high altitude long endurance recon uav

HIGH ALTITUDE LONG ENDURANCE UAV

Submitted by -

Sattwik Suman Das (SC08B108) Shashank.S (SC08B098)Tanveer Ali (SC08B003)

Dept. of Aerospace Engineer ingIndian Inst itute of Space science and Technology

Submitted as part of AE412 Aerospace Vehicle Design Project

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AE 412 AEROSPACE VEHICLE DESIGN PROJECT REPORT

1. Introduction:

Over the years, infiltration has increased manifold over India’s various borders. On the western front

along the Indo-Pak border, infiltration is a very serious concern due to the drastic increase in

terrorism over the years. Gross infiltration was estimated to be 342 in 2008, while it was 485 in

2009. Along the Indo-Bangladesh border, infiltration and subsequent illegal immigration remains a

serious issue. Rising military instalments along the Indo-China border also remains a cause for

concern for India’s defence forces.

The aim of our project is to address all these issues by making a High Altitude Long Endurance (HALE)

Unmanned Aerial Vehicle with Infrared sensing capabilities to check infiltration and report military

installations across India’s many borders. Alternatively, this UAV can also be used in search and

rescue missions during aircraft crashes etc.

Following are the requirements as specified by the customer:

1.1 Mission Capabilities

Patrol Area: 6500 m2

Patrol Duration: 40 hours of loiter

Weight Class: 1000-2000 kg

Launch Type: Conventional Runway (maximum 600m)

1.2 Performance Capabilities

Operational Ceiling: 19.8 kms (65000 ft)

Cruise Speed: 147 km/hr

Max Payload Weight: 113 kg

Rate of Climb: 5 m/s

The Primary factors that will be emphasized throughout the design of this UAV are its ability to

satisfy the mission requirements and the total cost when compared to existing systems

manufactured by foreign players in the market.

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2. Mission Description and Analysis

To properly design this UAV for its mission, analysis must be conducted to gain insight on the flight

manoeuvres required for the UAV. This section will outline a mission and calculate the duration,

flight speed, and distance covered over the various manoeuvres.

The flight can be broken up into three separate manoeuvres:

1. Takeoff, dash to surveillance area and climb to cruise altitude

2. Loiter for a total time of 40 hours

3. Cruise back to base, descend, and land

The UAV will initially take off and then dash at a speed of 147 km/hr to the centre of its 6500 sq. km.

surveillance area and follow the loiter pattern as shown below.

Figure 1: Loiter patern of the UAV

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Figure 2: Mission Profile of the UAV

2.1 Payload Analysis:

The requirements state that the maximum payload weight on the UAV can be 113 kg. Based on this

and based on the surveillance requirements, we have chosen the following payloads:

1. Synthetic Aperture Radar (SAR): The synthetic aperture radar is used for long range target

identification; therefore, a range of about 35 km is required to survey a 6500 sq. km. area.

We have chosen the Sandia National labs MiniSAR synthetic Aperture Radar* as it weighs

only 14 kg.

Figure 3: Sandia Labs MiniSAR

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2. Electro-Optic-Infrared Sensor: The electro-optic-infrared (EOI) sensor is used for close range

target identification. The Advanced EO/IR sensor from APM UAV Payloads Inc. is chosen. It

weighs 23 kg.

3. Data link: A line of sight data link is required to transmit the data collected by the EOI sensor

and synthetic aperture radar. The data link is also needed for communicating with the UAV.

The data link chosen for our UAV is the UAV Data Link by L-3 communications systems. This

data link is ideal because it has been used previously on other UAVs, has line of sight

capabilities and weighs less than a kg.

The payload package of the synthetic aperture radar, the EOI sensor, and line of sight data link

can meet all of the mission requirements. These three pieces of equipment in total weigh about

38 kg. This leaves 75 kg for auxiliary batteries which can be used in case of engine failure to

control the control surfaces of the UAV while gliding back safely to base.

2.2 Initial Sizing

From the given requirements

a) Cruise Speed = 147 km/h = 91.34 m/s

b) Endurance = 40 hrs

The weight fractions are found from “Airplane design by Jan Roskam” for a military patrol

aircraft to be:

Taxi – W1/Wo = 0.99

Take off – W2/W1 = 0.995

Climb – W3/W2 = 0.98

To find out the weight fraction for Loiter,

We have p = 0.77

Cp = 0.6

L/D = 16

Using Breguet’s Endurance Equation and incorporating above values, we have W4/W3 =

0.622

For descent, again from Roskam, we have Descent weight fraction, W5/W4 = 0.99

Similarly, for Landing-Taxi-Shutdown; the fuel fraction is W6/W5 = 0.992

Now, WE/WTO = 0.5896 (This gives WE(allowed))

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A graph was plotted between WE and WTO using data of various similar UAV as shown in the table

below.

