HADAFTM1404 Aircraft Design Book - fumblog.um.ac.irfumblog.um.ac.ir/gallery/869/HADAF CONCEPTUAL_20110620.pdf · 7.4 DISPOSITION OF LANDING GEAR AND STRUT 169 ... The students who

HADAFTM

1404 Aircraft Design Book

TABLE OF CONTESNTS

ABOUT THE GROUP 5

LIST OF SYMBOLS 6

1 WEIGHT SIZING 8

1.1 INTRODUCTION 8

1.2 MISSON SPECIFICATION 10

1.2.1 DETERMINATION OF MISSION PAYLOAD WEIGHT 11

1.2.2 DETERMINATION OF MISSION FUEL RESERVES 12

1.3 DATA ANALYSIS 13

1.4 WEIGHT SIZING 15

1.5 SENSETIVITY ANALYSIS 19

1.5.1 SENSITIVITY OF TAKEOFF WEIGHT TO PAYLOAD WEIGHT 20

1.5.2 SENSITIVITY OF TAKEOFF WEIGHT TO EMPTY WEIGHT 21

1.5.3 SENSITIVITY OF TAKEOFF WEIGHT TO RANGE, ENDURANCE, SPEED, SPECIFIC FUEL CONSUMPTION,

PROPELLER EFFICIENCY AND LIFT-TO-DRAG RATIO 21

1.6 APPENDIX: 24

1.7 REFERENCES 28

2 PERFORMANCE ESTIMATION 29

2.1 INTRODUCTION 29

2.2 SIZING TO STALL SPEED REQUIREMENTS 29

2.3 SIZING TO TAKE-OFF DISTANCE REQUIREMENTS 31

2.4 . SIZING TO LANDING DISTANCE REQUIREMENT 35

2.5 SIZING TO CLIMB REQUIREMENT 38

2.6 SIZING TO CRUISE SPEED REQUIREMENT 46

2.7 MATCHING OF ALL SIZING REQUIREMENT 50

2.8 ROAD MAP 53

2.9 APPENDIX 54

2.10 REFERENCES 58

3 SELECTION OF ENGINE 59

3.1 INTRODUCTION 59

3.2 SELECTION OF THE PROPULSION SYSTEM TYPE 59

3.3 DETERMINATION OF THE NUMBER OF ENGINES 62

3.4 DISPOSITION OF ENGINE 63

3.5 ENGINE BRANDS 64

3.6 PROPELLER DESIGN 75

3.7 DESIGN CHART (ABSTRACT): 77

PROPELLER DESIGN 77

3.8 REFERENCES 78

HADAFTM


4 THE GENERAL ARRANGEMENT AND FUSELAGE DESIGN 79

4.1 INTRODUCTION 79

4.2 OUTLINE OF CONFIGURATION POSSIBILITIES 79

4.2.1 OVERALL CONFIGURATION 80

4.2.2 ENGINE TYPE AND DISPOSITION 81

4.2.3 WING CONFIGURATION 82

4.2.4 EMPENNAGE CONFIGURATION 85

4.2.5 1.5. LANDING GEAR TYPE AND DISPOSITION 87

4.2.6 DETERMINATION OF THE CENTER OF VISION (COV) 90

4.3 OUTLINE OF FUSELAGE DESIGN 92

4.3.1 CROSS-SECTION DESIGN 93

4.3.2 FUSELAGE DIAMETER 93

4.3.3 THE SHEET-METAL TAIL CONE SECTION 94

4.3.4 FUSELAGE SHAPE 95

4.3.5 HADAF CONFIGURATION 97

4.4 DESIGNING DIAGRAM 98

4.5 APPENDIX 100

4.6 REFERENCES: 108

5 WING SIZING 109

5.1 INTRODUCTION 109

5.1.1 DECIDE 1DECIDE ON THE OVERAL WING/FUSELAGE ARRANGMENT 109

5.2 MORE DETAIL DESIGN PARAMETER 109

5.3 AIRFOIL PROFILE DESIGN 111

5.4 WING PLANFORM DESIGN 115

5.4.1 SWEEPANGLE 116

5.4.2 THICKNESS RATIO (T/C) 117

5.4.3 TAPER RATIO 118

5.4.4 TWIST ANGLE 119

5.4.5 INCIDENT ANGLE 119

5.4.6 DIHEDRAL ANGLE 120

