unmanned aircraft or UAVs. By

designing a UAV complete with monitoring equipment, capable of flying along

Australia’s coastlines the level of security in Australia can be enhanced. The UAV can

perform in monitoring routines which would otherwise take a long time to complete. The

UAV will not only quickly and effectively patrol Australia’s coastlines, but it will also

save human resources which can be directed elsewhere and put to better use.

In particular the UAV will look for

• Illegal fishing vessels

• Asylum seekers

• Drug Smuggling operations

• Lost vessels

This aircraft will be designed for surveillance operations including the monitoring of

Australia’s coastlines and will meet the criteria described in the following sections.


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6 Technical Task

6.1 Standard Requirements

As the UAV is used over Australian waters, it must conform to Australian standards and

Australian Air standards as dictated by CASA Part 101. If an Australian standard is not

available then the relevant international standards will be used. The UAV will also adhere

to the FAR 23 and FAR VLA standards.

6.2 Performance Parameters

Range

Australian maritime zones are classified as follows (www.customs.gov.au):

• The territorial sea (TS) – 12 nm from the baseline

• The contiguous zone (CZ) – 24 nm from the baseline

• The exclusive economic zone (EEZ) – 200 nm from the baseline

The baseline is also known as the water level line at low tide and is the point from which

all maritime zones are measured. The territorial sea is subject to Australian jurisdiction,

while in the contiguous zone Australia is able to exercise its customs, fiscal, immigration

or sanitary laws and regulations. Within the exclusive economic zone, Australia has

sovereign rights over all natural resources of the water, sea surface and subsoils. All 3

zones need to be monitored, however the TS and CZ are more critical to the security and

wellbeing of Australia as it is these zones that drug and people smuggling operations

must traverse in order gain access to the country. Consequently, our design will focus on

monitoring the TS and CZ zones.

The Australian Customs Coastwatch has bases in Broome, Cairns, Darwin, Horn Island-

Torres Strait and Gove (www.customs.gov.au). The distance between each of these bases


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is approximately 1500km (travelling by coast); therefore the UAV will need to have long-

range capabilities. It may be necessary to build a number of smaller bases along

Australia’s northern coastline to deploy and refuel UAV’s.

Loiter time:

For surveillance UAV’s, loiter and endurance are often more important properties than

range and cruise speed. The aircraft must be able to loiter for as long as possible and

cover a significant area in order to be effective. If the loiter time is too short, then the

effectiveness of the UAV as a coast watch aid will be questionable. In accordance with

the specified mission profile (see figure 1 next page), the aircraft will cruise out to the

surveillance zone and commence loiter, travelling along the coastline before cruise in to a

refuelling location. A suitable loiter time of 5 hours is chosen for the UAV. This value

was chosen for a number of reasons: it was desirable to keep the total mission below 8

hours; due to the fact that the UAV is to operate primarily during daylight hours. The

UAV is to be operated and monitored by Coastwatch; therefore Coastwatch personnel

would not have to work at odd hours while monitoring the UAV.

Time of Climb

There is no minimum time to climb for this application, however the time of climb does

not need to be fast as the UAV flies at low altitude. Hence, a 5 min climb to cruise

altitude is acceptable.

Cruise out/Cruise in distance

A certain cruising distance needs to be estimated to allow the aircraft to fly to the coast in

order to commence surveillance. This doesn’t need to be a large distance as the bases for

the coast watch UAV should be situated close to the sea. Therefore a distance of 50km or

27 nautical miles is enough for the UAV to reach the coast. If this distance were any

larger, it would detract from the surveillance loiter time available to the aircraft.

Cruise Speed

It is desirable for the aircraft to fly at a high velocity in order to increase the efficiency

and cover large distances quickly, however there is no requirement for the UAV to be fast


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or stealthy to avoid being attacked. On the other hand, if the aircraft is travelling too fast,

it will not be possible to obtain a clear view through the optical equipment being used and

hence successfully monitor coastlines. Therefore, a compromise must be made.

Typically, surveillance UAV’s have a cruise speed of approximately 100kph, hence this

value will be used for the loiter phase of the mission profile. For the cruise out/cruise in

phases, a speed of 150kph will be used.

Altitude

The same logic applies here as for cruise speed. If the aircraft is too high, visibility is

weakened where as, if the aircraft is too low, the effectiveness of the engine is limited

and the field of view of the camera becomes narrower and thus the UAV becomes less

effective as a surveillance aircraft. Light aircraft pilots recommend that an altitude of

1000ft will provide adequate visibility.

Mission Profile

The mission profile is given in Figure 6-1.

Figure 6-1: Mission profile


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6.3 Technical Level

The aircraft is to be designed for use in a number of different locations throughout

Australia for coast watch purposes. Consequently, the operation of the aircraft should be

kept simple to ensure that it can be used by people without a pilots licence. A small

amount of training should be sufficient to operate this aircraft. The design of the aircraft

should also be kept as simple as possible to make maintenance easier, particularly if the

aircraft is to be used in some remote areas where technical support is not readily

available. As Australia is such a large country, there are many different climates

depending on location, hence the UAV should be able to operate in a range of weather

conditions from high temperatures and humid conditions to wet and cold conditions.

6.4 Economical Parameters

Although this aircraft is in a program which has some government funding, the cost of

this aircraft should be kept to a minimum. It is desirable to keep the cost of the UAV low

and have more of the units in operation throughout Australia rather than increase the cost

and limit the number operating. It is conceivable to see this aircraft being used by other

organisations and not just coastal patrol such as park rangers to monitor the activities in

national parks and lifeguards to monitor popular beaches. Therefore, the aircraft as a

whole must be affordable to some of these private organisations. Furthermore, the aircraft

must be cheap to maintain and run. If the aircraft has high running costs, this will limit its

use. Adding to the cost of this aircraft is the added electronic equipment it must carry

such as cameras and GPS systems, these items are expensive but necessary. The aircraft

itself should not be more than twenty thousand Australian dollars to buy with running

costs not exceeding $5000 per year for use of the aircraft once per day, including fuel and

any maintenance to be carried out. (Note: the running costs will vary according to the

amount of use).


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6.5 Power plant type and requirements

The aircraft will utilise a small petrol engine to power a single propeller. Ideally, diesel

would be a suitable fuel for use, particularly in remote areas. After excessive storage

times, some of the volatile components in petrol evaporate whereas, this is less common

with diesel fuel (Campbell, CJ, 1991). However, diesel engines add excessive weight and

complexity to the aircraft and so a petrol engine is a good alternative. Furthermore, petrol

engines are commonly used in many applications and are a reliable power source. They

can also be made in light weight configurations producing high power to weight ratios.

The power to weight ratio of the engine is very important when considering the overall

performance and weight of the UAV. The smaller common UAV engines on the market

produce between 20 and 50 hp but weight can vary. Any power within this range will be

acceptable providing a high power to weight ratio. For further discussion on power, refer

to the engine selection section.

6.6 Main system parameters requirements

In order for this UAV to perform its surveillance operations, it needs to be fitted with

appropriate optical equipment. The UAV should contain an interchangeable camera

system to enable flexible surveillance operations. A GPS system will assist with

programming flight paths and determining the location of any suspicious activity detected

during surveillance.

The aircraft should be lightweight and transportable. Should the aircraft need to be

transferred to another surveillance location, take off location or maintenance depot a large

cumbersome aircraft makes this difficult to do.

The UAV should also possess short take off and landing requirements. The likely

operation area will be the coastline, which poses restrictions on available area for take off


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and landing. A short take off and landing will ensure successful operation in a large

number of locations.

6.7 Reliability and Maintainability

This aircraft is going to experience flight times in excess of 6 hours at a time and as such,

it needs to be reliable. Since the aircraft is unmanned, the level of reliability can be

lessened somewhat. However, since the aircraft engine has to be certified to 150 hours of

endurance, the remainder of the aircraft should also meet this standard. Hence no

servicing should be required until after 150 hours of flight time has been completed. The

aircraft should then be serviced every 3 months after this.


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7 Statistical Analysis

Before the design stage of any aircraft can proceed, research into the design and

performance characteristics of similar aircraft is a useful strategy in order to produce a

summary of engineering and performance benchmarks. This is done to gain an

understanding of what reasonable performances are possible for aircraft of similar design

parameters. All data obtained for this section of the report was drawn from Jane’s

information group, 2002.

The basic equation used for a statistical analysis is given in Equation 7-1, where A and B

are empirical constants for a particular type of aircraft. We are concerned with UAV’s in

this project.

Equation 7-1: Statistical Analysis equation (Roskam, 1994)

This report will use Roskam’s method for estimating take-off weight of the aircraft.

Therefore, the A and B values to be used in this equation will need to be established. This

is done by producing a graph of log(We) versus log(WTO) and fitting a line of best fit.

This will be of the form log(We) = y + xlog(WTO). This is then rearranged to match

Equation 7-1 from which the values for A and B can be found. The ten aircraft listed in

Table 7-1 were selected based on our initial weight and size estimated requirements for

our proposed aircraft.


