uav design 25kg category

AURORA INTEGRATED SYSTEMS PVT. LTD.

Preliminary design Report MALE UAV IITK

Brief Introduction to first concept of design

Vineet singh

3/25/2014

1 Mission Requirement ...................................................................................................................... 2

2 Weight Estimation and Wing loading ............................................................................................. 3

2.1 Wing loading Diagrams ........................................................................................................... 4

2.1.1 Stall speed ....................................................................................................................... 4

2.1.1.1 Clean wing at diff altitude ........................................................................................... 4

2.1.1.2 Flaps deployed ............................................................................................................ 5

2.1.2 Cruise Speed .................................................................................................................... 6

3 Power Estimation ............................................................................................................................ 9

3.1 Engine Specifications3W‐28iCs ............................................................................................. 10

4 Aircraft final design including control surface sizing .................................................................... 12

4.1 Aircraft geometry .................................................................................................................. 12

4.2 Airfoil selection ..................................................................................................................... 13

4.2.1 Xfoil analysis of SD 7037 ............................................................................................... 13

4.3 Wing analysis ........................................................................................................................ 14

4.3.1 Stability Coefficients for two different Cg location including trim analysis .................. 15

4.4 Final design ........................................................................................................................... 17

5 Controls Architecture .................................................................................................................... 19

5.1.1.2 Limits ......................................................................................................................... 21

5.1.1.3 Lateral controller ....................................................................................................... 21

5.1.1.4 Longitudinal controller .............................................................................................. 21

1 Mission Requirement A medium altitude long endurance (MALE) UAV is required to be developed which should satisfy

following criteria:

1. Maximum take‐off Weight: < 25 kg

2. Payload capacity: > 3 kg with option of 5 kg in place of lesser fuel

3. Take off length: < 100 m on concrete runway with option for 50 m

4. Climb rate: > 2.5 m/s at MSL

5. Operational altitude: 1500 m Above Ground Level ( from take‐off point)

6. Endurance: > 4 hrs.

7. Range: > 50 km ( 100 km optional with heavier communication system)

8. Engine: 2 stroke gasoline based Internal combustion engine, should have option for heavy

fuel (JP5 or JP8) for increasing endurance

9. Maximum take‐off altitude: > 5500 m ~ 18000 ft.

10. Landing: run way landing with option for net arrested landing in future

Mission constraints:

1. The UAV should be able to disassemble and carried in a cases of size 5 ft. by 3 ft. by 2 ft.

2. The CG should not vary beyond 5 % with fuel

3. The endurance should be demonstrated at take‐off from Mean sea level, climb 1500 m and

go to range of 50 km and come back and land.

4. After full endurance mission, there should be reserve fuel of more than 1 ltr

5. The engine is limited to internal combustion gasoline based propeller driven

6. Landing gear is not retractable

7. The avionics power is to be supplied by on board generator but there is lithium polymer

battery to power the avionics in case for generator failure for up to 1 hour with total

avionics power being 50 Watts

8. Sufficient air scoops are to be provided for engine cooling as engine is assumed to be air

cooled.

9. The UAV should be able to perform the mission in up to 30 kts ( 56 kmph/~ 15 m/s)

continuous average wind.

10. For landing trailing edge flaps can be deployed to decrease the touchdown airspeed and no other high lift mechanisms are to be tried out.

2 Weight Estimation and Wing loading The weight estimation and wing loading go hand in hand and often require iterations to arrive at

required number. The structural weight is dependent on the wetted area and that is dependent on

the components and weight of the airframe. The key parameter is wing loading which decides the

design point and other performance parameters. Once an estimate of weight is made, the wing

loading can be used to predict area of the wing, which in turn is used to predict the structural

weight. This process is repeated till converged to a particular value.

The final weight including the weight of the components is shown below:

Table 2‐1

Components Wts (Kgs) Incremental Wt

Fuselage 3.5 3.5

Stru

ctures

Wing 2.500 6 Empennage 1.6 7.6 Rear Landing Gear 0 7.6

Nose Landing Gear 0 7.6 Camera 1.5 9.1

Payload/ AP Avionics 0.25 9.35 Video Storage 0.25 9.6 9.6

Battery 1.5 1.5

Po

wer S

ystem

Engine + Accessories 2.5 4 Fuel Tank 0.3 4.3 Fuel Weight 5.6 9.9 Starter 0 9.9 Alternator 0 9.9 Power Cicuitry 0.2 10.1