5.4.7 WING TEST: 120

5.4.8 LATERAL CONTROL SURFACES 123

5.4.9 VERIFYING CLEAN AIRPLANE MAXIMUM LIFT COEFFICIENT AND SIZING THE HIGH

LIFT DEVICES 125

5.5 DECIDE ON THE OVERALL STRUCTURAL WING CONFIGURATION 129

5.6 COMPUTE THE WING FUEL VOLUME 130

5.7 ROAD MAP 131

5.8 REFERENCES: 132

6 PRELIMINARY TAIL SIZING 133


6.2 EMPENNAGE FUNCTIONS 134

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6.2.1 PITCH 135

6.2.2 YAW 136

6.2.3 ROLL 136

6.3 EMPENNAGE SIZING 137

6.3.1 EMPENNAGE CONFIGURATION 137

6.3.2 EMPENNAGE DISPOSITION 143

6.3.3 EMPENNAGE SIZE 143

6.3.4 FINAL CALCULATIONS 144

6.4 PLANFORM GEOMETRY OF EMPENNAGE 147

6.4.1 ASPECT RATIO 147

6.4.2 SWEEP ANGLE 149

6.4.3 TAPER RATIO 150

6.4.4 THICKNESS RATIO 150

6.4.5 DIHEDRAL ANGLE 151

6.4.6 INCIDENCE ANGLE 152

6.4.7 AIRFOIL SHAPE 152

6.5 CONTROL SURFACES SIZING 153

6.5.1 ELEVATOR 153

6.5.2 RUDDER 155

6.5.3 SIZE OF ELEVATOR AND RUDDER 156

6.6 REFERENCES 161

7 LANDING GEAR 162


7.2 FIXED / RETRACTABLE LANDING GEAR 164

7.3 LANDING GEAR CONFIGURATION TYPES 165

7.3.1 TAIL-WHEEL(TAIL-DRAGGER)[2] 165

7.3.2 NOSE WHEEL (TRICYCLE) 166

7.3.3 TANDEM 168

7.4 DISPOSITION OF LANDING GEAR AND STRUT 169

7.4.1 TIP-OVER CRITERIA: 169

7.4.2 GROUND CLEARANCE CRITERIA:[2] 170

7.5 COMPUTING THE MAXIMUM STATIC LOAD[4] 173

7.6 SELECTION OF TIRES[2] 174

7.7 LANDING GEAR DATABASE: 176

7.8 FINAL DRAWING OF LANDING GEAR SYSTEM 179

7.9 TABLE OF FINAL RESULTS 181

7.10 ROAD MAP 182

7.11 REFERENDES 183

8 WEIGHT AND BALANCE ANALYSIS 184


8.2 COMPONENT WEIGHT BREAKDOWN 186

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8.3 PRELIMINARY ARRANGEMENT DRAWING OF AIRPLANE AND EACH COMPONENT C.G LOCATION 186

8.4 CATEGORIZING THE X, Y, Z COORDINATE OF C.G OF EACH COMPONENT 188

8.4.1 FUSELAGE GROUP 188

8.4.2 WING GROUP 189

8.4.3 EMPENNAGE GROUP 189

8.4.4 ENGINE GROUP 189

8.4.5 LANDING GEAR GROUP 189

8.4.6 FIXED EQUIPMENTS GROUP 190

8.4.7 FUEL GROUP 191

8.4.8 PASSENGERS GROUP 191

8.5 CALCULATING THE XC.G&YC.GOF AIRPLANE 192

8.6 WEIGHT C.G EXCURSION DIAGRAM 193

8.7 C.G EXCURSION DIAGRAM ARGUMENT 195

8.8 REFRENCES 196

HADAFTM


ABOUT THE GROUP

A group of mechanical engineering students of Ferdowsi University of Mashhad

established the airplane-designing group of HADAF, on July 2009. Under the

instruction of Mr Mohammad JavadDarabiMahboub, the team started the conceptual

design phase of a2-seatedultra-light airplane called HADAFTM1

1404.