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log(We) vs. log(Wto)

y = 1.0321x - 0.2917

0

0.5

1

1.5

2

2.5

3

0 0.5 1 1.5 2 2.5 3

log(Wto)

log(W

e)

Aircraft We Wto log(We) log(Wto)

Sting 176.4 331 2.246499 2.519828

EMT Luna 44.1 66.1 1.644439 1.820201

Aerosonde 18.1 29.5 1.257679 1.469822

KAI 249 286.5 2.396199 2.457125

AAI Shadow 200.6 328 2.302331 2.515874

AAI Pioneer 276 419 2.440909 2.622214

Silver Arrow 48.5 79.4 1.685742 1.899821

BAE Phoenix 220 397 2.342423 2.598791

Silver Arrow mini 60 110 1.778151 2.041393

Aerosky 38.1 88.2 1.580925 1.945469

Table 7-1: Take off weights for similar aircraft

Figure 7-1: Empty weight versus take off weight


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The equation shown on the above graph is rearranged as discussed in the above section.

This yields the following values:

• A = 0.28263

• B = 0.9689

7.1 Wing Span/Aircraft Length

To obtain a reasonable figure for the wingspan of our proposed aircraft, it is necessary to

look at the wingspans of completed aircraft. A table of these values can be seen below in

Table 7-2.

Aircraft Wing Span

(m)

Aircraft Length

(m)

Sting 6 3.2

EMT Luna 4.17 2.24

Aerosonde 2.9 1.7

KAI 4.8 3.52

AAI Shadow 3.89 3.4

AAI Pioneer 5.11 4.26

Silver Arrow 3.57 2.56

BAE Phoenix 5.5 3.8

Silver Arrow

mini

3.66 2.74

Aerosky 4 n/a

Table 7-2: Wing span and aircraft length

Analysing Table 7-2, it can be seen that the smallest wing span value is 2.9m and the

largest is 6m. The maximum aircraft length is 4.26m and the minimum is 1.7m. This now


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gives the designers a reasonable idea of the basic dimensions of similar aircraft when

designing the proposed aircraft.

7.2 Take-off Methods

UAV’s can use different methods of take-off due to their light weight and ease of

transport. Larger UAV’s tend to use the common method of a run-way take-off as they

are too heavy and large for other methods. By studying the ten similar UAV’s (Table

7-3), it has been found that most similar to the proposed design, launch by method of

catapult or similar. From the data, it would seem that this would be the best method of

launch for the proposed UAV.

Aircraft Take-off Distance

Sting Bungee Launch

EMT Luna 4m Bungee Launch

Aerosonde Catapult/Car top at

40knots

KAI N/A

AAI Shadow Hydraulic Catapault

AAI Pioneer Catapault

Silver Arrow 82m

BAE Phoenix Hydraulic Catapault

Silver Arrow

mini

6m Catapault

Aerosky N/A

Table 7-3: Take off methods


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7.3 Landing Methods

Table 7-4 shows landing methods of similar aircraft to the proposed coast watch UAV.

Aircraft Landing

Sting Parachute or Landing with Arrestor Hook

EMT Luna Parachute

Aerosonde Belly landing, autonomously or under operator control

KAI Conventional wheeled landing standard; parachute for emergency

recovery

AAI Shadow Wheeled landing or parachute/parafoil retrieval

AAI Pioneer Wheel Landing, Tail Hook and Cables or Net

Silver Arrow Conventional wheeled landing

BAE Phoenix Parachute Airbag Method

Silver Arrow

mini

Parachute and Replacable Nose Cone

Aerosky Conventional wheeled landing

Table 7-4: Landing methods

Table 7-4 shows no obvious trend for landing methods. A landing method from this will

need to be chosen based on the application, climatic conditions and topography of the

area the proposed craft will be operating. Alternatively a different landing method could

be created.


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7.4 Summary of Benchmarks

From the analysis of the presented data, the following summary in Table 7-5 of

engineering benchmarks has been created.

Roskam's Equation values A = 0.3469, B = 0.9363

Wing Span > 2.9m, < 6m

Aircraft Length < 4.26m, >1.7m

Launch Distance/Method Catapault or similar

Landing method Many options

Table 7-5: Summary of statistical analysis


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8 Concept Sketches

Figure 8-1: Concept 1 sketch


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Figure 8-4: Concept 4, sketch 1


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9 Weight Estimation

This section of the report is concerned with estimating the take off and empty weight for

the UAV. To do this, data is required from the technical task and statistical analysis. The

following data is needed from the technical task:

• Cruise altitude = 1000 ft

• Cruise speed = 41.67 m/s

• Loiter speed = 27.78 m/s

• Range = 100 km

• Propeller efficiency = 0.8 (obtained from manufacturer)

From the statistical analysis, the required information is Equation 9-1.

( ) ( )emptyTO WW log9689.028263.0log +=

Equation 9-1: Take off weight equation using statistics

The next stage of the weight estimation is to form an equation of take off weight versus

empty weight using the various mission stages defined in the technical task. The general

equation used to do this is given in Equation 9-2.

emptyfuelpayloadcrewTO WWWWW +++=

Equation 9-2: General take off weight equation


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Equation 9-2 can then be solved along with Equation 9-1 to form an initial estimate for

the take off and empty weight. Firstly, all terms in Equation 9-2 need to be defined. Since

this project is concerned with designing a UAV, the crew and payload weight are zero.

The fuel weight can be calculated using the mass fuel fraction, as shown in Equation 9-3.

( )ffTOfuel MWW −= 106.1

Equation 9-3: Fuel weight equation

The mass fuel fraction is calculated using weight ratios for each mission stage of the

UAV, as shown in Equation 9-4.

∏=

+

=

n

i i

i

TO

ffW

W

W

WM

1

11

Equation 9-4: Mass fuel fraction

The remainder of this section will discuss calculation of the weight ratios. The first stage

in this process is to calculate the weight of fuel used for each stage. This is done by

defining the time, power requirements and fuel consumption for each stage. These values

are shown in Table 9-1. The values for cp and P have been taken from the Engine

selection section.

t (hr) cp (lbs/hp/hr) P (bhp)

Start-up 0.0833 0.57 10

Take off 0 - -

Climb 0.0185 0.57 22.5

Cruise out 0.33 0.52 19.1

Loiter 5 0.55 21

Cruise in 0.33 0.52 19.1

Descent 0.014 0.52 19.1

Landing/shutdown 0 - -

Table 9-1: Time, fuel consumption and power for each stage


-31-

Once this table has been formulated, the fuel weight for each stage can be calculated

using Equation 9-5.

)(lbsctPweightFuel p××=

Equation 9-5: Mission stage fuel weight

The next stage in the weight estimation is to form an initial guess for the UAV take off

weight. Then, using the fuel weight for each stage and the initial guess, the weight ratio

for each stage can be calculated. The weight ratios are then used to fully define Equation

9-2, which is then solved with Equation 9-1 to find the take off and empty weight. This

process is iterated until the initial guess matches the final value. Table 9-2 shows data for

the fuel weight for each stage, as well as the weight ratios.

Initial guess = 174 lbs

Fuel weight (lbs) Weight after stage

Weight ratio

Start-up 0.475 173.53 0.997

Take off 0 173.53 1

Climb 0.2375 173.29 0.9986

Cruise out 3.31 169.98 0.981

Loiter 57.75 112.23 0.6602

Cruise in 3.31 108.92 0.9705

Descent 0.1379 108.78 0.9987

Landing/shutdown 0 108.78 1

Table 9-2: Fuel weight and weight ratio for each stage

Using these number, the mass fuel fraction is calculated using Equation 9-4 and is

0.6336. Equation 9-2 can now be written as shown in Equation 9-6.

emptyTO WW ×= 6351.1

Equation 9-6: Final take off weight equation


-32-

Weight Estimation

0

50

100

150

200

250

300

350

400

0 50 100 150 200 250

We (lbs)

Wto (lbs)

Calculation

Statistics

The graph of Equation 9-1 and Equation 9-6 is shown in Figure 9-1. Using this figure,

and defining the solution as the point of intersection, the take off and empty weights are

172.3 and 103.9 lbs respectively.

Figure 9-1: Weight estimation graph


-33-

10 Sensitivity Analysis

Following the weight estimation, a sensitivity analysis is required to find the parameters

to which the take off weight is highly dependant. Sensitivity was calculated for both the

endurance and range case for the following parameters:

• Specific Fuel Consumption

• Propeller Efficiency

• Cruise Velocity

• L/D Ratio

The sensitivity to all these parameters to take-off weight can be found using Equation

10-1 where F is defined in Equation 10-2. The values of y

R

∂∂

and y

E

∂∂

can be found using

Table 10-1 for the propeller driven aircraft case.

y

RF

y

Wto

∂∂

=∂∂

y

EF

y

Wto

∂∂

=∂∂

Equation 10-1: Take-off Weight Sensitivities for Range and Endurance Cases

ffreserve MMDBCWtoWtoBF )1())1(()( 12 +−−−= −

Equation 10-2: Equation to calculate F


-34-

Table 10-1: Values for Sensitivity Equations (From Airplane Design by J. Roskam)

To calculate F, values are required from the weight estimation section. C and D are calculated using Equation 10-3 and Equation 10-4. The values used to calculate C and D are shown in Table 10-2. These values are also used in the calculation of the sensitivities as can be

seen in Table 10-1. The values obtained for A, B, C and D were 0.28263, 0.9689, 0.6286

and 0 respectively.

