Coast Watch UAV
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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
Name
Id Number
Kelly Balnaves
1132985
Bradley Cook
1133395
Alex Horstmann
1131838
Ryan Middleton
1133404
Christian Rogers
1130940
Criteria
Signature
Mark (Total – 100) /10
Project Definition Research Activities
/15
Technical Calculations
/25
Drawings
/25
Format of the Report
/10
Novelty of the Solution
/15
Name
Group Mark
Peer Assessment
Kelly Balnaves Bradley Cook Alex Horstmann Ryan Middleton Christian Rogers
1
Total Mark
Aircraft Design Project
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1 Executive Summary This project presents the conceptual design and sizing of a CoastGuard 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 takeoff 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
6
7
5.1
AIMS .............................................................................................................................................8
5.2
SCOPE ...........................................................................................................................................8
5.3
BACKGROUND INFORMATION ........................................................................................................8
5.4
SIGNIFICANCE ...............................................................................................................................9
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
STATISTICAL ANALYSIS..............................................................................................................16 7.1
WING SPAN/AIRCRAFT LENGTH ..................................................................................................18
7.2
TAKEOFF 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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FAR 23.77 Climb Gradient Sizing ....................................................................................42
11.5
CRUISE SIZING ............................................................................................................................43
11.6
OVERALL SIZING CHART .............................................................................................................43
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
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 61: MISSION PROFILE.................................................................................................................12 FIGURE 71: EMPTY WEIGHT VERSUS TAKE OFF WEIGHT .........................................................................17 FIGURE 81: CONCEPT 1 SKETCH .............................................................................................................22 FIGURE 82: CONCEPT 2 SKETCH .............................................................................................................23 FIGURE 83: CONCEPT 3 SKETCH .............................................................................................................24 FIGURE 84: CONCEPT 4, SKETCH 1 .........................................................................................................25 FIGURE 85: CONCEPT 4, SKETCH 2 .........................................................................................................26 FIGURE 86: CONCEPT 5, SKETCH 1 .........................................................................................................27 FIGURE 87: CONCEPT 5, SKETCH 2 .........................................................................................................28 FIGURE 91: WEIGHT ESTIMATION GRAPH ...............................................................................................32 FIGURE 111: SIZING CHART ...................................................................................................................44 FIGURE 112 : SIZING CHART WITHOUT LANDING AND TAKEOFF ............................................................45 FIGURE 121 FINENESS RATIO TERMS .....................................................................................................47 FIGURE 122: FINENESS RATIO FOR SUBSONIC AIRCRAFT .......................................................................47 FIGURE 123: AEROFOIL LIFT CURVES (MODELFOIL) .............................................................................50 FIGURE 124 – NACA 4415 (MODELFOIL) ..............................................................................................51 FIGURE 125 – NACA 0012 PROFILE (MODELFOIL)................................................................................52 FIGURE 126: TWIN BOOM TAIL CONFIGURATION (HTTP://AEROWEB.LUCIA.IT/RAP/PARIS97)................55 FIGURE 127: VERTICAL STABILISER DIMENSIONS..................................................................................59 FIGURE 128: HORIZONTAL STABILISER DIMENSIONS .............................................................................60 FIGURE 129: GENERAL TRACTOR AND PUSHER CONFIGURATIONS (RAYMER, 1992) ..............................66 FIGURE 1210: ENGINE POSITIONS FOR PUSHER CONFIGURATION (RAYMER, 1992).................................69 FIGURE 1211: CATAPULT LAUNCH SYSTEM FOR SURVEILLANCE UAV .................................................70 FIGURE 1212: HOOK LANDING SYSTEM .................................................................................................72 FIGURE 1213: DETACHABLE EQUIPMENT BAY .......................................................................................73 FIGURE 131: CENTRE OF GRAVITY ENVELOPE .......................................................................................77