Figure 4: Data of existing UAVs

Regression was used to find WE(Tentative) for different values of WTO.

When WE(Allowed) and WE(Tentative) were within 0.09% of one another, that value of WE was taken as the

final value and corresponding WTO was accepted as the final WTO for this stage of design.

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2.3 Airfoil Selection

To proceed with the design of an aircraft, some of the critical parameters like lift-to-drag ratio, CL

needed to be found out. For, this we had to choose an Airfoil. In order to judge the performance, we

have chosen the following measures.

2.3.1 Airfoil Selection Criteria

The performance of the Airfoil is measured using the criteria - maximum lift coefficient, aerodynamic

efficiency, and off-design aerodynamic performance. The next three sections outline the importance

of each of these criteria.

2.3.2 Maximum Lift Coefficient

One of the most desirable characteristics of the selected airfoil is its lift coefficient. The lift

coefficient dictates how well the aircraft will generate lift during lift-intensive manoeuvres, such as

take-off and landing. Given that the airfoil generally has a lift coefficient higher than that of the

entire wing, the airfoil to be chosen has to have a maximum lift coefficient higher than the value of

the wing. In addition to meeting the maximum lift coefficient requirement, we would like an airfoil

which has superior lift characteristics in order to minimize the wing area.

Figure 5: Plotted graph of data from previous table

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2.3.4 Aerodynamic Efficiency

The second most important criterion is the aerodynamic efficiency, given by the maximum lift-to-

drag ratio. To reduce drag and thereby conserve fuel, the aircraft will have to fly in such a state as to

achieve maximum aerodynamic efficiency. Since the aircraft spends the majority of its flight time

loitering, the airfoil selected must have the highest aerodynamic efficiency at loitering conditions.

2.3.5 Off-design Aerodynamic Characteristics

The final criterion we considered was off-design performance of the airfoil. Having a high efficiency

at a single angle of attack does not guarantee reasonable aerodynamic performance throughout the

entire flight envelope. Therefore, the airfoil should have a reasonable lift-to-drag ratio over a broad

range of angles of attack. The airfoil must also be able to operate over a wide range of conditions.

2.4 Analysis of Airfoil

After the analysis of a wide range of airfoils, which were categorized as high lift and low drag airfoils,

we chose the NACA 23015 illustrated in Fig. 6

Figure 6: NACA 23015

The following airfoils were analysed:

NACA 5 digit 63 series and 23 series

NASA General Aviation airfoil series

The aerodynamic data was obtained from a program called JAVAFOIL. But, this software had

limitations of modelling the flow separation during stall. However, it used empirical results to find

out the stalling angle. Based on the stalling behaviour and the aerodynamic efficiencies, we chose

the NACA 23015 airfoil. This airfoil was analysed in ANSYS FLUENT for various angles of attack at the

expected Reynolds’s number of 1x107 and the data plotted as shown in Fig. 7 and Fig. 8

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Figure 7: CL vs. Alpha for NACA 23015

Figure 8: Drag Polar for NACA 23015

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Parameter Value

Max CL 1.173

Max CL angle 15.0

Max L/D 37.011

Max L/D angle 5.0

Stall angle 8.5

Zero-lift angle -1.0

Figure 9: NACA 23015 airfoil at an angle of attack 9 degree

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3. Wing Design

After selection of the airfoil, the next major step is to make the finite wing using the airfoil. The wing

configuration chosen was low wing and owing to the low Mach number, the wing did not require a

sweep back. Although the low wing has an inherent disadvantage of bad field performance in rough

terrain, it was ignored as the military base runways are always paved. The wing parameters obtained

are given below. Note that these are for the angle of attack at which endurance is maximum i.e. 6o,

unless mentioned otherwise.