Funuseableffreserve MMMC −−+−= )1)(1(1

Equation 10-3: Equation for ‘C’ Calculation

crewPL WWD +=

Equation 10-4: Equation for ‘D’ Calculation


-35-

WTO 272.23

We 166.5

Mreserve 0

mff 0.633558

mfunusable 0.005

WPL 0

Wcrew 0

Table 10-2: Values used for calculating ‘C’ and ‘D’

The following values displayed in Table 10-3 are also required for the sensitivity

analysis.

Parameter Cruise Loiter

cp 0.52 0.55

np 0.8 0.8

L/D 10 8

V (mph) 93.33333 62.22222

R (sm) 62.5

E 5

Table 10-3: Required values for sensitivity analysis

Using Equation 10-2, F is calculated as 5412.11. This value can then be used in the

sensitivity equations shown in Table 10-1 to calculate y

R

∂∂

and y

E

∂∂

for each of the

parameters mentioned. Table 10-4 shows the values obtained for the sensitivity for each

parameter. Note that the velocity has no effect for the range case.


-36-

Sensitivities

Range Endurance

0.94 R / E 77.17

225.50 cp 701.57

-146.58 np -482.33

V 6.20

-11.73 L/D -48.23

Table 10-4: Sensitivities to the Main Parameters

Range case

The sensitivities in Table 10-4 can be interpreted as follows. For every mile added to the

range of the aircraft, the take off weight will increase by 0.94 pounds. If the specific fuel

consumption increases by 0.2, the take off weight would increase by 0.2 x 225.5 lbs. If

the propeller efficiency and L/D are increased, the take off weight will decrease, as

shown by the negative sensitivities.

Endurance case

The sensitivities for the endurance case are much higher than the range case. This is due

to the aircraft being designed primarily for loiter as opposed to cruise. This was discussed

in the Technical Task. For every hour added to the loiter time, the aircraft take off weight

will increase by 77.17 pounds. The sensitivity to specific fuel consumption has the most

potential for decreasing the UAV take off weight. As will be discussed in the Engine

Selection section, it was possible to select an engine with a specific fuel consumption of

0.33 lbs/lbs/hr. This would result in significant weight benefits for the aircraft. However,

this engine could not be chosen due to geometry restrictions which posed a greater

disadvantage than the advantages associated with the lower fuel consumption.


-37-

11 Aircraft Sizing

In the absence of well-defined UAV standards for climb, cruise and takeoff and landing

distances, the FAR 23 standards for small aircraft were applied. Outlined below are the

calculations that were used, the values used within the calculations and the final sizing

graph. Some of the values used are closely coupled with the weight estimation section.

All of the results were iterated until all results sufficiently matched.

11.1 Stall Speed Sizing

Stall speed sizing was undertaken using a standard stall speed of 61kts. The following

equation yields a value for W/S which is constant for all values of W/P.

W

S=1

2ρVstall

2 CLmax

Equation 11-1: Stall speed equation

Wing Loading for Stall Speed

W/S (Vstall) 14.79

11.2 Take off distance sizing

Sizing to FAR 23 Takeoff Req

CLmaxTO 0.85

Sto 300 ft

Stog 498 ft

TOP23 87.55


-38-

The equation used for take-off sizing is given in Equation 11-2. In this case, takeoff is

occurring at h=0, therefore σ = 1.

W

S

TO

= TOP23 ⋅σ ⋅CLmaxTO ⋅hp

W

TO

Equation 11-2 Take-off sizing equation

Equation 11-2 is solved and the results are displayed below in Table 11-1.

Table 11-1: Take-off distance solution

11.3 Landing distance sizing

The values for landing distance sizing are given in Table 11-2.

Sizing to FAR 23 Landing Req

Slg 500 ft

Sl 969 ft

Takeoff Sizing Table

W/S (lb/ft2) (W/P)TO

(lb/hp)

5 14.88

10 7.44

15 4.96

20 3.72

25 2.97

30 2.48

35 2.12

40 1.86

45 1.65

50 1.49


-39-

VsL 43.42 kts

WTO 172 lbs

WL 105 lbs

WTO/ WL 1.64

ClmaxL 1.05

(W/S)L 2.30 lbs/ft2

(W/S)TO 3.78 lbs/ft2

Table 11-2: Landing distance sizing

11.4 Climb Sizing

As our UAV is to have only a single engine, the FAR 23.67 climb requirements for One

Engine Inoperative (OEI) will be neglected. In terms of the value for coefficient of lift,

maximum values associated with common UAVs were chosen and are outlined in Table

11-3.

CLTO MAX 0.85

CLTO 0.45

CLland MAX 1.05

CLland 0.56

CL cruise 0.65

Table 11-3: Lift coefficient for all configurations

In order to calculate a first estimate of the zero lift drag coefficient (CD0) it is required

that we obtain an equivalent equivalent skin friction drag (Cfe) and a value for Swet/Sref, as

shown in Equation 11-3. Taking Cfe to be similar to that of a light aircraft with a single

engine and assuming a value of Swet/Sref, the values in Table 11-4.

CD0 = C fe

Swet

Sref

Equation 11-3: Zero lift drag coefficient estimation


-40-

Cfe 0.01

Swet/Sref 5

CD0 0.05

Table 11-4: Drag polar values

The zero lift drag coefficient can then be used with Equation 11-4 to help determine the

Coefficient of Drag (CD).

CD = CD0 +1

π ⋅ A ⋅ e⋅CL

2

Equation 11-4: Drag polar equation

Here, ‘A’ denotes the Aspect Ratio and ‘e’ denotes the Oswald efficiency factor which

are both assumed and shown in Table 11-5.

A 7

e (Takeoff) 0.8

e (Landing) 0.75

Table 11-5: Aspect ratio and Oswald efficiency factor

Thus for each of the configurations required for the FAR23 climb sizing, the expression

for CD are detailed below in Table 11-6. It is assumed that takeoff flaps add 0.015 to the

drag polar and that landing flaps add 0.065 to the drag polar. Note that the effects of

landing gear drag are not being considered at this stage as the UAV is planned to be

launched via catapult and caught in a net.

Flaps Landing Gear Coefficient of Drag (CD)

Takeoff n/a 2

8.07

1015.005.0 LtakeoffD CC ⋅

⋅⋅++=π

Landing n/a 2

75.07

1065.005.0 LlandingD CC ⋅

⋅⋅++=π

Table 11-6: Drag polar equations for FAR23 climb sizing


-41-

11.4.1 FAR 23.65 Rate of Climb Sizing

Rate of climb specified in FAR23.65 as being greater than 300fpm. Following from that

the value of RCP = (33000)-1 x RC. The final relationship between W/P and W/S is

determined using Equation 11-5.

RCP =ηp

(W /P)

−

(W /S)1/ 2

19((CL )3 / 2 /CD )maxσ

1/ 2

Equation 11-5: FAR23 RoC sizing equation

Here, the propeller efficiency is taken to be ηp = 0.8 , the density ratio σ = 1 (as this

standard is calculated for Sea level conditions). In the below table an additional

calculation is required to convert the W/P values to (W/P)TO values. Dividing the W/P

values by 1.1 will take into account the thrust required for takeoff and return a value for

(W/P)TO .

11.4.2 FAR 23.65 Climb Gradient Sizing

Using an estimation for CLclimb and the value for CDclimb calculated earlier,(L/D)climb can

be calculated. As defined by the FAR23.65 standard, CGR =1/12rad and CGRP can be

found using Equation 11-6.

CGRP =CGR + (L /D)−1( )

CL

1/ 2

Equation 11-6: FAR23.65 CG sizing equation

The relationship between W/S and W/P is then found using Equation 11-7. Once again,

the propeller efficiency is taken to be ηp = 0.8 , the density ratio σ = 1. Also, we must

again correct the W/P value to the takeoff value by dividing by 1.1.


-42-

CGRP =18.97 ⋅ηp ⋅σ

1/ 2

(W /P)(W /S)1/ 2

Equation 11-7: Final equation for FAR23.67 sizing

11.4.3 FAR 23.77 Climb Gradient Sizing

Sizing to the FAR23.77 standard is very similar to the FAR 23.65 climb gradient sizing

outlined above. However, in this case we have CL and CD both for landing configuration

and CGR = 1/30rad. Again, we must adjust the W/P value for takeoff thrust by dividing it

by 1.1. The final climb sizing values are presented in

Table 11-7.