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3 Table of Tables TABLE 71: TAKE OFF WEIGHTS FOR SIMILAR AIRCRAFT .........................................................................17 TABLE 72: WING SPAN AND AIRCRAFT LENGTH .....................................................................................18 TABLE 73: TAKE OFF METHODS .............................................................................................................19 TABLE 74: LANDING METHODS ..............................................................................................................20 TABLE 75: SUMMARY OF STATISTICAL ANALYSIS ..................................................................................21 TABLE 91: TIME, FUEL CONSUMPTION AND POWER FOR EACH STAGE ....................................................30 TABLE 92: FUEL WEIGHT AND WEIGHT RATIO FOR EACH STAGE.............................................................31 TABLE 101: VALUES FOR SENSITIVITY EQUATIONS (FROM AIRPLANE DESIGN BY J. ROSKAM).............34 TABLE 102: VALUES USED FOR CALCULATING ‘C’ AND ‘D’...................................................................35 TABLE 103: REQUIRED VALUES FOR SENSITIVITY ANALYSIS ..................................................................35 TABLE 104: SENSITIVITIES TO THE MAIN PARAMETERS .........................................................................36 TABLE 111: TAKEOFF DISTANCE SOLUTION ..........................................................................................38 TABLE 112: LANDING DISTANCE SIZING.................................................................................................39 TABLE 113: LIFT COEFFICIENT FOR ALL CONFIGURATIONS ....................................................................39 TABLE 114: DRAG POLAR VALUES .........................................................................................................40 TABLE 115: ASPECT RATIO AND OSWALD EFFICIENCY FACTOR .............................................................40 TABLE 116: DRAG POLAR EQUATIONS FOR FAR23 CLIMB SIZING ..........................................................40 TABLE 117: FINAL CLIMB SIZING VALUES ..............................................................................................42 TABLE 118: CRUISE SIZING VALUES .......................................................................................................43 TABLE 119: FINAL SIZING VALUES .........................................................................................................45 TABLE 121: COMMON AIRFOIL SECTIONS FOR UAV AIRCRAFT (LEDNICER, 2007).................................49 TABLE 122: AIRCRAFT VOLUME COEFFICIENT DATA (AVALAKKI ET AL, 2007) ....................................57 TABLE 123: MAIN WING PROPERTIES ....................................................................................................57 TABLE 124: ENGINE SELECTION TABLE ..................................................................................................66 TABLE 131: WEIGHT BREAKDOWN ........................................................................................................74 TABLE 132: AIRFOIL LIFTCURVE SLOPES (MODELFOIL) ......................................................................78
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4 Table of Equations EQUATION 71: STATISTICAL ANALYSIS EQUATION (ROSKAM, 1994) .....................................................16 EQUATION 91: TAKE OFF WEIGHT EQUATION USING STATISTICS ............................................................29 EQUATION 92: GENERAL TAKE OFF WEIGHT EQUATION .........................................................................29 EQUATION 93: FUEL WEIGHT EQUATION ................................................................................................30 EQUATION 94: MASS FUEL FRACTION ....................................................................................................30 EQUATION 95: MISSION STAGE FUEL WEIGHT ........................................................................................31 EQUATION 96: FINAL TAKE OFF WEIGHT EQUATION ...............................................................................31 EQUATION 101: TAKEOFF WEIGHT SENSITIVITIES FOR RANGE AND ENDURANCE CASES .....................33 EQUATION 102: EQUATION TO CALCULATE F ........................................................................................33 EQUATION 103: EQUATION FOR ‘C’ CALCULATION ...............................................................................34 EQUATION 104: EQUATION FOR ‘D’ CALCULATION ...............................................................................34 EQUATION 111: STALL SPEED EQUATION ...............................................................................................37 EQUATION 112 TAKEOFF SIZING EQUATION ..........................................................................................38 EQUATION 113: ZERO LIFT DRAG COEFFICIENT ESTIMATION ..................................................................39 EQUATION 114: DRAG POLAR EQUATION ...............................................................................................40 EQUATION 115: FAR23 ROC SIZING EQUATION.....................................................................................41 EQUATION 116: FAR23.65 CG SIZING EQUATION ..................................................................................41 