Parameter Value Obtained from

Span Effectiveness Factor, e 0.8 Historical Data

Zero-Lift Drag, CDo 0.008 Airfoil Selection

Lift Coefficient, CL 0.55619 Calculation

Aspect Ratio, AR 20.77 Empirical Relation

Taper Ratio, 0.4 Historical Data

Wing Loading, W/S 169.2 kg/m2 Constraint Analysis

Wing Reference Area, S 9.323 m2 Calculation

Lift-to-Drag ratio, L/Dmax 40 Calculation

Span, b 13.91 m Calculation

Root chord, cr 1.74 m Raymer’s formula

Tail chord, ct 0.696 m Raymer’s formula

Mean Aerodynamic Chord, ̅ 1.29 m Raymer’s formula

MAC position, ̅ 5.44 m Raymer’s formula

Figure 10: Wing details

Fowler flaps were chosen for the wing as they are the least complex and provide a considerably

good amount of lift. The flapped wing area, Sflapped, was determined to be 4.94 m2 from empirical

relations.

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4. Constraint analysis

Constraint analysis is carried out to determine the T/Wo and Wo/S values that meet the user

specified requirements like stalling speed, take-off and landing distances, cruise Mach number and

ceiling and the air worthiness requirements such as the missed approach gradient and the second

stage gradient. Wo/S depends is subjected to ceiling, stalling and landing field length constraints.

4.1 Stalling:

W/S= 1/2 *ρ*Vstall2*CLmax

Constraint on stall puts upper limit on W/S.

4.2 Landing distance constraint:

Slanding(ft)=80*(W/S(lb/ft2))(1/σCLmax)+ Sa(ft)

Figure 11: Constraints (MATLAB Code in Appendix)

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Constraint on the landing field length puts an upper limit on W/S.

4.3 Constraint on ceiling:

(W/S)=q√

Constraint on ceiling puts a lower limit on W/S.

Take off ground roll puts a constraint on both W/S and hp/W given by:

W/S=(TOP)σCLmax(hp/W)

Where, TOP= take off parameter= Take off Field length/k. All the above constrains are graphically

represented by the graphs shown here and the design point is suitably selected.

Figure 12: P/W vs. W/S- Results of Constrain analysis (MATLAB Code in Appendix)

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4.4 Results of Constraint Analysis:

Wing Loading= W/S= 169.2 Kg/m2=34.65lb/ft2.

Loiter Velocity= Vloiter=62.5m/s.

Power to weight ratio= P/W=0.11hp/lb.

From this, we get that the power required is over 300 hp.

As per the constraint analysis, the power required turned out to be more than what all the available

IC engines could deliver. The IC engines also suffered from the problem of drastic power reduction at

high altitudes (altitudes above 10 km) even with a turbo super charger. This forced us to look for an

alternative propulsion system and we chose the turboprop due to its high fuel efficiency at low

subsonic Mach numbers when compared to turbofan and turbojet. This required us to perform the

initial sizing again.

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5. Initial sizing II

For this initial sizing, we used DARCORP’s Advanced Aircraft Analysis software v2.5

The screenshot of the same is shown below with the required figures

Figure 13: Weights of different UAVs

Figure 14: Input Parameters

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5.1 Result

From this, we obtain the weight of the UAV as 1579.15 kg. This weight will be used in calculations

hereafter.

Figure 15: Output Weights

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6. Tail sizing

6.1 Introduction

Tails provide trim, stability and control to an aircraft. Trim refers to the generation of lift forces that,

by acting through a moment arm about the centre of gravity, balances the other moment produced

by the aircraft.

The following empennage arrangements were considered for the pusher configuration of our UAV:

V-Tail: Theoretically, the V-tail reduces wetted area and will be beneficial to the aircraft design by

reducing the aircraft weight. Also, the interference drag and spiralling tendencies are significantly

reduced when we use a V-tail design. However, extensive NACA research suggests that the V-tail

surfaces need to be enlarged so that they have the equivalent wetted area as a conventional design

in order to provide good stability and control.

H-Tail: A twin boom configuration was considered because such a configuration can accommodate a

pusher prop layout while allowing the heavy engine to be located near the centre of gravity of the

aircraft. The long slender booms also allow the tail of the aircraft to be positioned farther aft of the

wing, maximizing the moment arm of the tail surfaces without having to incur the full weight penalty

of building an equivalently long fuselage. The twin boom tail configuration however could force the

wing structure to be more robust than a traditional design, because the booms are usually fixed to

the wing. Also, the booms could create additional wetted area, which could increase the drag on the

aircraft. Figure 16 illustrates the twin boom tail configuration. However, tail will be directly exposed

to the propeller wash.

Figure 16: A typical H tail configuration

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Only the above two were considered because a conventional tail design will not will not cater to a

pusher configuration. The H-Tail would position the vertical tails in undisturbed air during high angle

of attack conditions. However, the drawback is that the H-Tail is comparatively heavier.