FAR 23 Climb Sizing

FAR 23.65 RC FAR 23.65 CGR FAR 23.77 CGR

W/S W/P W/P (TO) W/P W/P (TO) W/P W/P (TO)

5 37.69 34.26 23.39 21.26 31.86 28.97

10 30.47 27.70 16.54 15.04 22.53 20.48

15 26.57 24.15 13.50 12.28 18.40 16.72

20 23.98 21.80 11.69 10.63 15.93 14.48

25 22.08 20.07 10.46 9.51 14.25 12.95

30 20.61 18.73 9.55 8.68 13.01 11.83

35 19.42 17.65 8.84 8.04 12.04 10.95

40 18.43 16.75 8.27 7.52 11.27 10.24

45 17.58 15.98 7.80 7.09 10.62 9.66

50 16.85 15.32 7.40 6.72 10.08 9.16

Table 11-7: Final climb sizing values


-43-

11.5 Cruise Sizing

Sizing for cruise can be achieved by taking the power index, Ip from a chart and applying

the following calculation. Taking Ip = 1.1 gives a relationship for W/S to W/P. As it can

be assumed that 75% of power is used during cruise, the value for W/P can be adjusted to

take-off power by multiplying 0.75. The values are shown in Table 11-8.

Cruise Sizing

W/S W/P W/P(TO)

5 3.830851195 2.873138396

10 7.661702389 5.746276792

15 11.49255358 8.619415188

20 15.32340478 11.49255358

25 19.15425597 14.36569198

30 22.98510717 17.23883038

35 26.81595836 20.11196877

40 30.64680956 22.98510717

45 34.47766075 25.85824556

50 38.30851195 28.73138396

Table 11-8: Cruise sizing values

11.6 Overall Sizing Chart

The overall sizing chart is shown in Figure 11-1. From Figure 11-1 it can be seen that the

takeoff distance have a profound effect upon the overall sizing of the aircraft. In order to

make the sizing chart more reasonable, it is required to dramatically increase both the

landing and takeoff distances up to 1000ft. As the aircraft in question is to be a UAV able

to be launched from a variety of locations, this increase in the required landing and


-44-

takeoff distances is undesirable. An alternative method of launch and capture is thus

considered. Thus, by launching the UAV by means of a catapult and by landing the UAV

into a large net, the takeoff and landing distances can be neglected from the sizing

discussion. The revised sizing chart is shown in Figure 11-2.

Figure 11-1: Sizing Chart


-45-

Figure 11-2 : Sizing chart without Landing and Takeoff

It can be observed that a matching point exists where:

• W/S = 15

• W/PTO = 8.5

From these values and using W =175lbs, the values required for PTO and S can be

determined. These values are outlined in Table 11-9.

W 175 lbs

S 11.7 ft2

PTO 20.6 hp

Table 11-9: Final sizing values


-46-

12 Overall Configuration Design

12.1 Fuselage Design

Various factors need to be considered for the fuselage design. For a UAV, this may not

hold as much importance as for a larger aircraft, but still warrants discussion. Various

pertinent factors relating to fuselage design are discussed below.

For a UAV, the main design parameters for the fuselage are aerodynamic performance

and component storage. Aerodynamic performance can be broken down into friction and

pressure drag. Compressibility drag can be ignored at such low speeds, and induced drag

needs only be considered for a flying wing aircraft.

Friction drag is mostly dependant on the fineness ratio of the aircrafts fuselage. The

fineness ratio of an aircraft is given by the length divided by the average diameter as

shown in Equation 12-1. The terms are shown in Figure 12-1. The value of Lf for the

designed aircraft is 1500mm and the value for Df is 600mm (found by an average of the

side on and top view diameter). This gives a fineness ratio for the coast watch UAV of

2.5. A graph of typical values of fineness ratio for subsonic aircraft is shown below in

Figure 12-2. The graph shows that once the fineness ratio gets below 2, the drag ratio

increases dramatically. The designed coast watch UAV is on the favourable side of this

limit.

Equation 12-1: Fineness ratio


-47-

Figure 12-1 Fineness Ratio terms

Figure 12-2: Fineness ratio for Subsonic Aircraft

Lf

Df


-48-

The Profile and base drag is mainly dependant on the profile of the fuselage of the

aircraft as the name suggests. This type of drag is increased with increased separation of

the flow from the fuselage. This means that blunt profiles have a larger profile drag than

streamline profiles. The designed aircraft has a large upsweep which can lead to large

separation and increased profile drag. The pusher propeller design assists to re-energise

the boundary layer of the aft-end of the fuselage and keep the separation of flow to a

minimum.

12.2 Aerofoil Selection

One of the more important features of an aircraft is the aerofoil. An aerofoil forms the

cross-section of an aircraft wing and consequently, is the primary source of lift. It is

possible to design a custom aerofoil to suit our design, however, this would be an

expensive and unnecessary process as there are many existing aerofoil designs that will

suit this application. Various aerofoil sections will be investigated, and the most

appropriate will be chosen for our design. It is not feasible to examine all aerofoil cross

sections, therefore, in conjunction with the statistical analysis, only aerofoils used in

similar aircraft will be investigated.

Since we are designing a Coast watch UAV, there are particular features which we are

looking for. In particular, we seek:

• short take off, hence high lift at low speed

• low drag in cruise configuration to keep fuel requirements to a minimum

• rigid wings to reduce flutter (making control programming easier)

• a main wing that stalls before the tail

To meet these requirements, we will investigate other aerofoils which are currently in use

in small aircraft.


-49-

Wing Profile Selection

Table 12-1 shows some small aircraft and the aerofoil which they use for the main wing.

Aircraft Aerofoil

AAI AA-2 Mamba NACA 4412

AAI Shadow 200 NACA 4415

AAI Shadow 400 NACA 4415

BAI Aerosystems Dragon Drone NACA 63A012

Table 12-1: Common airfoil sections for UAV aircraft (Lednicer, 2007)

As can be seen in Table 12-1, all of the aircraft use NACA (National Advisory

Committee for Aeronautics) profiles. NACA are not the only aerofoil designs available

but they are widely used, particularly in small aircraft. Furthermore, NACA 4000 series

and 2000 series aerofoils are used in many of the smaller, manned aircraft produced for

private use. The Air Tractor AT-802 uses a NACA 4415 aerofoil, the Cessna 205 uses a

NACA 2412 and the Jabiru LSA uses a NACA 4412 (Lednicer, 2007). Both of the 2000

and 4000 series are cambered aerofoils. For a complete analysis, symmetrical aerofoils

will also be considered in this section.

The aerofoils mentioned in Table 12-1, along with two symmetrical aerofoils (0012 and

0016) will be further investigated to consider their lift and drag qualities. The lift-curves

of the aerofoils are shown in

Figure 12-3.


-50-

Figure 12-3: Aerofoil Lift Curves (Modelfoil)

The cambered designs provide a higher coefficient of lift at lower angles of attack and

can reach higher lift coefficient prior to stalling, this is a desirable characteristic for the

wing. This characteristic will enable simpler assembly of the aircraft with respect to the

aerofoil incidence angle. The required lift coefficient at take off is 1.2 (see Aircraft Sizing

section). This is very close to stall for the two symmetrical aerofoils, as can be seen in

Figure 12-3. This could be avoided through the use of flaps, which will increase lift and

delay stall. However, the addition of flaps will add cost and weight to the aircraft. For a

simple UAV design, it will be much easier to use the cambered design which provides

higher lift. If required, flaps can still be added to the cambered design, however these will

not be as complex as those required for the symmetrical design. Because of this, a

cambered aerofoil will be selected.

Of the two thickness ratios considered (12% and 15%) the data obtained from Modelfoil

showed only minor differences in the lifting characteristics, however it is known that

thick airfoils (t/c>14%) have different stall characteristics to thin aerofoils. Separation


-51-

begins at the trailing edge of a thick aerofoil and gradually moves toward the leading

edge, giving warning that stall is about to occur, whereas thin aerofoils stall fairly

suddenly without warning (Raymer, 1992). In addition to this, the increased moment of

inertia of the thicker aerofoil section results in smaller bending stresses and wing

deflection, meaning that wing structural weight is less.

Figure 12-4 – NACA 4415 (Modelfoil)

Stall characteristics are dependent on wing properties as well as aerofoil profile. Twist,

dihedral, taper, sweep and aspect ratio are 3D effects that contribute to the stall

characteristics of the aerofoil, these are discussed in Wing Design and Positioning. It is

desired for the tail to stall later than the main wing. This will cause positive stability

characteristics, as if the main wing stalls first, the aircraft will return to its original

position. However, if the tail were to stall first, the aircraft will become unstable as the

tail is no longer providing balancing forces and moments. In general, the simplest way to

ensure that the wing stalls before the tail is to use a larger aspect ratio for the wing. The

aspect ratio of the wing was chosen as 7 based on statistical methods, while the horizontal

tail had an aspect ratio of 6, as calculated in Tail Sizing.


-52-

Tail Profile Selection

For ease of manufacture and simplicity, the vertical and horizontal tail airfoils can have

the same profile. The selected airfoils should be uncambered, as the vertical tail should

generate no lift under cruise conditions. The aerofoils considered were the symmetrical

NACA aerofoils at thicknesses of 12%c and 16%c. Both are commonly used as

horizontal and vertical stabilisers on small aircraft. A thinner section results in less drag

and a higher stall angle which are desirable characteristics for the tail. Therefore the

aerofoil profile selected for the horizontal and vertical tails is the NACA 0012

symmetrical aerofoil. The NACA 0012 is shown in Figure 12-5.