The students who attended in this project are:

1. Mojtaba Balaj

2. Mehdi BehnamVashani

3. Abbas Daliry

4. Amir Faghihi

5. Sina Heidari

6. Ali Mehrkish

7. Seyyed Mohammad Naghavizadeh

8. Hassan Nami

9. SomayyeNorouzi

10. Ali Omidi

11. Amir Kimiagaran

12. Mohsen Shamsabadi

13. Seyyed Ali Sahhaf

14. Saeid Zare

15. Saman Zare

HADAFTM

1404 Aircraft Design Book List of Symbols

List of Symbols

Performance Estimation

Wing area S

Take-off trust TTO

Take-off power PTO

Maximum required take-off lift coefficient

with flaps up

CL ,max (clean)

Maximum required lift coefficient for take-off CL ,max TO

Maximum required lift coefficient for landing CL ,max L , CL ,max PA

Wing loading W/S

Thrust loading, T/W

power-off stall speed

Density

Aerodynamic drag coefficient CD

ground friction coefficient µG

take-off ground roll STOG

take-off distance

Landing weight WL

Approach speed VA

landing ground run SLG

aspect ratio A

Oswald e

Weight Sizing

Take off gross weight WTO

Empty weight WE

Mission fuel weight WF

Operating empty weight WOE

Payload weight WPL

Trapped fuel & oil weight Wtfo

Crew weight Wcrew

Manufacturer empty weight WME

Fixed equipment weight WFEQ

Range R

Endurance of loiter Eltr

Cruise Velocity Vcr

Fuel Reserve weight

Propeller efficiency ηp

Lift-to-drag ratio L/D

Specific fuel consumption CP

HADAFTM

1404 Aircraft Design Book List of Symbols

zero-lift coefficient

equivalent area

wetted area Swet

power index

Selection of Engine

Mach number

propeller diameter

blade power loading Pb

Wing Sizing

Size S

Aspect ratio A

Sweep angle

Thickness ratio t/c

Taper ratio

Incident angle

Dihedral angle

Preliminary Tail Sizing

static loading factor ns

total gear weight

HADAFTM

1404 Aircraft Design Book Weight Sizing

1 Weight Sizing

1.1 INTRODUCTION

Airplanes normally meet very stringent range, endurance, speed and cruise speed

objectives while carrying a given payload. It is important to predict the minimum

airplane weight and the weight of fuel which is needed to accomplish a given mission.

This report focuses on the processes of Mission specification, weight sizing &

sensitivity analysis.

Figure 1-1the preliminary sizing process as covered in this report

Having the mission specification of our ultra-light 2-seated aircraft in mind, in this

report we‟ll give an estimation of:

- Take off gross weight, WTO

- Empty weight, WE

- Mission fuel weight, WF

Breaking down the takeoff gross weight we have the following formulation:

WTO= WOE+ WF+ WPL (1.1)

WOE = WE + Wtfo + Wcrew (1.2)

Preliminary Sizing

WTO WE WF

Sensitivity Analysis

Definition of R&D Needs

Refinement of Preliminary

Sizing

MISSION

SPECIFICATION

HADAFTM


WE = WME + WFEQ (1.3)

Where:

WTO= Take off gross weight

WOE = Operating empty weight

WF = Fuel weight

WPL = Payload weight

WE = Empty weight

Wtfo = Trapped fuel & oil weight

Wcrew = Crew weight

WME = Manufacturer empty weight

WFEQ = Fixed equipment weight

At this Junction, two key points must be made:

Point1: It is not difficult to estimate the required mission fuel weight WF from

very basic considerations.

Point2: According to Roskam Method, there exists a linear relationship

between log10WTO and log10WE for homebuilt airplanes.

Based on these two points, the process of estimating values for WTO, WE and WF

consists of the following steps:

step1. The mission payload weight, WPL will be determined.

step2. A likely value of take-off weight, WTO will be guessed.

step3. The mission fuel weight, WF will be determined.

step4. A tentative value for WOE will be calculated from:

(1.4)

step5. A tentative value for WE will be calculated from:

(1.5)

Although Wtfo often gets neglected for some airplanes, in this report it is assumed

to amount as much as 0.5% of WTO at this stage in the design process.

step6. The allowable value of WE will be found.

step7. The values for and for WE, as obtained from steps 5 and 6, will be

compared. Next, an adjustment to the value of will be made and steps 3

through 6 will be repeated. This process continues until the values of and

agree each other to within some pre-selected tolerance. A tolerance of 0.5%

is usually sufficient at this stage in the design process.

HADAFTM


Figure 1-2roadmap of weight sizing process done by HADAFTM group

After estimating takeoff gross weight and aircraft empty weight is completely done by

using Breguit equations, some coefficients called growth factors will be calculated.

This part of weight estimation process will be fully discussed in the last part of this

report.

Weight estimating process is actually the most important part of plane designing

process, because all upcoming calculations in the other parts will be taken into

account based on information gained in this part. So, this part must be done with

much more efforts and strict rational reasoning.