EQUATION 117: FINAL EQUATION FOR FAR23.67 SIZING ......................................................................42 EQUATION 121: FINENESS RATIO ...........................................................................................................46 EQUATION 122: VERTICAL TAIL VOLUME COEFFICIENT .........................................................................56 EQUATION 123: HORIZONTAL TAIL VOLUME COEFFICIENT.....................................................................56 EQUATION 124: VERTICAL TAIL AREA ...................................................................................................58 EQUATION 125: HORIZONTAL TAIL AREA...............................................................................................58 EQUATION 126: PROPELLER TIP VELOCITY EQUATION (RAYMER, 2006) ................................................62 EQUATION 127: REARRANGED PROPELLER TIP VELOCITY EQUATION (RAYMER, 2006)..........................62 EQUATION 131: LIFT CURVE EQUATION CONVERSION (NELSON, 1989)..................................................78 EQUATION 132: DOWNWASH ANGLE ......................................................................................................79 EQUATION 133: DOWNWASH ANGLE WITH VARIATION IN ANGLE OF ATTACK (NELSON, 1989)..............79 EQUATION 134: NEUTRAL POINT (NELSON, 1989) .................................................................................79 EQUATION 135: 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 AP3C 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 IslandTorres Strait and Gove (www.customs.gov.au). The distance between each of these bases 10
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is approximately 1500km (travelling by coast); therefore the UAV will need to have longrange 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 11
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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 61.
Figure 61: 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 14
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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 71, where A and B are empirical constants for a particular type of aircraft. We are concerned with UAV’s in this project.
Equation 71: Statistical Analysis equation (Roskam, 1994)
This report will use Roskam’s method for estimating takeoff 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 71 from which the values for A and B can be found. The ten aircraft listed in Table 71 were selected based on our initial weight and size estimated requirements for our proposed aircraft.
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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 71: Take off weights for similar aircraft
log(We) vs. log(Wto)
3 2.5 y = 1.0321x  0.2917 log(We)
2 1.5 1 0.5 0 0
0.5
1
1.5
2
log(Wto)
Figure 71: Empty weight versus take off weight
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2.5
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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 72.
Aircraft
Wing Span Aircraft (m)
(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 3.66
Length
2.74
mini Aerosky
4
n/a
Table 72: Wing span and aircraft length
Analysing Table 72, 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 Takeoff Methods UAV’s can use different methods of takeoff due to their light weight and ease of transport. Larger UAV’s tend to use the common method of a runway takeoff as they are too heavy and large for other methods. By studying the ten similar UAV’s (Table 73), 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
Takeoff Distance
Sting
Bungee Launch
EMT Luna
4m Bungee Launch
Aerosonde
Catapult/Car
top
40knots KAI
N/A
AAI Shadow
Hydraulic Catapault
AAI Pioneer
Catapault
Silver Arrow
82m
BAE Phoenix
Hydraulic Catapault
Silver Arrow 6m Catapault mini Aerosky
N/A Table 73: Take off methods
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7.3 Landing Methods Table 74 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 Parachute and Replacable Nose Cone mini Aerosky
Conventional wheeled landing Table 74: Landing methods
Table 74 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 75 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 75: Summary of statistical analysis
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8 Concept Sketches
Figure 81: Concept 1 sketch
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Figure 82: Concept 2 sketch
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Figure 83: Concept 3 sketch
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Figure 84: Concept 4, sketch 1
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Figure 85: Concept 4, sketch 2
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Figure 86: Concept 5, sketch 1
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Figure 87: 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 91. log(WTO ) = 0.28263 + 0.9689 log(Wempty ) Equation 91: 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 92. WTO = Wcrew + W payload + W fuel + Wempty Equation 92: General take off weight equation
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Equation 92 can then be solved along with Equation 91 to form an initial estimate for the take off and empty weight. Firstly, all terms in Equation 92 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 93. W fuel = 1.06WTO (1 − M ff
)
Equation 93: Fuel weight equation
The mass fuel fraction is calculated using weight ratios for each mission stage of the UAV, as shown in Equation 94.