Figure 12-5 – NACA 0012 Profile (Modelfoil)

12.3 Wing Design and Positioning

The design and positioning of a wing can have an effect on numerous aircraft aspects

including safety, visibility, drag, weight, speed and stall of the aircraft. The main

considerations in regards to our UAV are outlined below.


-53-

Vertical Position

As our aircraft will primarily be used for surveillance, surveillance equipment needs to be

mounted on the underside of the aircraft. A high-wing configuration will allow for a

greater field of view for this surveillance equipment. While a low wing aircraft can be

safer in the event of a crash, our UAV will not be carrying a cargo in the main body and

thus this crash safety is not a requirement. The decreased drag associated with a mid-

wing aircraft will also be negligible as our aircraft is small and will be flying at low

speeds. A high-wing position will also be lighter than a mid-wing position, allowing for a

greater range.

Based on the parameters discussed above, a high-wing position was chosen for our UAV.

Wing Sweep

The main benefits of wing sweep are observed when the aircraft in question is travelling

at speeds nearing or in excess of the speed of sound. In terms of weight and cost, both

forward and aft sweeping of the wing increases the weight and cost of the aircraft. As our

UAV will be travelling at speeds significantly lower than the speed of sound, the

compressibility drag generated by a non-swept wing will be negligible, therefore, the

wing of our UAV will be non-swept.

Wing Aspect Ratio

While a high aspect ratio wing gives a lower induced drag and high lift-curve slopes, low

aspect ratio wings exhibit better aero-elastic stability and lateral stability. A high aspect

ratio wing also requires significantly more weight to reinforce than a low aspect ratio

wing.

As our UAV will be launched from a catapult and caught by a net, the stresses exerted

upon the wings are likely to be considerable. Therefore, a high aspect ratio wing may be

more likely to fail during one of these activities.


-54-

Thus a wing aspect ratio of 7 was chosen. This aspect ratio is a moderate one which can

exhibit some of the benefits of a high-aspect ratio wing while retaining the strength of a

low aspect ratio wing.

Wing Thickness Ratio

In terms of wing thickness, thick wings contribute more to profile drag at subsonic speeds

but are also lighter due to a higher stiffness. A thick wing also allows for greater fuel

carrying capacity.

A medium value of 15 was chosen for this project in order to decrease weight, increase

strength and keep profile drag to a manageable level.

Wing Taper Ratio and Wing Twist

Wing taper ratio and wing twist are both methods to modify the lift distribution on the

wing.

The complexity of both tapering and twisting means that they are both costly operations.

As our UAV is aiming to be a cost effective solution, any increase in cost is undesirable,

thus, the wings will not exhibit any taper or twist.

Dihedral

The inherent stability of a high-wing aircraft renders the increased stability gained by a

slight dihedral irrelevant. Also, the mounting of tail booms would be made more complex

if the wing were at an angle. Following from this, our UAV will not exhibit a dihedral.

Incidence Angle

Using the lift coefficient required by the main wing in Figure 12-3, it can be seen that an

angle of attack of 2° is required to obtain this. The aircraft will be easier to manufacture if

the wings are mounted with an angle of incidence of 2°. This will save time, money and

weight associated with the design and manufacturing of flaps.


-55-

12.4 Tail Design

The coast watch UAV incorporates a twin boom tail design. This allows for vertical

stabilisers to be mounted on each boom with a horizontal tail joining the two booms. An

example of the proposed configuration is shown in Figure 12-6.

Figure 12-6: Twin Boom Tail Configuration (http://aeroweb.lucia.it/rap/Paris97)

The horizontal tail will be mounted high on the vertical stabilisers, which means the

stabiliser will experience clean airflow rather than being in the wake of the propeller and

main wing.

Tail Booms

The tail booms connect the horizontal and vertical tail assembly to the fuselage via the

wing. The tail booms are attached to the lower surface of the wing rather than directly to

the fuselage, this was necessary in order to allow sufficient space for the propeller, which

had a diameter of 33inches (838mm). This spacing between the booms also allowed the

horizontal stabiliser to have a larger span without overhanging the sides of the vertical

stabilisers. The tail booms use a curved design for aesthetics and total length of 1.7m

each. Due to the large loads transferred to the wings, the joint surface is fairly large in

order to spread the load evenly. Having the booms attached directly to the wings may


-56-

also require additional structural work when joining the wings to the fuselage due to the

extra loading on the wing.

Profile Selection

For ease of manufacture and simplicity, the vertical and horizontal tail airfoils will have

the same profile. The selected airfoils should be uncambered, as the vertical tail should

generate no lift under cruise conditions. The airfoil profile selected for the horizontal and

vertical tails is the NACA 0012 symmetrical airfoil, which has a maximum thickness of

12%c.

Sizing

Calculations of the required sizes of the horizontal and vertical tail were based on the

methods presented by Raymer (2006). The vertical and horizontal tail volume

coefficients, shown in Equation 12-2 and Equation 12-3 respectively, are a statistics

based method for finding a suitable areas, using parameters of the aircraft, namely wing

chord (cw), wing span (bw) and length of moment arm (L).

vt

wwVV

L

SbVS =

Equation 12-2: Vertical tail volume coefficient

ht

wwHH

L

SCVS =

Equation 12-3: Horizontal tail volume coefficient

Suitable values for the vertical and horizontal tail volume coefficients were found by

analyzing the specifications of similar UAV’s and using the data presented by Raymer

(2006). Typical values for both vertical and horizontal coefficients obtained from

statistics are given in Table 12-2.


-57-

Aircraft HV VV Reference

BAI ’Javelin’ 0.6364 0.0372 Janes-Information-Group

(2002)

INTA ’Alo’ 0.5935 0.0337 Janes-Information-Group

(2002)

Aerosonde ’Aerosonde’ 0.93 0.0201 Janes-Information-Group

(2002)

Homebuilt Aircraft 0.50 0.04 Raymer (2006)

General Aviation Aircraft 0.70 0.04 Raymer (2006)

Average 0.67 0.035

Table 12-2: Aircraft Volume Coefficient Data (Avalakki et al, 2007)

The values used for the coast watch UAV were selected using Table 12-2 and are:

HV = 0.6

VV = 0.035

Using the volume coefficient data and information on the main wing and fuselage of the

UAV, the stabilisers could be sized. The properties of the main wing are shown in Table

12-3.

Aspect Ratio 7

Span 2.8 m

Chord 0.4 m

Sweep 0

Area 1.12 m2

Table 12-3: Main Wing Properties


-58-

Raymer suggests that for aft-mounted engines such as one mounted for “pusher

propeller” operation; the tail arm is approximately 50% of the fuselage length. Hence the

moment arm used in calculations will be 1.5m.

Using Equation 12-2, vertical tail area becomes:

5.1

12.18.204.0 ××=vtS

2084.0 mSvt =

Equation 12-4: Vertical tail area

Using Equation 12-3, horizontal tail area becomes:

5.1

12.14.06.0 ××=htS

218.0 mSht =

Equation 12-5: Horizontal tail area

The areas calculated by Equation 12-4 and Equation 12-5 are used as approximate values

only; ideally they should be considered as minimum values for conceptual design and

then refined during detailed design once the dynamics of the aircraft have been analysed.

There are two vertical stabilisers; hence the total area calculated in Equation 12-4 is

divided by two to obtain the area for each vertical stabiliser.

It is desired to avoid the wakes created by the main wing and propeller as much as

possible, thus it is proposed to mount the horizontal stabiliser between the tips of the two

vertical stabilisers, thus placing constraints on the design. The vertical stabilisers must be

of sufficient height to avoid the trailing vortices and the span of the horizontal tail must

be approximately equal to the spacing between the two tail booms.


-59-

It was decided that to meet the area requirements, the vertical stabilisers would have span (height) of b=0.6m, taper ratio of zero and sweep of zero. This gives the rectangular geometry shown in Figure 12-7 and a total area of 0.132m2.

Figure 12-7: Vertical Stabiliser Dimensions

The spacing between the tail booms restricts the span of the horizontal stabiliser. Tail

boom spacing is dependent on propeller diameter as discussed earlier is equal to 0.9m.

Therefore, the horizontal stabiliser will have a span of 0.9m and a chord length

approximately equal to the tip chord length of the vertical stabiliser, this gives the

geometry in Figure 12-8 .Total horizontal stabiliser area is 0.162m2, slightly less than that

calculated by the volume coefficient method.

0.11m

0.6m 0.48m

Rudder


-60-

Figure 12-8: Horizontal Stabiliser Dimensions

12.5 Control Surface Sizing

The control surfaces on the coast watch UAV are the ailerons (roll), the rudders (yaw)

and the elevator (pitch). For preliminary conceptual design, the analysis presented in the

following section is sufficient, however, during detailed design it should be checked that

the control surfaces provide adequate control authority and are structurally sound.

Ailerons

Raymer (2006) suggests that ailerons chord length should be between 15% and 25% of

the main wing chord length. They should extend from the 50% span to 90% to be in the

aerodynamically optimum position. This allows the ailerons to operate in relatively

undisturbed airflow, whilst avoiding the wing tip vortices.