The V-Tail would be structurally lighter, and can be mounted easily with the pusher configuration

with an engine at the aft of the fuselage. A conventional tail would require us to have the vertical tail

mounted over the engine section in the fuselage and that will make it structurally weaker. V-Tails

have reduced wetted area and reduced induced drag however it brings in complexity of control.

Considering all these factors, we choose the V tail for our UAV.

Sl.No. Aircraft Wing

Ref

Area,

S (Ft2)

Horizontal

Tail Area,

Sh

(Ft2)

Proportionally

Reduced Value

(Ft2)

Vertical

Tail Area,

Sv

(Ft2)

Proportionally

Reduced Value

(Ft2)

1 Lockheed C130E 1745 536 102.49 300 57.36619

2 Lockheed P3c 1300 322 82.649 176 45.17514

3 Antonov An12 BP 1310 319 81.254 205 52.2171

4 Antonov An26 807 213 88.05 171 70.70543

5 Grum E2c 700 174 82.94 199 94.86046

6 Aerital G222 889 255 95.71 207 77.69602

Average Value 88.84883 66.33672

Figure 17: Data from various Aircrafts

The conventional tail design is carried from historical data as tabulated above. However for our UAV,

we have chosen a V-Tail configuration because in the H-Tail configuration the horizontal tail would

be exposed directly to the propeller wash.

For the V-Tail sizing the area of the tail is such that the projection of the tail on the horizontal and

vertical planes will be equal to the values obtained for the conventional tail analysis. The area of the

V-Tail is thus obtained to be 5.27m2. The dihedral angle or tilt angle of the V-tail with the horizontal

is calculated as shown below:

Г = tan-1(Sv/Sh) = 36.74o.

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7. Landing gear

7.1 Introduction

For our UAV, given the fact that we will be having a turboprop engine at the back of the aircraft, we

have opted for the Tricycle configuration landing gear for our aircraft. The landing gear system will

be retractable with a simple actuator mechanism for retraction.

According to “Aircraft Design: A conceptual approach” by Raymer, typically the main tires carry 90

percent of the aircraft weight whereas the front nose tire carries only 10 percent of the aircraft

weight.

For tire sizing, we use the statistical method proposed by Raymer.

Raymer also says that if an airplane is to be operated under FAR 25 regulations, a 7 % margin must

be put to all calculated wheel loads. He also states that it is common to put a 25 % margin.

After Initial sizing 2, the aircraft weight was revised to 1300 kg.

7.2 Front Wheel

Wheel load acting on front wheel = 10% of total weight = 130 kg

Now considering FAR regulations and the design margin, we get the load of front wheel as 171.6 kg

D= A x WB

Where D= Diameter of the wheel in cm

W= Load acting on wheel in kg

A, B= Appropriate Constants from Raymer

So,

D (cm) = 5.1*(171.6)0.349 = 30.71 cm

Similarly,

W= A x WB

Where W= Width of the wheel in cm

W= Load acting on wheel in kg

A, B= Appropriate Constants from Raymer

So,

W (cm) = 2.3*(171.6)0.312 = 12 cm

7.3 Back Wheel

Wheel load acting on both back wheels = 90% of the total weight = 1170 kg

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Considering FAR regulations and the design margin, we get the load of front wheel as 1544.4 kg

Wheel load acting on one back wheel = 1544.4/2 = 772.2 kg

Now,

D= A x WB = 5.1*(772.2)0.349 = 52 cm

W= A x WB = 2.3*(772)0.312 = 18.3 cm

Parameter Diameter (cm) Width (cm)

Front wheel 30.71 12

Back Wheel 52 18.3

Figure 18: Wheel sizing Results

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8. CONFIGURATION AND LAYOUT

8.1 Pusher

The propeller can be placed at one of the 8 possible combinations of Tractor, Pusher and Fuselage,

Wing, Pod and tail. The configuration we have chosen is the fuselage pusher configuration because

of the following advantages:

It frees up the nose of the aircraft, which allows the payloads to be placed in the front part

of the fuselage. This aids in the efficient functioning of the payloads.

It reduces the skin friction drag because the pusher location allows the aircraft to fly in

undisturbed air.

It allows a reduction in the aircraft wetted area by shortening the fuselage.

This configuration also suffers from some disadvantages as given below:

The propeller has a reduced efficiency because it is forced to work with the disturbed airflow

from fuselage, wing and tails.