Pump 0 10.1 Propeller 0.15 10.25

19.85

Modem 0.07 0.07

Co

mm

un

icatio

n S

ystems

Antennae (Comm) 0.03 0.1 Antennae(Video) 0.03 0.13 Antennae(Iridium) 0 0.13

GPS (1) 0.03 0.16 wires 0.5 0.66

20.51

Fuselage Tailboom joint 0.05 0

Misc. S

tructu

res

Gimbal Mount 0.25 0.25 Wing Joint 0.05 0.3 Nose shock absorber 0.05 0.35 Rear shock Absorber 0.05 0.4

Motor Mount Bulkhea 0.05 0.45 Servo (Flap1 & 2) 0.1 0.55

Actua

tion

Servos(Aileron L &R) 0.1 0.65 Servo (Elevator) 0.05 0.7 Servo (Rudder) 0.05 0.75

Servo (Engine) 0.04 0.79

Video Transmitter 0.1 0.89 Unaccounted Wt Total TOGW 21.4 Total Empty Weight 15.8

2.1 Wing loading Diagrams

2.1.1 Stall speed

2.1.1.1 Clean wing at diff altitude

It is assumed that the clean wing has CLMAX = 1.0.

For this value of CLMAX, the wing loading required for different stall speeds an altitude is obtained.

Wing loading is targeted so as to have stall speed of around 17 m/s at MSL. This would mean that at

the time of of landing, the touchdown speed can be ~ 21 m/s at MSL. Higher landing speed would

mean higher loads on structures.

Moreover, the requirement of operations from high altitude at 5000m requires the stall speed to be

much higher. Hence, it should be noted that trailing edge flaps are required for landing and take off.

Structures, 7.92, 39%

Comm. Systems & Others, 0.87, 4%

Avionics, 0.12, 1%

Payload, 1.63, 8%

Propulsion system and power, 9.7, 48%

Weight Distribution

Structures Comm. Systems & Others Avionics Payload Propulsion system and power

2.1.1.2 Flaps deployed

CLMax = 1.4 using flaps deployed. Hence, the flaps should be sized such that CLMAX = 1.4 is achievable.

0

25

50

75

100

125

150

175

200

225

250

275

300

325

350

375

0 1000 2000 3000 4000 5000 6000 7000

W/S

(N

/m^

2)

Altitude (Meters)

W/S @ Diff. Stall Speeds & Altitude

W/S @ Vs = 17 m/sec. W/S @ Vs = 18 m/sec. W/S @ Vs = 19 m/sec.

W/S @ Vs = 20 m/sec. W/S @ Vs = 21 m/sec. W/S @ Vs = 22m/sec.

W/S @ Vs = 23 m/sec. W/S @ Vs = 24 m/sec. W/S Chosen

0

25

50

75

100

125

150

175

200

225

250

275

300

325

350

375

0 1000 2000 3000 4000 5000 6000 7000

W/S

(N

/m^

2)

Altitude (Meters)

W/S @ Diff. Stall Speeds & Altitude

W/S @ Vs = 17 m/sec. W/S @ Vs = 18 m/sec. W/S @ Vs = 19 m/sec.

W/S @ Vs = 20 m/sec. W/S @ Vs = 21 m/sec. W/S @ Vs = 22m/sec.

W/S @ Vs = 23 m/sec. W/S @ Vs = 24 m/sec. W/S Chosen

2.1.2 Cruise Speed For different cruise speed and aspect ratio, the optimum wing loading is plotted for different

altitude. For reference the selected wing loading is also shown for three different aspect ratios.

160180200220240260280300320340360380400420440460480500520540560580600620

0 1000 2000 3000 4000 5000 6000 7000

W/S

(N

/ sq

. m

eter

s)

Altitude (Meters)

Wing Loading Vs Altitude for Diff. Cruise Velocities for AR = 12

W/S: Vc = 20 W/S: Vc = 21 W/S: Vc = 22 W/S: Vc = 23 W/S: Vc = 24

W/S: Vc = 25 W/S: Vc = 26 W/S: Vc = 27 W/S: Vc = 28 W/S Chosen

100

120

140

160

180

200

220

240

260

280

300

320

340

360

380

400

420

0 1000 2000 3000 4000 5000 6000 7000

W/S

(N

/ sq

. m

eter

s)

Altitude (Meters)

Wing Loading Vs Altitude for Diff. Cruise Velocities for AR = 9 for best CLcruise



The worthwhile thing to note in the above graphs is that the most optimal cruise speed is close to

the stall speed as the most optimal Cruise lift coefficient is also very high.