Caileron = 0.2Cht

⇒ Caileron = 0.06m

Elevator

The Coastwatch UAV utilises a single elevator located on the horizontal tail. General

guidelines for elevator design are that the chord length should be between 20% and 30%

of the chord length of the horizontal stabiliser (Simons, 2002).

0.9m

0.18m

Elevator

0.63m


-61-

Celevator = 0.3Cht

mCelevator 054.0=⇒

The horizontal stabiliser extends from 15% of the span to 85% in order to give a surface

that is as large as possible.

Rudder

The twin-boom design of this aircraft lends itself to the use of two rudders that operate simultaneously. General guidelines for rudder design are the same as for elevators, the chord length should be between 20% and 30% of the chord length of the vertical stabiliser (Simons, 2002) and it should have the same taper ratio. The selected rudder geometry is shown in Figure 12-7. The span of the rudder extends from 10% of the span to 90% in order to give

a surface that is as large as possible.

vtrudder CC 3.0=

mCrudder 035.0=⇒

12.6 Propulsion System

The first basic decision to be made is propeller or jet powered engine. The major

consideration here is the cruising speed. The cruising speed for the UAV is 150 km/h

(and 100 km/h for the loiter phase). This is a relatively low speed, and a propeller engine

will definitely be sufficient. A propeller engine has further benefits over a jet engine with

regard to cost and weight. The following sections will discuss the required propeller

sizing and engine selection. As shown in the analysis below, one engine will be sufficient

to meet the requirements.


-62-

Propeller Sizing and Selection

The Coast Watch UAV will be propeller driven and hence selection of a suitable

propeller is important. The optimal performance of this aircraft is preferable in cruise and

as the aircraft does not have vigorous climb requirements, a “cruise prop” will be used. In

theory, the larger the diameter of the propeller, the more efficient it becomes (Raymer,

2006), however, the maximum diameter is limited by the propellers tip speed.

The size of the propeller is limited by the vector sum of the propeller’s tip speed and the

forward speed of the aircraft, as shown in Equation 12-6.

22

, cruisetiphelixicaltip VVV +=

Equation 12-6: Propeller tip velocity equation (Raymer, 2006)

The velocity limitation depends on the material of the propeller, as shown below.

• Wooden propellers, fpsV helixicaltip 850, <

• Metal propellers, fpsV helixicaltip 950, <

But in order to reduce propeller noise, it is recommended that fpsV helixicaltip 700, <

(Raymer, 2006). Since the coast watch UAV will be flying at altitudes lower than

conventional aircraft, the propeller will be sized such that it does not produce excessive

noise i.e. fpsV helixicaltip 700, < . This restriction will also allow the UAV to utilise wooden

propellers which are significantly cheaper than composite material and metal propellers

as discussed below.

The cruise speed of the aircraft is 100km/hr but the cruise out and cruise in speeds are

higher at 150km/hr. Equation 12-6 is rearranged to from Equation 12-7, which is used to

calculate the propeller diameter.

22

, cruisehelixicaltiptip VVV −=

Equation 12-7: Rearranged propeller tip velocity equation (Raymer, 2006)


-63-

The following are calculations used to size the propeller:

rpmin speed engine is

Where

60

diameter maximum Now,

/23.1901.1

25.209

gusts for wind allow toused be will10% offactor safety A

/25.209

/67.41/150

/36.213700

max

,

n

n

VD

smV

smV

smhrkmV

smfpsV

tip

tip

tip

cruise

helixicaltip

π=

==∴

=⇒

==

==

Maximum engine power is produced at 6700rpm (Zanzottera Technologies ltd, 2007),

hence the maximum diameter can now be calculated.

inchmD 34.215422.0max ≈=∴

However, as can be seen in the technical drawings, a propeller of this size will be

ineffective due to the larger cross-sectional area of the fuselage. By examining the

geometry of the tail boom, the distance between the propeller centre and the tail boom

can be calculated

There is a maximum clearance of 460mm between the propeller centre and the tail boom.

Therefore, by reducing the operating speed of the engine to 4200rpm, the diameter of the

propeller can be increased. It is acceptable to reduce the engine operating speed to this

value as the power required by the aircraft is approximately half of the available power

generated by the engine. Hence the diameter becomes,

100

450

460≈


-64-

m

D

952.0

4200

6025.205max

=×

×=⇒

π

Incorporating the 10% safety factor, gives a diameter of:

inchmmD 5.3385.0 ≈=

As shown above, the maximum distance limiting the propeller diameter is

460×2=920mm. Therefore, by using a propeller with a 33 inch diameter, the propeller

size is maximised whilst leaving sufficient clearance from the tail boom.

In addition to the propeller diameter, the pitch of the propeller can be sized using the

figure from reference (Simons, 2002). For a cruise speed of 150km/hr, a pitch size of 12

inches was found to be appropriate.

The Desert Aircraft Company (Desert Aircraft, 2007) is a supplier of parts and

accessories for UAVs and small model aircraft. Desert Aircraft have catalogued a number

of different propellers from various manufacturers. It is noted that by inspection of the

Desert Aircraft catalogue, wooden propellers are much cheaper than composite or metal

propellers. Therefore, using a wooden propeller will assist in lowering the cost of the

aircraft. Based on this and the above calculations, an MSC 33 x 12 inch wooden propeller

was selected. MSC manufacture wooden propellers which offer high performance but

produce lower levels of noise which is desirable. Lastly, the propeller will have two

blades as this further decreases cost and will provide adequate thrust to the aircraft. The

engine selected for the aircraft will provide more than the required power for the aircraft

to fly and hence justifies the use of a 2 bladed propeller as opposed to 3 blades.

The manufacturers suggested that a propeller efficiency value of 0.8 would be sufficient.


-65-

Engine Selection

The design of the coast watch UAV is to be as compact and small as practically possible.

As mentioned in the Technical Task, the power-weight ratio is an important factor in

engine selection. In order to select a suitable engine, the smallest available engines for

use in UAV’s were researched. From this search, the possible engines for this UAV have

been narrowed down to four options, which are shown in Table 12-4. The following is a

discussion of the pertinent points relating to the engine selection.

The AR741 is a Wankel rotary engine specifically designed for surveillance use whereas

the remaining three engines are piston engines. All of the engines are of a comparable

size in terms of volume but their weight and therefore, power-weight ratio varies. In

terms of weight reduction and high power-weight ratio, the two standout engines are the

UAV Engines AR741 and the Zanzottera Technologies 498ia, with the latter engine being

slightly lighter. The discussion from this point will be limited to these two engines.

Although both engines have a similar volume, the largest dimension of the AR741 is the

length, whereas the largest dimension of the 498ia is the width. The larger width will

make implementation more difficult as the cross-sectional area of the fuselage will

increase. Therefore, it is easier to design an aircraft around a long but slender engine such

as the AR741. It could also be said that piston engines have a longer life span than

Wankel rotary engines, but the AR741 has been specifically designed for extended use in

surveillance operations and has passed the FAR-33 type endurance test (Desert Aircraft,

2007).

The AR741 engine has a major disadvantage in the fuel consumption. The 498ia will

result in a much lighter aircraft due to less fuel and structure weight to carry the fuel (see

Sensitivity Analysis). However, the AR741 will produce a more aerodynamic fuselage

due to the geometry discussed above, and hence decrease the drag of the aircraft.

Although the low fuel consumption will be ideal for this application, the AR741 has been

chosen due its ease of implementation. A longer fuselage is also likely to produce a more

attractive aircraft.


-66-

Manufacturer

Model Power

(bhp)

Weight

(lbs)

Length

(inches)

Width

(inches)

Height

(inches)

Fuel

Consumption

(lbs/hp/hr)

UAV Engines (UAV

Engines ltd, 2004)

AR741 38 23 23.6 9.3 10.3 0.57

Bernard Hooper

Engineering ltd

SPV

580

40.6 40.12 14.17 14.33 14.53 0.57

Zanzottera

Technologies

498ia 39 18.564 10.23 18.639 10.266 0.33

Lightning Aircraft 302D2-

FI

28 33 17.5 14.5 12 2.75 gal/hr

Table 12-4: Engine selection table

12.7 Propulsion Integration

12.7.1 General Configuration

This section is concerned the positioning and mounting of the propeller engine. The

position of the engine is limited to two broad categories, tractor and pusher. A tractor

configuration has the propeller ahead of its installation point, where as the pusher

configuration has the propeller behind its point of installation. These configurations are

shown in Figure 12-9 and will be discussed for the remainder of this section.

Figure 12-9: General tractor and pusher configurations (Raymer, 1992)


-67-

Pusher

A pusher configuration has significant benefits with regard to aerodynamics. If the

propeller is mounted behind the fuselage, it reenergizes the boundary layer and forces the

flow to remain attached to the fuselage. This causes a reduction in form drag. The

efficiency of the main wings are also increased as they do not experience propeller wash.

The effectiveness of the tail is also increased as the pusher configuration will cause a high

velocity air stream over these surfaces.