1.2 MISSON SPECIFICATION

In order to define a mission for the goal plane, different aviation regulations must be

considered such as FAR and JAR and mix the information gained this way with our

especial needs and create a mission profile. Correctness of this profile is so important

because any mistake in this step, may accuse every assumption that has been made

earlier, and therefore all other design processes would be incorrect. Not having a

correct and fit view to flight mission profile, causes the designer to be confused in

gathering data for the database too. So second relation between WTO and WE will

HADAFTM


become incorrect. Like all traditional designs, a mission profile must be drawn which

gives all its specifications here.

Figure 1-3 mission profile of HADAF1404 ultra-light 2-seated aircraft

Destination of HADAF 1404

flight, taking off from Mashhad, is considered to be Tehran

and it is known that distance between Mashhad and Tehran is about 924 km so:

R = 926km = 500 nm

An endurance of about 1 hour during loiter phase near the destination is required so:

Endurance = Eltr = 1 hour

Also flying 115 mph during the cruise phase is desirable. So:

Velocity = Vcr = 115mph =185 km/h

This data will be used in the rest of weight estimation process. Now data analysis

based on the mission profile will be started.

1.2.1 DETERMINATION OF MISSION PAYLOAD WEIGHT

Mission payload weight, WPL, is normally specified in the mission specification. This

payload weight usually consists of one or more of the following:

1. Passengers and baggage

2. Cargo

For passengers in a commercial airplane an average weight of 175 lbs. per person

and 30 lbs. of baggage is a reasonable assumption for short to medium distance

flights. As FAR23 certified the airplanes of the homebuilt class, they are usually

operated by Owner/Pilots and it is unusual to define the crew weight as part of the

payload in these cases, as the pilot weight is considered as payload in this project.

As defined in the mission specification of HADAF1404, there are two passengers.

Each passenger weighs 80kg (176lbs) carrying a baggage of 20kg (44lbs). Another

HADAFTM


additional 30kg cargo is taken into account. So the total mission payload will be

230kg (507lbs). This payload weight presents a high payload weight for this type of

ultra-light airplanes.

This additional cargo is considered to meet the target applications of HADAF1404.

As a family airplane HADAF1404 can carry a child up to 25Kg (and 5kg for baby-

chair). For Urban, rescue, meteorological, or forestry purposes, this additional cargo

is considered for extra equipment carried by airplane.

1.2.2 DETERMINATION OF MISSION FUEL RESERVES

Fuel reserves are normally specified in the mission specification. They are also

specified in those FAR regulations. Due to Roskam method, fuel reserves are

generally specified in one or more of the following types:

1. As a fraction of WF,used.

2. As a requirement for additional range so that an alternate airport can be reached

3. As a requirement for loiter time

Since a long loiter time has been assumed in mission specifications of HADAF1404,

no additional fuel reserve was held in considerations. So:

Table 1-1 mission specification for HADAFTM1404

Airplane code: HADAF1404

Airplane type: Homebuilt airplane

Payload: Two passengers at 80kg each (includes pilot), 40kg

total baggage and 30kg additional weight

Range: 926km (500 nm) with maximum payload

(No reserved fuel is considered.)

Endurance: 1hour loiter

Altitude: 12,000 ft. (for the design range)

Certification base: FAR23

HADAFTM


1.3 DATA ANALYSIS

Data gathering is the next stage. This stage is very important and simultaneously too

time consuming. Gathering the data started by denoting a range for takeoff gross

weight. Having flight mission and this range in mind, team members created a

database which consists of 150 ultra-light aircraft. The initial database included 2-

seated, 3-seated and low weight 4-seated aircrafts. As soon as this database

completed, a number of items were omitted based on some other factors. These factors

are as listed below:

i. Material: Since it was decided to build the aircraft with composite materials

before the start of the design process, all metallic or wooden aircrafts were not

applicable as the entries of the database. Therefore some of these aircrafts

were omitted from the database.

ii. Range: As it is assumed, the range of flight to be about 500 nm, the planes

that their ranges were out of 450 – 600 nm range were omitted from the

database.

iii. Type of plane: Some planes in the database have irrelevant applications. So,

it doesn‟t make any sense to put these planes data in the database.

iv. Lack of data: data sets of some planes were incomplete and despite of many

searches their missing data could not be found.

v. Similarity of some data: For some pairs of planes data, there is a close

similarity. Therefore, one of them should be omitted. Because similarity of data

leads to error when plotting WTO vs. WE diagram and of course leads to gain

incorrect coefficients.