Moreover, the optimal CLCruise becomes closer and closer to CLMAX as aspect ratio is increased. This

can be found in the case of Aspect ratio 12 and 15 compared to aspect ratio of 9 in the above

graphs.

As we have seen earlier, since the stall speed is close to 20 m/s for clean wing. The Cruise speed

should be at least 25 m/s so that any disturbances due to gust and instantaneous changes in the

airspeed do not lead to stall. It is therefore required to look at the carpet diagram of power loading

v/s wing loading for different aspect ratio. This should give an idea of the required thrust for cruise

condition as the aspect ratio is varied. This acts as a decision point for aspect ratio. The aspect ratio

should be chosen so as to be not very high as it also leads to structural difficulties. Neither should it

be so low so that it causes undue induced drag leading to decreased endurance.

180200220240260280300320340360380400420440460480500520540560580600620640660680700

0 1000 2000 3000 4000 5000 6000 7000

W/S

(N

/ sq

. m

eter

s)

Altitude (Meters)

Wing Loading Vs Altitude for Diff. Cruise Velocities for AR = 15



For Aspect ratio of 9, T/W required for cruise is 0.073. For aspect ratio of 12, T/W required is 0.063

and for 15, T/W is 0.056. As the aspect ratio is increased, although the best cruise speed is lower, the

Thrust by weight becomes lower. A thrust by weight of 0.06 should be good enough for the required

mission. Therefore, aspect ratio should be close to around 13.5.

One of the constraints of the mission is that the whole aircraft should be modular and carried in a

case and carried in cases of size 5ft by 3 ft by 2ft. This means the longest length which is wing section

should be less than 4 ft (giving half feet for the shock absorbing foam on each side).

It is intended that the wing be made in three pieces, whereby, each piece is less than 1.2 m.

Based on the constraint and the design insight gathered from the plots for cruise speed, the wing

span was chosen to be 3.3 m which can be broken down to 1.1 m three wing planform. This gives an

aspect ratio of 13.7 with chosen wing area of 0.8 m2.

0

0.05

0.1

195 210 225 240 255 270 285 300 315 330 345 360 375 390 405 420 435 450 465 480 495 510 525 540

T/W

W/S (Newton/Square Meter)

Power Loading Vs Wing Loading for Diff. AR at Different Cruise Speeds at best CLcruise

AR = 9 AR = 12 AR = 15 V = 20 : Se alevel

V = 21: Sea Level V = 22: Sea Level V = 23: Sea Level V = 24: Sea Level

V = 25: Sea Level V = 26: Sea Level W/S Chosen

3 Power Estimation The propulsion unit is required to be 2 stroke internal combustion gasoline based propeller driven

engine. The constraint of using gasoline based engine is to ensure that the fuel is readily available.

Also, two stroke engine produce higher power compared to four stroke engine and and are easily

available. To estimate the power required a table is made for estimating the climb rate for the given

weight of 20.5 kg and Cd0 assumed to 0.03 which is on the higher side. An important factor is the

propeller efficiency. A basic propeller curve shows that propeller efficiency is dependent on the

incoming air speed, rotational speed of the propeller and diameter of the propeller. Efficiency, as

high as 0.76, is predicted by blade element theory. However, as the worst case scenario, the

maximum propeller efficiency is assumed to 0.66. For achieving climbr ate of 4.5 m/s, the power

required at the engine shaft is back calculated using the propeller efficiency for different aspect

ratio.

The power input above is taken to be 2647 watts which is 3.55 horsepower.

This is available from a 28 cc 3W engine. 3W engines are the most widely used highly reliable engine

from Germany. The power required shows that even for worst case scenario of lesser propeller

efficiency a climb rate of 4.5 m/s is easily achievable. In fact, climb rate as high as 6 m/s can be

achieved. This excess power is especially valuable for operating at higher altitude where the cruise

speed, take off speed as well as the landing speed increases to about 1.5 times.

0

500

1000

1500

2000

2500

3000

3500

4000

4500

20 24 28 32 36 40

Po

wer

(W

atts

)

Velocity (m/sec.)

Power Requirements for Target Climb Rate 4.5 m/sec. @ 1-g load, at W = 20.5 Kg

Pout Req.: 20.5 Kg, AR = 9 Pin Req.: 20.5 Kg, AR = 9 Pout Req.: 20.5 Kg, AR = 12

Pin Req.:20.5 Kg, AR = 12 Pout Req.: 20.5 Kg, AR = 15 Pin Req.: 20.5 Kg, AR = 15

Power Input: 2647 Watts

It is recommended that the next higher engine 3W 55i CS should also be kept as back up for theflight

trials in case the weight due to payload becomes more.