The major disadvantage of a pusher configuration is the uneven distribution of weight

towards the tail (they become tail heavy). However, this can be accounted for by the

positioning of fuel and tail sizing. Another disadvantage associated with a pusher

configuration is encountered during take off and landing. This configuration is more

likely to be damaged by rocks and debris during take off and landing. Further to this, the

aircraft will require longer landing gear as the propeller will dip close to the ground when

the nose is lifted.

Tractor

A tractor configuration has a main advantage in propeller efficiency. As the propeller is

in front of the fuselage, it is placed in undisturbed air and hence the propeller efficiency is

higher. Tractor configurations also possess a more favourable weight distribution and

hence greater longitudinal stability than a pusher configuration. Also, this configuration

does not suffer from the restrictions associated with take off and landing that the pusher

configuration does.

The main disadvantage of a tractor configuration is the propeller plane endangers crew

and payload. As this is a UAV, there is no such crew or payload, however expensive

equipment such as control systems and cameras could be endangered during a propeller

mishap. This has the potential to place limitations on the propeller location if a tractor

configuration is chosen.


-68-

Selection

The configuration chosen for this aircraft is the pusher propeller. The major determining

factor in this decision is the landing mode (net landing, see Landing Section). A tractor

configuration is not feasible with a net landing as the propeller is likely to get caught in

the net, hence ruining the net and potentially the propeller. As mentioned, the major

disadvantages for a pusher configuration can be accounted for by fuel positioning and tail

sizing. The positioning of the propeller should also be simpler with a pusher

configuration as it is unlikely to endanger any sensitive equipment. The restrictions on

landing gear are also avoided due to the catapult take off method being implemented.

12.7.2 Position

Within the pusher configuration, the position of the engine needs to be specified. The

engine position is split into broad categories as shown in Figure 12-10 overleaf. The

fuselage position can decrease the fuselage wetted area, hence decreasing skin friction

drag. The wing position is limited to two engine aircraft, hence not applicable to our

application. The tail and pod configuration are used in exceptional circumstances, such as

seaplanes, where a large clearance is required between the take off surface and the

propeller. These two options produce a high thrust line which causes undesirable control

characteristics. Because of these reasons, a fuselage mounted engine is chosen.


-69-

Figure 12-10: Engine positions for pusher configuration (Raymer, 1992)

Fuselage position Wing position Pod configuration

Tail position


-70-

12.8 Take off Methods

The statistical analysis showed common launch methods for UAV’s. The most common

of these was the catapult launch method. The major advantage of the catapult launch is

the diversity of possible launch locations. The aircraft can be launched from any location

that the catapult unit can access. The major disadvantage of the catapult system is the

obvious cost associated with purchasing the unit.

The beaches which would implement the coast watch UAV are those which are located in

highly populated towns or cities. In these areas there are generally no spaces in which an

aircraft could perform a run-way take-off. The cost associated with purchasing a catapult

launching system in justified for the reason of providing a launch point conveniently

close to the area the craft is required to perform its loiter. Figure 12-11 shows the catapult

launch system for a military surveillance UAV.

Figure 12-11: Catapult Launch System for Surveillance UAV


-71-

12.9 Landing Methods

With the chosen launch method not incorporating a landing gear, a method of landing

which accounted for this needed to be designed. The statistical analysis provided a

number of possible solutions for landing methods. A runway landing was not an option as

this would need space which is not readily available as discussed above. The other

methods discovered during the statistical analysis were the parachute method, the net

method and the hook method.

The parachute method was not a feasible option as UAV’s which incorporate this method

of landing have a ‘crumple zone’ which takes the impact due to the downward velocity of

the landing aircraft. The crumple zone would be located where the surveillance

equipment needs to be so this method of landing wasn’t an option. The ‘crash landing’

into a net landing option was not considered as it is too easy for the crafts propeller to get

tangled in the net or the entire craft could bounce off of the net. The hook and cable

method was chosen as it is an easily designed and implemented system and has the least

chance of causing damage to the craft.

When the craft is coming into land, its engine shuts down and the craft glides to its

landing position which consists of two poles with cable between them as shown in Figure

12-12 overleaf. The aircraft deploys a hook on a cable and drags the hook along until it

becomes hooked on the landing apparatus.


-72-

Figure 12-12: Hook Landing System


-73-

12.10 Detachable Equipment Bay

In order to appeal to a variety of markets our UAV will exhibit a detachable equipment

bay. This bay will serve to house a variety of equipment depending upon the desired

operation of the aircraft. One of the main appeals of the bay will be to maintain a constant

aerodynamic profile and hence exhibit constant drag independent of the equipment in use.

The detachable bay is shown in Figure 12-13.

Figure 12-13: Detachable Equipment Bay


-74-

13 Weight and Stability Analysis

13.1 Weight Analysis

The weight of the aircraft was analysed using a combination of known values for some

components of the aircraft such as the propulsion system and statistical methods for the

unknowns, such as structural weight. A breakdown of the aircraft weight is shown in

Table 13-1.

Aircraft Weight Breakdown

Propulsion System 25lbs Based on engine and propeller selection

Fuel 64lbs Calculated from mission profile

Structural Weight 66lbs Based on Statistics (Wstructure=0.38WTO)

(Fuselage) 29.7lbs Wfuselage= 45%

(Empennage) 9.9lbs Wempennage= 15%

(Wing) 26.4lbs Wwing= 40%

Payload Weight 10lbs Based on available cameras to suit

application

Instrumentation 10lbs Based on Statistics (Wsystem=0.06WTO)

Total Take-Off Weight 175lbs

Table 13-1: Weight breakdown

The various components of the aircraft were positioned in order to produce favourable

stability characteristics. The following section discusses the reasoning behind the

positioning of the components.

Propulsion System

The position of the engine and propeller was essentially fixed based on the chosen

fuselage design and pusher propeller configuration. The engine was located at the rear of

the fuselage with the propeller directly coupled to the output shaft. This is the simplest


-75-

method of mounting the engine as it requires no heavy gearbox components that would

be required if the axis of rotations were not aligned.

Camera (Payload)

The camera and surveillance equipment was considered as a payload due to the fact that

the instruments included on the UAV are designed to be interchangeable, this allows for

change in the types of missions performed by the UAV and introduces flexibility into the

design. The payload makes up only a small portion of the total weight, hence it could be

placed almost anywhere on the aircraft without significantly affecting CG location.

However, the cameras had to be placed for maximum range of visibility, hence the

obvious location was below the fuselage as this gave the camera a 360° field of view.

Aesthetics were also considered when placing the cameras and it was decided by the

group that mounting the cameras below the fuselage was the most aesthetically pleasing

option.

Fuel Tank

The fuel tank made up the largest portion of the aircraft weight; therefore careful

consideration was given to its placement within the fuselage. Fuel tanks are commonly

placed in the wings, however the team decided this was not the most economical option

for a low-cost UAV, as it would add unnecessary complexity to the fuel delivery system.

In addition to this the quantity of fuel required is large and tanks in the wings do not have

sufficient volume to carry the required fuel. It was decided to use a cylindrical fuel tank

located in the centre of the fuselage, as this would minimise the amount of CG travel

during flight. Having the fuel tank located close to the engine also simplifies the fuel

delivery system, reducing weight and improving reliability.

Instrumentation

Instrumentation such as the UAV’s navigation system and power source was placed at the

front of the fuselage behind the nose. It was decided to locate the power source close to

the cameras to minimize the amount or wiring required. There are also benefits of


-76-

mounting this equipment at the front of the UAV as it balances the weight of the tail and

engine and brings the CG forward slightly.

Wing, Fuselage and Empennage

Structural components constitute 38 percent of the aircraft’s take-off weight. As the UAV

is launched by catapult there is no landing gear and the weight of the catapult release

mechanism was assumed to be negligible for this conceptual analysis. The empennage is

attached to the wing via two tail booms, thus the moment arm of the vertical and

horizontal stabilisers is fixed. To change longitudinal stability and static margin the wing

and tail can be moved. Moving the wing and tail back makes the aircraft more stable

while moving it forward brings the CG and neutral point closer together making it

unstable.

The calculation of the centre of gravity (CG) location was performed by considering the

weights and locations of each of the main components and calculating a weighted

average. The CG was calculated in the x and y directions only, it was assumed that the

weight was distributed evenly on both sides of the aircraft in the spanwise direction,

therefore eliminating the need to analyse zcg. The CG location is not constant, and

changes under different configurations, it is therefore necessary to calculate the CG

envelope. The CG envelope was determined by calculating the CG location for four

different configurations:

1. Empty weight only, no fuel or payload.

2. Empty weight and payload weight, no fuel.

3. Empty weight and fuel weight, no payload.

4. Take-off weight.

A spreadsheet was constructed in order to analyse each component of the aircraft and its

contribution to the total weight and CG (see Appendix A). Figure 13-1 was obtained by

analysing the four configurations mentioned above.