It requires a longer landing gear because the aft location causes the propeller to dip closer to

the runway as the nose is lifted for takeoff.

8.2 Propeller

The detailed design of the propeller, such as the blade shape, twist etc., are not required to layout a

propeller-engine aircraft. Using empirical data from Raymer, we get the propeller diameter as

D = 2.3 m

A constant-speed propeller is used to maximise the efficiency by changing the pitch angles so as to

maintain the engine at its optimal rpm.

8.3 Pratt & Whitney Canada PT6

The Pratt & Whitney Canada PT6 is one of the most popular turboprop aircraft engines in history and

are produced by Pratt & Whitney Canada. The PT6 family is particularly well known for its extremely

high reliability, with ‘Mean Time Between Outages’ on the order of 9000 hours in some models.

A number of advantages are derived from the design of the PT6A engine which has proven valuable

in routine field operation. They are discussed below.

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During an engine start, only the compressor section of the PT6A engine needs be rotated by the

starter-generator. By comparison, a fixed-shaft engine must spin all rotating components including

the reduction gearbox and propeller during an engine start, resulting in a requirement for heavier

starting systems.

The PT6A engine free turbine design allows the propeller RPM to be reduced and the propeller

feathered during ground operation without shutting down the engine. This facilitates permits very

quiet ground operation. Propeller RPM can also be varied in flight (on most applications) permitting

propeller RPM to be set for quieter cruise and optimum efficiency.

Figure 19: Pratt and Whitney Canada PT 6A

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9. Layout Configuration

Figure 20: Cut section view of UAV

The payloads and the control systems (shown in purple in the figure) are kept near the nose of the

aircraft. The auxiliary batteries (weighing more than 60 kg) are kept in the fuselage portion between

the wings (denoted in black in the figure). The fuel is stored in tanks inside the wings like any

standard aircraft, The tail does not contain any fuel tanks.

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10. Refined sizing:

In refined sizing, as the engine is already is already fixed and the power known, we can go for a fixed

engine analysis. The quantities Mmax, Vmax, Aspect ratio (AR), hp/WO, WO/S have all got well defined

values after we have gone through the process of constraint analysis.

Mmax= 0.678;

Vmax= 79.3736 kts;

Wo/S=169.2 kg/m2= 34.6548 lb/ft2

P/W= 0.11 hp/lb

AR=20.77

Using the refined sizing procedure given in Raymer, for a turboprop aircraft we have

Wf/Wo= 0.3049(1+0.05)=0.32014 We/Wo= 0.641

Wo_initial=1577.543 kg

Wo after refined sizing = 1547.889 kg

Therefore, after refined sizing, we have the take-off weight as 1547.889 kg.

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11. Material Selection and weight estimation

11.1 Fuselage

Choice of materials emphasizes not only strength/weight ratio but also:

• Comparably large strength allied to lightness;

• Strong stiffness and toughness for the rear rod;

• Low cost and weight for all parts.

• Fracture toughness

• Crack propagation rate

• Stress corrosion resistance

• Exfoliation corrosion resistance

Today, the main material used is aluminum alloys for all kinds of aircraft, which is pure aluminum

mixed with other metals to improve its strength.

Below is a comparison of material property comparison for different kinds of possible materials for

aircraft fuselage, aluminum sheet, wood, plastics, and carbon fibres.

Material Density(g/cm3) Tensile

Strength(at 73

for 22.7 oC)

Young’s

Modulus, E

(MPa)

Method of

Manufacturing

Price

Aluminum Sheet 2.7 30,000 Psi 70,000 Forging Moderate

Wood 0.8 550 Psi 10,000 Adhesive Bonding Cheap

Plastics(PVC) 1.15 7,000 Psi 3,000 Vacuum forming Very

Cheap

Carbon Fiber 1.78 100,000 Psi 50,000 Epoxy resin Very

Expensive

Composite ----- --------- ----- SCRIMPTM Moderate

Figure 21: Material Comparison

The fuselage material is chosen to be a mixture of carbon and quartz fibres blended in a composite

with Kevlar. Below this material, we place a wood laminate in layers. Between each layer of

laminate, a sturdy fabric is sandwiched in to provide insulation to internal components. The

composites are proprietary information and hence, their property data was unavailable. But, on

evaluating a number of other UAVs the fuselage is almost always made of composites. Hence, a

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research is required into the making of the composite material and thereafter, manufacturing a

fuselage out of it.

