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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
Aircraft Design Project Group 7
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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 8-2: CONCEPT 2 SKETCH .............................................................................................................23
FIGURE 8-3: CONCEPT 3 SKETCH .............................................................................................................24
FIGURE 8-4: CONCEPT 4, SKETCH 1 .........................................................................................................25
FIGURE 8-5: CONCEPT 4, SKETCH 2 .........................................................................................................26
FIGURE 8-6: CONCEPT 5, SKETCH 1 .........................................................................................................27
FIGURE 8-7: CONCEPT 5, SKETCH 2 .........................................................................................................28
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 8-2: Concept 2 sketch
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Figure 8-3: Concept 3 sketch
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Figure 8-4: Concept 4, sketch 1
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Figure 8-5: Concept 4, sketch 2
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Figure 8-6: Concept 5, sketch 1
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Figure 8-7: Concept 5, sketch 2
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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
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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
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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
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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
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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
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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.
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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.
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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
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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
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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
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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
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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.
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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
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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
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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
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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
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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
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Figure 12-1 Fineness Ratio terms
Figure 12-2: Fineness ratio for Subsonic Aircraft
Lf
Df
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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.
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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.
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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
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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.
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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.
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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.
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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.
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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
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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.
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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
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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.
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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
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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
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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.
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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)
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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≈
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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.
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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.
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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)
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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.
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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.
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Figure 12-10: Engine positions for pusher configuration (Raymer, 1992)
Fuselage position Wing position Pod configuration
Tail position
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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
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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.
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Figure 12-12: Hook Landing System
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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
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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
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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
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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.
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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
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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
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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.
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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
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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.
Aircraft Design Project Group 7
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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
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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>
Aircraft Design Project Group 7
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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.
Aircraft Design Project Group 7
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16 Drawings See overleaf.
Title:
Drawn By:
Drwg No.
Proj: Scale:
Date: Sheet:
Mechanical EngineeringNorth Terrace, Adelaide
All dimensions in mm unless otherwise stated.All tolerances 0.1mm or 1 unless stated
Coast Watch UAV
02-June-2008
Group 7
Rendered Image
A31
Title:
Drawn By:
Drwg No.
Proj: Scale:
Date: Sheet:
Mechanical EngineeringNorth Terrace, Adelaide
All dimensions in mm unless otherwise stated.All tolerances 0.1mm or 1 unless stated
Coast Watch UAV
02-June-2008
Group 7
2780
2800
400
838
600
970 1700
900
360200
8731700
600
420
340
500
Three View Drawing
1:20
2 A3
Title:
Drawn By:
Drwg No.
Proj: Scale:
Date: Sheet:
Mechanical EngineeringNorth Terrace, Adelaide
All dimensions in mm unless otherwise stated.All tolerances 0.1mm or 1 unless stated
Coast Watch UAV
02-June-2008
Group 7
Exploded View
3
1:13.33
AileronMain WingTail BoomVertical StabiliserHorizontal StabiliserElevatorRudderMotorPropellorHook CasingEquipment BayEquipmentLanding HookFuel TankNavigation EquipmentNose ConeMain Body
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Parts List
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