-77-

CG envelope

0

20

40

60

80

100

120

140

160

180

200

0 5 10 15 20 25 30 35 40 45

CG location (%MAC)

Aircraft W

eight (lbs)

Figure 13-1: Centre of Gravity envelope

The x component of the CG was found to vary between 9.65% MAC at take-off to

32.15% MAC at empty weight with no payload or fuel. When doing stability

calculations, the most aft CG is always used. The y-component of CG determines the

effect the propeller has on the stability of the aircraft. The thrust produced by the

propeller creates a moment about the CG if it is not directly on the axis of propeller

rotation. The location of the CG varies from between 33.5mm to 68.7mm above the line

of action of the propeller, meaning that there is always a small nose-up moment on the

aircraft. This propeller-moment term adds significant complexity to stability calculations,

as there is now velocity dependence, hence the small moment will be ignored for stability

calculations.

Take-Off (Most fore CG)

Empty Weight (Most aft CG)

Static Margin


-78-

13.2 Stability Analysis

An analysis of the longitudinal stability of the aircraft was conducted to ensure that the

UAV was statically stable and to make adjustments to the wing layout if the aircraft was

found to be unstable. An aircraft is stable in the longitudinal axis only if a positive change

in angle of attack produces a negative pitching moment about the aircraft centre of

gravity (CG). Therefore, mathematically, for stability:

0<αd

dCm

Airfoil Characteristics

The characteristics of the wing and tail were analysed in order to determine the neutral

point of the aircraft. The airfoil profile used for the main wing is the NACA 4415 with an

aspect ratio of 7 and the profile used for the tail is a NACA 0012 with an aspect ratio of

6. The lift-curve slope of these wings is shown in Table 13-2. The lift-curve slope of the

aerofoil was converted to a slope for a finite span wing using Equation 13-1.

AR

ddC

ddC

d

dC

l

lL

πα

αα +

=1

Equation 13-1: Lift curve equation conversion (Nelson, 1989)

Airfoil αd

dCL (rad-1)

Aspect Ratio

(AR) αddCl (rad-1)

αε

d

d (rad-1)

Wing - NACA 4415 4.173 7 5.151 0.38

Tail - NACA 0012 4.455 6 5.833 -

Table 13-2: Airfoil Lift-Curve Slopes (Modelfoil)

Downwash Angle

Downwash behind the main wing causes the effective angle of attack of the horizontal

tail to change. Downwash is a result of the bound and trailing vortices (Nelson, 1989), the

exact amount of downwash experienced by the tail depends on the location of the tail


-79-

relative to the main wing. The horizontal stabiliser of the Coastwatch UAV was placed

high on the tail in an attempt to avoid the vortices as much as possible. Downwash angle

for any lift coefficient can be estimated using Equation 13-2. Equation 13-3 calculates the

change in downwash angle with angle of attack.

w

Lw

AR

C

πε

2=

Equation 13-2: Downwash angle

w

Lw

AR

d

dC

d

d

πα

αε 2

=

Equation 13-3: Downwash angle with variation in angle of attack (Nelson, 1989)

Neutral Point

The neutral point of an aircraft is also known as the aircraft’s aerodynamic centre. It can

be considered as the point through which the lift force of the entire aircraft acts and also

the point about which moment stays constant with angle of attack. In general, the neutral

point is calculated by summing the moments about a point, located distance npx from the

wing leading edge. This equation is then differentiated with respect to angle of attack and

set to zero; the equation can then be solved for npx . This expression is shown in Equation

13-4.

−+−=αε

ηα

α

α

α

d

d

C

CV

C

C

c

X

c

X

wL

tLH

wL

fmacwnp1

Equation 13-4: Neutral point (Nelson, 1989)

Where:

acX is the distance from the wings leading edge to the aerodynamic centre of the wing.

η is the tail efficiency and is defined as the ratio of dynamic pressures of the main

wing and tail. For our calculations this is assumed to be unity.


-80-

fmC α is the contribution of the fuselage to the moment about the CG. Because the

fuselage is not designed to be a lifting surface, its contribution is quite small and

can be ignored for these basic stability calculations.

HV is the tail volume ratio and is given by:

cS

SlV tt

H =

tl is the moment arm of the horizontal stabiliser and is fixed at 1.5m

All other variables have already been defined.

A spreadsheet was used to calculate the neutral point of the aircraft based on Equation

7-1 (see Appendix A). Based on the configuration of the current conceptual design, the

location of the stick-fixed neutral point was found to be 0.157m aft of the wing leading

edge, i.e. the neutral point is at 39.25% MAC. This point is included on Figure 13-1.

Static Margin

The static margin of the aircraft is defined as the distance between the aircraft’s most aft

centre of gravity and the neutral point. The static margin is expressed graphically on

Figure 13-1 and is calculated using Equation 13-5.

c

XXSM

cgnp −=

Equation 13-5: Static Margin

%15.3225.39 −=⇒ SM

%1.7=∴SM


-81-

A static margin of 7.1% is sufficiently stable for a UAV, as the aircraft is presumably

computer controlled. Aircraft of the general aviation and homebuilt type have larger static

margins, as their pilots may be less experienced.


-82-

14 Performance Analysis and Conclusion The outcome of the project is a UAV with the performance parameters shown below. The

range and endurance are the same as specified in the mission profile and technical task.

Through the design process, several key parameters were improved; rate of climb was

increased allowing the UAV to reach operational altitude quicker. The designed UAV is

aesthetically pleasing while still meeting all requirements set by the technical task. The

appropriate hand calculations are shown in Appendix B.

Wing Geometry (mm)

Main Wing Tail Section

Aerofoil section NACA 4415 NACA 0012

Chord 400 180

Span 900 Horizontal – 900 Vertical – 600

Lever arm - 1700

Dihedral 0 0

Incidence angle 2 0

Sweep 0 0

Taper ratio 1 1

Twist 0 0

Aircraft Geometry (mm)

Fuselage length 1700

Fuselage width 420

Fuselage height 600

Aircraft length 2780


-83-

Performance Parameters

Range 54nm

Endurance 5 hours

Stall Speed

• Take-Off 73.17 knots (123.6 ft/s)

• Cruise 83.71 knots (141.41 ft/s)

• Landing 65.87 knots (111.26 ft/s)

Cruise Speed 52.4 knots (88.46 ft/s)

Time to Climb 27s

Rate of Climb 2196 fpm

Climb Gradient 0.3 rad

Absolute Ceiling 36500 ft


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15 References

Tech task and intro references

http://www.defenseindustrydaily.com/cobham-catches-a-1b-australian-coastwatch-

contract-01695/

http://www.customs.gov.au/site/page.cfm?u=4238

Australian Government Border Protection Command, Maritime Zones

http://www.customs.gov.au/webdata/resources/files/BPC_FactSheet_MaritimeZones1.pd

f

Campbell, CJ, 1991, The Golden Century of Oil, 1950-2050: The Depletion of a

Resource, Kluwer, Boston

Statistical Analysis

Jane’s All the World UAV’s ,2002, Jane’s Information Group.

Engine selection

UAV Engines ltd, 2004, Engine specifications, viewed 19 April 2008 <http://www.uavenginesltd.co.uk/index.php?id=397> Bernard Hooper Engineering ltd, Engine specifications, viewed 19 April 2008 <http://users.breathe.com/prhooper/spv580ds.htm> Lightning Aircraft Inc, 2007, Engine specifications, viewed 19 April 2008 <http://www.lightningaircraft.com/302D2-FI.htm> Zanzottera Technologies ltd, 2007, Engine specifications, viewed 19 April 2008 <http://www.zanzotteraengines.com/498.html>

Propeller selection

Desert Aircraft, 2007, Propeller Specifications, Viewed 13 May 2008,

<http://www.desertaircraft.com/page.php?Page=MSC>


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Raymer, D, 2006, Aircraft Design: A Conceptual Approach Fourth Edition, American

Institute of Aeronautics and Astronautics Inc, USA

Simons, M, 2002, Model Aircraft Aerodynamics Fourth Edition, Special Interest Model

Books, Great Britain

Zanzottera Technologies ltd, 2007, Engine specifications, viewed 19 April 2008 <http://www.zanzotteraengines.com/498.html>

Aerofoil selection

Lednicer, D 2007, The Incomplete Guide to Airfoil Usage, viewed 04/08/2007

<http://www.ae.uiuc.edu/m-selig/ads/aircraft.html>

Roskam, J 2005, Airplane Design, Roskam Aviation and Engineering Corp, Ottawa.

Stability References

Avalakki, N. Bannister, J. Chartier, B. Downie, T. Gibson, B. Gottwald, C. Moncrieff,

P.Williams, M. 2007. Design, Development and Manufacture of a Search and Rescue

Unmanned Aerial Vehicle. The University of Adelaide.

Nelson, R. 1989. Flight Stability and Automatic Control. United States. McGraw-Hill

Simons, M. 2002. Model Aircraft Aerodynamics Fourth Edition. Great Britain.: Special

Interest Model Books Ltd.

Hanley, P.E. 1999. Modelfoil Version 2.0. Software

Janes-Information-Group. 2002. Janes Unmanned Aerial Vehicles and Targets. USA:

Janes Information Group.

Raymer, D. 2006. Aircraft Design: A Conceptual Approach Fourth Edition. USA:

American Institute of Aeronautics and Astronautics Inc.


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16 Drawings See overleaf.

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Coast Watch UAV