W M ff = 1 WTO
n Wi +1 ∏ W i = 1 i
Equation 94: 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 91. The values for cp and P have been taken from the Engine selection section.
t (hr)
cp (lbs/hp/hr)
P (bhp)
Startup
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 91: 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 95. Fuel weight = P × t × c p
(lbs )
Equation 95: 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 92, which is then solved with Equation 91 to find the take off and empty weight. This process is iterated until the initial guess matches the final value. Table 92 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
Startup
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 92: Fuel weight and weight ratio for each stage
Using these number, the mass fuel fraction is calculated using Equation 94 and is 0.6336. Equation 92 can now be written as shown in Equation 96. WTO = 1.6351 × Wempty Equation 96: Final take off weight equation
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The graph of Equation 91 and Equation 96 is shown in Figure 91. 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.
Weight Estimation 400 350 Wto (lbs)
300 250 200 150 Calculation
100
Statistics
50 0 0
50
100
150 We (lbs)
Figure 91: 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 takeoff weight can be found using Equation 101 where F is defined in Equation 102. The values of
∂R ∂E and can be found using ∂y ∂y
Table 101 for the propeller driven aircraft case.
∂Wto ∂R =F ∂y ∂y
∂Wto ∂E =F ∂y ∂y
Equation 101: Takeoff Weight Sensitivities for Range and Endurance Cases
F = − B(Wto) 2 (CWto(1 − B) − D) −1 (1 + M reserve ) M ff Equation 102: Equation to calculate F
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Table 101: 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 103 and Equation 104. The values used to calculate C and D are shown in Table 102. These values are also used in the calculation of the sensitivities as can be seen in Table 101. The values obtained for A, B, C and D were 0.28263, 0.9689, 0.6286 and 0 respectively.
C = 1 − (1 + M reserve )(1 − M ff ) − M Funuseable Equation 103: Equation for ‘C’ Calculation
D = WPL + Wcrew Equation 104: Equation for ‘D’ Calculation
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W TO
272.23
We
166.5
Mreserve
0
mff
0.633558
mfunusable
0.005
W PL
0
W crew
0
Table 102: Values used for calculating ‘C’ and ‘D’
The following values displayed in Table 103 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 103: Required values for sensitivity analysis
Using Equation 102, F is calculated as 5412.11. This value can then be used in the sensitivity equations shown in Table 101 to calculate
∂R ∂E and for each of the ∂y ∂y
parameters mentioned. Table 104 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
L/D
48.23
11.73
Table 104: Sensitivities to the Main Parameters
Range case The sensitivities in Table 104 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 welldefined 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 1 2 = ρVstall CL max S 2 Equation 111: 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 takeoff sizing is given in Equation 112. In this case, takeoff is occurring at h=0, therefore σ = 1.
W hp = TOP23 ⋅ σ ⋅ CL max TO ⋅ S TO W TO Equation 112 Takeoff sizing equation
Equation 112 is solved and the results are displayed below in Table 111.
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
Table 111: Takeoff distance solution
11.3 Landing distance sizing The values for landing distance sizing are given in Table 112.
Sizing to FAR 23 Landing Req Slg
500
ft
Sl
969
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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 112: 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 113.
CLTO MAX
0.85
CLTO
0.45
CLland MAX
1.05
CLland
0.56
CL cruise
0.65
Table 113: 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 113. 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 114.
CD 0 = C fe
Swet Sref
Equation 113: Zero lift drag coefficient estimation
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Cfe
0.01
Swet/Sref
5
CD0
0.05
Table 114: Drag polar values
The zero lift drag coefficient can then be used with Equation 114 to help determine the Coefficient of Drag (CD). CD = C D 0 +
1 2 ⋅ CL π ⋅ A⋅e
Equation 114: Drag polar equation
Here, ‘A’ denotes the Aspect Ratio and ‘e’ denotes the Oswald efficiency factor which are both assumed and shown in Table 115.
A
7
e (Takeoff)
0.8
e (Landing)
0.75
Table 115: 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 116. 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
C D = 0.05 + 0.015 +
1 2 ⋅ C Ltakeoff π ⋅ 7 ⋅ 0 .8
Landing
n/a
C D = 0.05 + 0.065 +
1 2 ⋅ C Llanding π ⋅ 7 ⋅ 0.75
Table 116: Drag polar equations for FAR23 climb sizing
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11.4.1
Group 7
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 115. ηp (W /S)1/ 2 − RCP = 3/2 1/ 2 (W / P) 19((CL ) /CD ) max σ Equation 115: 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 116.