The edges of the wings are made of titanium and are dotted with microscopic holes that allow an

ethylene glycol solution to seep out of internal reservoirs and breakdown ice that forms on the

wings during flight. This is particularly important while flying at high altitudes like 19.8 km.

11.2 Wing skin

The upper and lower wing skins are subject to different loading conditions in flight. This influences

the choice of materials in each case. There are important differences in property requirements.

Upper Wing Skin Material must have

Compressive strength

Stiffness

Fatigue resistance

Fracture toughness

Compressive strength is the key property for upper wing skin.

Lower Wing Skin Material must have

Tensile strength

Fatigue resistance

Fracture toughness

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There are a number of age hardenable Al alloys that do exceed the minimum strength and fracture

toughness requirements. However, it is also critical that the lower wing skin is damage tolerant and

is able to resist failure by fatigue crack growth.

Figure 22: Al 2024 properties (Source ASM)

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Based on the above factors, we have decided to go with 2024 Aluminum alloy for the upper and

lower wing skin. It is an aluminium alloy having density of 2.78 g/cc, with copper as the primary

alloying element. It is used in applications requiring high strength to weight ratio, as well as good

fatigue resistance. To improve the corrosion resistance, which will be an important factor in coastal

conditions we will be coating the material with Zinc.

11.3 Weight estimation

The weight estimates can be done by following a scaling process using data from some other similar

aircrafts. Data of various UAVs like those of Predator, Heron Turboprop, Altair etc. were used.

Component Weight

Fuselage 124.96 kg/ 275.5 lbs

Payload 37 kg

Control system 19 kg

Wing 152.951 kg

Tail 52 kg

Fuel 567.22 kg

Engine 193 kg

Landing gear front 42.45 kg

Landing gear rear 84.186 kg

Auxiliary battery 60 kg

Propeller 4.789 kg

Figure 23: Component Weights

Fuselage weight has to be somewhere around 11% of the total weight as per the Cessna estimation

model given Raymer. However this analysis cannot be done as Cessna is a passenger aircraft and

here we have a pusher configured UAV.

The wing design is done by taking the ratio of empty weights of the UAVs under consideration, and

the tail weight is estimated by scaling down the weight of the wing to the size of the tail that was

estimated earlier in the tail sizing section.

The values of fuel fraction and GTOW were taken from the refined sizing results in order to estimate

the fuel weight and a reserve fuel of 5% was considered.

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The landing gear weights were also scaled from the empty weight values and the propeller weight

was estimated from historical data.

The payload weights including those of auxiliary batteries and control system have specific values as

the payloads are well defined under the payloads section of this report.

11.4 Determination of CG

Component Weight (kg) Distance from nose tip (m)

Moment (*9.8 Nm)

Fuselage 124.96 4.545 567.9432

Payload 37 1.2 44.4

Control system 19 1.2 22.8

Wing 152.951 -

Tail 52 7.19 373.88

Fuel 567.22 -

Engine 193 8.14 1571.02

Landing gear front 42.45 4.2 178.29

Landing gear rear 84.186 6.2 521.9532

Auxiliary battery 60 -

Propeller 4.789 9.09 43.53201

Net moment about nose tip 3323.81841

net weight of components whose weight have been considered for moment evaluation

557.385

C.G. location (from nose tip) 5.963m

Figure 7: CG Determination details

As evident from the above table, the Centre of gravity is located at a distance of 5.963 m from the

nose tip of the UAV. Please note that many of the above weights are estimates and there may be

slight inaccuracies in the arrived value.

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12. References

[1] Raymer, Daniel P. Aircraft Design: A Conceptual Approach Third Edition. Reston, VA: American

Institute of Aeronautics and Astronautics, Inc., 2006

[2] http://www.sandia.gov/RADAR/minisar.html

[3] http://www.l-3com.com/csw/Product/docs/31-Mini%20UAV%20Data%20Link%20-

MUDL%20gen2.pdf

[4] Airplane Design by Jan Roskam

[5] http://articles.timesofindia.indiatimes.com/2010-05-04/india/28296738_1_infiltration-indo-pak-

border-belt

[6] Unmanned Aerial Systems: Design, Development and Deployment by Reg Austin

[7] Aircraft Design Lecture notes by Prof. R.K Pant, IIT Bombay

[8] http://www.engbrasil.eng.br/index_arquivos/FAR23.pdf

http://www.sandia.gov/RADAR/minisar.html

http://www.l-3com.com/csw/Product/docs/31-Mini%20UAV%20Data%20Link%20-MUDL%20gen2.pdf