3.1 Engine Specifications3W‐28iCs

Techn. specifications

Cylinder capacity 1.74 cu in

Power 3.55 HP

Power Rating 2.64 KW

Speed range 1500 ‐ 8500 rpm

Oil / Gasoline Ratio 1 : 50 / 2% Mix

IIS ‐ Ignition 6 ‐ 8,4 V

Propeller 2‐bladed 16x10;18x8;18x10;20x8

Propeller 3‐bladed 16x8;16x10

3W‐55 XiCS

Techn. specifications

Power 5.9 HP

Power Rating 4.399 KW

Speed range 1300 ‐ 8500 rpm

Oil / Gasoline Ratio 1:50 / 2% Mix

IIS ‐ Ignition 6 ‐ 8,4 V

Propeller 2‐bladed 22x8; 22x10; 22x12; 24x8

Propeller 3‐bladed 19x12; 20x8; 20x10; 22x8

Apart from the engine itself, there are other accessories like intake manifold, exhaust manifold and

muffler and throttle actuator required to run the engine.

Different propellers can be mounted to vary the final output power. Tests need to be carried out to

validate the best propeller that should be used which provides the best fuel efficiency and sufficient

static thrust to have short take off.

4 Aircraft final design including control surface sizing

4.1 Aircraft geometry

The wing is divided into three sections with each section being 1100 mm long. The wing shape is

taken to be close to semi‐elliptical to have maximum aerodynamic efficiency. The downside is that

such wing requires precise manufacturing of the mould.

Two design variants are sought, one using inverted V‐Tail and the other inverted U tail. The inverted

V‐tail has the advantage of saving weight and having proverse roll‐yaw coupling . However, it also

requires precise fabrication of the molds for alignment and also leads to coupling in elevator and

rudder if the control surface movement is not symmetric.

4.2 Airfoil selection Many different airfoil were studied for selection. Among the low Reynolds number airfoil, sd7037

was found to be the best suited. It should also be known that sd7037 has been widely used on many

UAVs and also on different models and tested in wind tunnel by different labs extensively.

4.2.1 Xfoil analysis of SD 7037

Green colour in following plot is for Reynolds number 200,000. Magenta is for Reynolds number

500,000

4.3 Wing analysis

Wing span : 3.3m

Wing Area : 0.79m

MAC : 0.24m

Root Chord Length : 0.30m

Tail span : 1m

Fuselage length : 1.2m

Tail arm length (from leading edge of wing root to tail root) : 1m

Vertical tail height : 0.5m

Wing Airfoil : SD 7037

Tail airfoil : NACA 0012

4.3.1 Stability Coefficients for two different Cg location including trim analysis

Neutral point : 0.235m from leading edge of root chord

Cruise velocity : 24m/s

Takeoff weight : 21kg

Cruise CL : 0.7

Case 1 : CG at 0.15m from leading edge of root chord

Case 2 : CG at 0.19m from leading edge of root chord

Case 1 Case 2

Other important coefficients are

CL_alfa : 6.2

Cl_aileron: 0.45

Cn_aileron : 0.045

Cl_rudder : 0.0045

Cn_rudder : 0.2

CL_flap : 0.91

CM_flap : 0.98

Cl_beta : ‐0.035

Cl_p : ‐0.6

Cl_r : 0.2

Cn_r : ‐0.2

Cn_p : ‐0.12

4.4 Final design It is recommended that CG be kept at 190 mm from the leading edge of the wing (case 2 above). This

leads to zero trim elevator deflection at around 4 degree angle of attack which corresponds to CL=

0.7, corresponding to cruise speed of 24 m/s at MSL.

The choice of CG is important as this leads to the trim elevator which should be close to zero to have

minimum trim drag.

The wing is not given any dihedral and spiral is left slightly unstable. A measurement of spiral

stability is the ratio of Cl_beta*Cn_r and Cn_beta*Cl_r. If this is less than 1, spiral is unstable and if it

is more than one spiral is stable. Typically, spiral is left unstable for the aircrafts as it is slow mode

and the pilot can intuitively correct for slight disturbance.

Dutch rill on the other hand is a competing mode. This mode tends to become unstable if spiral is

made stable and vice‐versa. The dutch roll is dependent on the vertical tail volume and the dihedral

angle of the wing. In the current design, the vertical tail volume is kept high so as to have stable

dutch roll. A stable dutch roll is required, especially, during landing at low speeds.