(CGR + (L / D) ) CGRP = −1
CL
1/ 2
Equation 116: FAR23.65 CG sizing equation
The relationship between W/S and W/P is then found using Equation 117. 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 ⋅ η p ⋅ σ 1/ 2 (W /P)(W /S)1/ 2
Equation 117: 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 117.
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 117: 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 takeoff power by multiplying 0.75. The values are shown in Table 118.
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 118: Cruise sizing values
11.6 Overall Sizing Chart The overall sizing chart is shown in Figure 111. From Figure 111 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 112.
Figure 111: Sizing Chart
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Figure 112 : 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 119.
W
175 lbs
S
11.7 ft2
PTO
20.6 hp
Table 119: 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 121. The terms are shown in Figure 121. 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 122. 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 121: Fineness ratio
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Df
Lf
Figure 121 Fineness Ratio terms
Figure 122: Fineness ratio for Subsonic Aircraft
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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 reenergise the boundary layer of the aftend 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 crosssection 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 121 shows some small aircraft and the aerofoil which they use for the main wing.
Aircraft
Aerofoil
AAI AA2 Mamba
NACA 4412
AAI Shadow 200
NACA 4415
AAI Shadow 400
NACA 4415
BAI Aerosystems Dragon Drone
NACA 63A012
Table 121: Common airfoil sections for UAV aircraft (Lednicer, 2007)
As can be seen in Table 121, 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 AT802 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 121, along with two symmetrical aerofoils (0012 and 0016) will be further investigated to consider their lift and drag qualities. The liftcurves of the aerofoils are shown in Figure 123.
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Figure 123: 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 123. 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 50
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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 124 – 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 125.
Figure 125 – 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 highwing 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 midwing aircraft will also be negligible as our aircraft is small and will be flying at low speeds. A highwing position will also be lighter than a midwing position, allowing for a greater range.
Based on the parameters discussed above, a highwing 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 nonswept wing will be negligible, therefore, the wing of our UAV will be nonswept.
Wing Aspect Ratio While a high aspect ratio wing gives a lower induced drag and high liftcurve slopes, low aspect ratio wings exhibit better aeroelastic 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 highaspect 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 highwing 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 123, 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 126.
Figure 126: 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 55
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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 122 and Equation 123 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).
SV =
VV bw S w Lvt
Equation 122: Vertical tail volume coefficient
SH =
VH C w S w Lht
Equation 123: 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 122.
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Aircraft
VH
VV
Reference
BAI ’Javelin’
0.6364
0.0372
JanesInformationGroup (2002)
INTA ’Alo’
0.5935
0.0337
JanesInformationGroup (2002)
Aerosonde ’Aerosonde’
0.93
0.0201
JanesInformationGroup (2002)
Homebuilt Aircraft
0.50
0.04
Raymer (2006)
General Aviation Aircraft
0.70
0.04
Raymer (2006)
Average
0.67
0.035
Table 122: Aircraft Volume Coefficient Data (Avalakki et al, 2007)
The values used for the coast watch UAV were selected using Table 122 and are:
V H = 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 123.
Aspect Ratio
7
Span
2.8 m
Chord
0.4 m
Sweep
0
Area
1.12 m2 Table 123: Main Wing Properties
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Raymer suggests that for aftmounted 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 122, vertical tail area becomes:
S vt =
0.04 × 2.8 × 1.12 1 .5
S vt = 0.084m 2 Equation 124: Vertical tail area
Using Equation 123, horizontal tail area becomes:
S ht =
0.6 × 0.4 × 1.12 1 .5
S ht = 0.18m 2 Equation 125: Horizontal tail area
The areas calculated by Equation 124 and Equation 125 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 124 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 127 and a total area of 0.132m2.