http://www.l-3com.com/csw/Product/docs/31-Mini%20UAV%20Data%20Link%20-MUDL%20gen2.pdf

http://articles.timesofindia.indiatimes.com/2010-05-04/india/28296738_1_infiltration-indo-pak-border-belt

http://articles.timesofindia.indiatimes.com/2010-05-04/india/28296738_1_infiltration-indo-pak-border-belt

http://www.engbrasil.eng.br/index_arquivos/FAR23.pdf

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Appendix 1

Following are the 3d diagrams of the UAV made using Google Sketchup 8. (All dimensions accurate)

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Appendix 2

MATLAB Codes used in constraint analysis

i=1;

Clmax=1.173;

for i=1:257

v(i)=(50+i)*5/18;

q=0.5*0.096*v(i)^2;

cd=0.008;

k=(1/pi*0.8*20.77);

ws(i)=q*sqrt(3*cd/k);

Clmax_to=0.9*((2.5*(16.458/30.034))+(1.5*(14.576/31.034)));

ws_stall(i)=0.5*1.23*18.005^2*Clmax;

ws3(i)=((600-137)/5)*1.827;

ws4(i)=q*0.9;

end

plot(v,ws)

hold on

plot(v,ws_stall)

plot(v,ws3)

plot(v,ws4)

%%

hold off

clto=0.8*Clmax_to;

for i=1:15;

pw(i)=0.01+i/100;

ws5(i)=210*clto*pw(i);

pw2(i)=0.237;

ws6(i)=34.6548;%fps

P a g e | 34


n=1.000001243;

ws7=36.5:((969-36.5)/69):969;

q_loiter=0.5*.096*62.5^2;

pw3(i)=(62.5/0.82)*(((q_loiter*cd)/ws7(i))+(ws7(i)*(n^2/(q_loiter*pi*20.77*0.8))));

end

hold on

plot(ws5,pw)

plot(ws5,pw2)

plot(ws6,pw)

plot(ws7,pw3)

P a g e | 35


Appendix 3

Design Comparisons with similar UAVs

MQ-1 Predator

GENERAL CHARACTERISTICS

Length 27 ft (8.22 m)

Wingspan 48.7 ft (14.8 m); MQ-1B, Block 10/15: 55.25 ft

(16.84 m)

Height 6.9 ft (2.1 m)

Wing Area 123.3 sq ft (11.5 m²)

Empty weight 1,130 lb (512 kg)

Loaded weight 2,250 lb (1,020 kg)

Max take-off weight 2,250 lb] (1,020 kg)

Power plant 1× Rotax 914F turbocharged four-cylinder

engine, 115 hp (86 kW)

PERFORMANCE

Maximum speed 135 mph (117 knots, 217 km/h)

Cruise speed 81–103 mph (70–90 knots, 130–165 km/h)

Stall speed 62 mph (54 knots, 100 km/h) (dependent on

aircraft weight)

Range >2,000 nmi (2,300 mi/3,700 km)

Endurance 24 hours

Service Ceiling 25,000 ft (7,620 m)

The General Atomics MQ-1 Predator is an unmanned aerial vehicle (UAV) used primarily by the

United States Air Force and Central Intelligence Agency. Initially conceived in the early 1990s for

reconnaissance and forward observation roles, the Predator carries cameras and other sensors but

has been modified and upgraded to carry and fire two AGM-114 Hellfire missiles or other

ammunitions. The aircraft, in use since 1995, has seen combat over Afghanistan, Pakistan, Bosnia,

Serbia, Iraq, and Yemen.

The USAF describes the Predator as a "Tier II" MALE UAS (medium-altitude, long-endurance UAV

system). The UAS consists of four aircraft or "air vehicles" with sensors, a ground control station

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(GCS), and a primary satellite link communication suite. Powered by a Rotax 912 UL engine and

driven by a propeller, the air vehicle can fly up to 400 nautical miles (740 km) to a target, loiter

overhead for 14 hours, and then return to its base.

Comparison

The predator has an inverted V tail when compared to our design and also has a longer fuselage with

fully retractable landing gears.