Flaps and Aileron are sized so as to have control surface derivatives closely equal to the general

acrobatic aero models. This can lead to a bit of sensitive controls for manual mode.

5 Controls Architecture AVES autopilot is intended to be used for final integration. AVES autopilot is provides generic PId

control architecture with different gains and limits which can be configured for the aircraft. The

gains are required to be tuned for the particular aircraft to achieve the best performance. The AVES

Ground Control Station also supports the gain tuning interface which provides the capability to give

step commands to desirec variable and see the response and accordingly change the gains of the

selected control loop.

The controls are divided mainly in lateral controls which can be configured for aileron and rudder/

aileron only/ rudder only models and the two longitudinal controls elevator and throttle. User

should note than no separate control structure is required for v‐tail or inverted v‐tail or a model with

elevons. That has to be done in servo configuration explained later on in the manual.

Each control loop has UAV parameters, switches and the PID blocks. The PID blocks are not strictly

simple traditional PID logic. In various places the control uses aircraft geometry and aerodynamic

parameters inside the PID blocks.

Please go through the gain tuning manual for understanding the methodlogy of tuning the control

gains and the aircraft parameters which also includes trims and limits.

5.1.1.1.1.1.1 Lateral control: Aileron/Rudder

Figure 1: Aileron/Rudder Control loop

Aileron control is meant for the bank angle control of the UAV. The commanded roll can come

either from heading to roll block which is enabled in navigation phase of autopilot or directly from

the loiter phase. If it is rudder only model, the control loop adds the yaw rate feedback to aileron

output and thus, rudder servo should be connected to the aileron output of autopilot. If it is aileron

only model, the rudder should be left disconnected.

5.1.1.1.1.1.2 Longitudinal control: elevator

Figure 2: Elevator control loop

Elevator can be controlled through altitude or air speed. Depending on the altitude window

configured in aircraft trims in configurator, the control is either in altitude to pitch or air speed to

pitch. It should be understood that both air speed and altitude cannot command desired pitch

simultaneously. The desired pitch is controlled through pitch to elevator PID block.

5.1.1.1.1.1.3 Longitudinal control: throttle

Figure 3: Throttle control loop

Throttle can be controlled through altitude or air speed. Depending on the altitude window

configured in aircraft trims in configurator, the control is either in altitude to throttle or air speed to

throttle. It should be understood that both air speed and altitude cannot command desired throttle

simultaneously. The desired throttle is passed through throttle slew rate to give final throttle.

5.1.1.2 Limits

Each control inner or outer control loop in Aves control structure has saturation limits to prevent

overmovement of the servos or superfluous commands to be generated. These limits are very

important for the controller and the controller is designed so as to not exceed any of the limits. The

default values of these limits should be used. If improvement in performance is desired, some of the

relevant limits can be altered.

5.1.1.3 Lateral controller Table 2: Lateral Controller Limits

S. no. Parameter Meaning

1 Max Cmd Roll Maximum Commanded roll angle from heading to roll controller

2 Min Cmd Roll Minimum Commanded roll angle from heading to roll controller, should be negative

3 Max Roll Error Maximum commanded roll rate from roll angle hold controller

4 Min Roll Error Minimum commanded roll rate from roll angle hold controller. Should be negative

5 Max Aileron Maximum commanded aileron from roll controller

6 Min Aileron Minimum commanded aileron from roll controller

7 Pitch saturation Value of tangent of pitch veyond qhich pitch effects should not be taken in the roll controller. Should be kept at default value.

8 Min Pdyn Minimum dynamic pressure that can be achieved by the aircraft. This value should be set to half the value of air density (1.2125 kg/m^3) times the square of minimum air speed.

9 Max Rudder Maximum rudder commanded from yaw rate damping.

10 Min Rudder Minimum rudder commanded form yaw rate damping, should be negative.

5.1.1.4 Longitudinal controller Table 3: Longitudinal controller Limits

S. no. Parameter Meaning

1 Max Cmd Pitch Maximum Commanded pitch from altitude or air speed.

2 Min Cmd Pitch Minimum commanded pitch form altitude or air speed, should be negative.

3 Max elevator Maximum elevator commanded through pitch control.

4 Min elevator Minimum elevator commanded through pitch control.

5 Max throttle Maximum commanded throttle form throttle control.

6 Min throttle Minimum commanded throttle form throttle control.

uav design 25kg category

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