Rudder
0.11m
0.48m
0.6m
Figure 127: 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 128 .Total horizontal stabiliser area is 0.162m2, slightly less than that calculated by the volume coefficient method.
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0.9m
0.63m 0.18m
Elevator Figure 128: 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).
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Celevator = 0.3Cht ⇒ C elevator = 0.054m
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 twinboom 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 127. 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. C rudder = 0.3C vt
⇒ C rudder = 0.035m
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 126. 2 Vtip ,helixical = Vtip2 + Vcruise
Equation 126: Propeller tip velocity equation (Raymer, 2006)
The velocity limitation depends on the material of the propeller, as shown below. •
Wooden propellers, Vtip ,helixical < 850 fps
•
Metal propellers, Vtip ,helixical < 950 fps
But in order to reduce propeller noise, it is recommended that Vtip ,helixical < 700 fps (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. Vtip ,helixical < 700 fps . 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 126 is rearranged to from Equation 127, which is used to calculate the propeller diameter.
2 Vtip = Vtip2 ,helixical − Vcruise
Equation 127: Rearranged propeller tip velocity equation (Raymer, 2006)
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The following are calculations used to size the propeller: Vtip ,helixical = 700 fps = 213.36m / s Vcruise = 150km / hr = 41.67 m / s ⇒ Vtip = 209.25m / s A safety factor of 10% will be used to allow for wind gusts 209.25 = 190.23m / s 1 .1 Now, maximum diameter
∴Vtip =
Dmax =
60Vtip
πn
Where n is engine speed in rpm
Maximum engine power is produced at 6700rpm (Zanzottera Technologies ltd, 2007), hence the maximum diameter can now be calculated. ∴ Dmax = 0.5422m ≈ 21.34inch
However, as can be seen in the technical drawings, a propeller of this size will be ineffective due to the larger crosssectional 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
≈ 460
100
450
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, 63
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⇒ Dmax =
205.25 × 60 π × 4200
= 0.952m Incorporating the 10% safety factor, gives a diameter of: D = 0.85mm ≈ 33.5inch
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 powerweight 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 124. 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, powerweight ratio varies. In terms of weight reduction and high powerweight 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 crosssectional 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 FAR33 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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Model
Power
Weight
Length
Width
Height
Fuel
(bhp)
(lbs)
(inches)
(inches)
(inches)
Consumption (lbs/hp/hr)
UAV Engines (UAV
AR741
38
23
23.6
9.3
10.3
0.57
SPV
40.6
40.12
14.17
14.33
14.53
0.57
Engines ltd, 2004) Bernard
Hooper
Engineering ltd
580
Zanzottera
498ia
39
18.564
10.23
18.639
10.266
0.33
302D2
28
33
17.5
14.5
12
2.75 gal/hr
Technologies Lightning Aircraft
FI Table 124: 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 129 and will be discussed for the remainder of this section.
Figure 129: 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 1210 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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Fuselage position
Wing position
Pod configuration
Tail position
Figure 1210: Engine positions for pusher configuration (Raymer, 1992)
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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 runway takeoff. 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 1211 shows the catapult launch system for a military surveillance UAV.
Figure 1211: 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 1212 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 1212: 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 1213.
Figure 1213: 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 131.
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
Total TakeOff Weight
175lbs
Based on Statistics (Wsystem=0.06WTO)
Table 131: 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 lowcost 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 takeoff 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. Takeoff 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 131 was obtained by analysing the four configurations mentioned above.
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CG envelope 200
180
160
TakeOff (Most fore CG)
Aircraft Weight (lbs)
140
120
100
80
Empty Weight (Most aft CG)
60
40
Static Margin
20
0 0
5
10
15
20
25
30
35
40
45
CG location (%MAC)
Figure 131: Centre of Gravity envelope
The x component of the CG was found to vary between 9.65% MAC at takeoff to 32.15% MAC at empty weight with no payload or fuel. When doing stability calculations, the most aft CG is always used. The ycomponent 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 noseup moment on the aircraft. This propellermoment term adds significant complexity to stability calculations, as there is now velocity dependence, hence the small moment will be ignored for stability calculations. 77
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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:
dC m
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