Figure 1: MQ-1 Predator

Lockheed Martin RQ-3 DarkStar

CHARECTERISTIC DATA

Length 4.57 m (15 ft 0 in)

Wingspan 21.03 m (69 ft 0 in)

Height 1.52 m (5 ft 0 in)

Weight max: 3900 kg (8600 lb)

Speed > 460 km/h (285 mph)

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Ceiling 19800 m (65000 ft)

Endurance 12 h

Propulsion Williams F129 turbofan; 8.45 kN (1900 lb)

The DarkStar UAV was developed as a LO-HAE (Low-Observable High-Altitude Endurance) UAV. The

ACTD (Advanced Concept Technology Demonstration) contract for this UAV was awarded to

Lockheed Martin in June 1994. A major subcontractor with a 50% share was Boeing, which was

responsible for wing development and production.

Figure 2: Lockheed Martin Darkstar

The RQ-3A was a LO flying-wing design with a very slightly forward-swept wing and a "flying saucer"-

shaped fuselage section. It was powered by a single Williams F129 (Model FJ44-1A) turbofan engine,

and could cruise for about 12 hours at an altitude of up to 19800 m (65000 ft). For the planned fully

autonomous missions, the DarkStar was equipped with a GPS/INS navigation system, which could be

changed in flight. Communication was done via two-way data links (command and control uplink,

sensor data downlink), either a wideband line-of-sight link or a J-band SATCOM link. The payload

P a g e | 38


bays in the lower fuselage could accommodate various types of sensors, but the primary options

were a Northrop Grumman AN/ZPQ-1 TESAR (Tactical Endurance Synthetic Aperture Radar)

surveillance radar or a Recon/Optical CA-236 electro-optical camera system.

Comparison

Compared to our UAV, the Darkstar uses swept forward wings with a wingspan of over 21 m and a

turbofan engine. The fuselage of the Darkstar is also very uniquely saucer shaped. Other than this,

the top speed of the Darkstar is 3 times that of our UAV.

Northrop Grumman RQ-4 Global Hawk

CHARECTERISTICS DATA

Length 13.53 m

Wingspan 35.42 m

Height 4.64 m

Weight 12130 kg

Speed 648 kmph

Ceiling 19800 m

Range 21720 km

Endurance 36 h

Propulsion Rolls-Royce/Allison F137-AD-100 turbofan; 33.8 kN (7600 lb)

The Global Hawk was the U.S. Air Force's first operational UAV in the HAE (High Altitude Endurance)

category. Its development began in 1994, when DARPA issued a request for proposals for their "Tier

II+" HAE UAV requirement. In March 1995 an ACTD (Advanced Concept Technology Demonstration)

contract was awarded to Teledyne Ryan Aeronautical (now part of Northrop Grumman), and in

January 1997, the designation RQ-4A was officially allocated to the Global Hawk UAV. The first of

five ACTD vehicles flew for the first time in February 1998.

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Figure 8: Northrop Grumman RQ-4 Global Hawk

The RQ-4A is powered by a Rolls-Royce/Allison F137-AD-100 (model AE 3007H) turbofan. It takes off

and lands on conventional runways using a retractable tricycle landing gear. The airframe has the

typical layout of a high-endurance UAV, and the prominent nose bulge houses the wideband

SATCOM antenna of 1.2 m (4 ft) diameter. The vehicle can reach an altitude of 19800 m (65000 ft)

and has a maximum endurance of at least 36 (and possibly up to 42) hours. A Global Hawk system

consists of two RQ-4A UAVs and two major ground stations, the RD-2A Mission Control Element

(MCE) and the RD-2B Launch and Recovery Element (LRE). The LRE is used to load autonomous flight

data into the UAV's GPS/INS navigation system, control the vehicle during take-off and landing, and

monitor its flight performance. The MCE personnel controls and monitors the UAV's sensor systems.

Both LRE and MCE can control three RQ-4As simultaneously. The main components of the RQ-4A's

ISS (Integrated Sensor Suite) for its surveillance, reconnaissance and target acquisition missions are

an SAR/MTI (Synthetic Aperture Radar/Moving Target Indicator) and IR/EO (Infrared/Electro-Optical)

sensors. For self-defence, the UAV is equipped with an AN/ALR-69 radar warning receiver and

AN/ALE-50 towed decoys.

Comparison

Compared to our UAV, the Global Hawk is a much larger UAV belonging to an entirely different

weight class and has three times the wingspan. The similarities between the two UAVs are the

operating ceiling and the endurance. We have selected a turboprop whereas the Global hawk flies a

Turbofan.

“All images used in Appendix 3 are taken from the websites of the manufacturers of the UAVs”

high altitude long endurance recon uav

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