The database, purified from incorrect data and fully coincident to the flight mission,

was prepared as follows.

Table 1-2 Final database of ultra-light airplanes matched to the specified mission

Max Gross

Wt(kg) log wto

Standard

Empty Wt(kg) log we

Skylark 599 2.777426822 296 2.471291711

Pioneer 200 472 2.673941999 260 2.414973348

F99 Rambo 470 2.672097858 285 2.45484486

SportCruiser 1 599 2.777426822 306 2.485721426

CT2K 480 2.681241237 258 2.411619706

Remos 598 2.776701184 303 2.481442629

Jabiru j-170 545 2.736396502 290 2.462397998

TL 3000 Sirius 472 2.673941999 295 2.469822016

T-10 Frigate 550 2.740362689 315 2.498310554

Sport 600 599 2.777426822 295 2.469822016

P2004 Bravo 599 2.777426822 331 2.519827994

TECNAM P92 600 2.77815125 325 2.511883361

F99 LSA 594 2.773786445 308 2.488550717

HADAFTM


In section 2.1 of Part1 of Roskam method, point 2 raised the issue of the existence of a

linear relationship between log10WE and log10WTO. Once such relationship is

established, it should be easy to obtain WE from WTO.

It is desirable as small as value for WE for any given WTO. Therefore, it is reasonable

to assume, that a manufacturer will always try to make WE as small as possible for

any given takeoff weight.

For that reason, at any value of WTO in table1-2, the corresponding value of WE

should be viewed as the 'minimum allowable' value at the current 'state-of-the-art' of

airplane design.

The trend of log WTO vs. log WE is plotted and calculation of the coefficients, A and B,

is performed.

Log10 (WTO) = A + B.Log10 (WE) (1.6)

Our based-on-database plot of Log10(WTO)vs.Log10(WE) and the calculated

coefficients of the upcoming equation are:

Log10(WTO) = 1.069 Log10(WE) + 0.096 (1.7)

So A= 0.096 & B= 1.069

Figure 1-4 based-on-database plot of Log10 (WTO) vs. Log10 (WE)

y = 1.069x + 0.096 R² = 0.539

2.66

2.68

2.7

2.72

2.74

2.76

2.78

2.8

2.4 2.42 2.44 2.46 2.48 2.5 2.52 2.54

HADAFTM


1.4 WEIGHT SIZING

Illustrated in figure1-2, stages of this part were explained in the previous sections.

Now it‟s the time to present the calculated data for each weight element. The purpose

of this section is the detailed calculations of each term in takeoff gross weight, one by

one, as shown in the introduction.

- Payload weight

Payload weight is assumed to be equaled to 230 kg

- Trapped fuel & oil

Wtfo = 0.005 WTO (1.8)

- Reserved Fuel

Wres = 0 lit

- Fuel Fractions

According to table 2.1 of Roskam book fuel fractions suggested for a homebuilt

aircraft is as listed below:

Table 1-3 table of fuel fractions except cruise and loiter phase

Phase

Engine

start,

warm-up

Taxi Take-

off Climb Descent

Landing,

Taxi

Shut

Down

Homebuilt 0.998 0.998 0.998 0.995 0.995 0.995

- Breguit equations

For cruise and loiter phases, fuel fractions cannot be chosen from such a table,

because for these phases, fuel fractions depend on factors L/D, CP, R, P , V and E. So

for different cases different values for fuel fractions is expectable. For the goal plane,

values as listed in tables1-3 and 1-4 for cruise and loiter phases are assumed.

Fuel fraction of the phase of climb was calculated by Breguit equations, too. But the

fuel fractions presented by statistical information of Roskam were used.

i. Cruise phase:

According to table 2.2 of Roskam book, the suggested value for L/D is 8 to 10. But it is

clear that airplanes with smooth exteriors and/or high wing loadings can have L/D

HADAFTM


values which are considerably higher. So a good estimation for L/D can be made by

using the drag polar of common airfoil used for homebuilt or single engine aircrafts.

Table 1-4 Suggested values for L/D,Cj,ηp and Cp for cruise and loiter phase(table 2-2 in Roskam method)

The numbers in this table represent values based on existing engines. So the specific

fuel consumption is calculated according to the available engines specification such

as Rotax and Jabiru engines as below:

(1.9)

Table 1-5 suggested values for L/D ,CP& according to Roskam method for cruise

L/D CP(lbs/hp/hr) Rcr(nm) P

Cruise 13 0.39 500 0.8

1

ln375)(

i

i

crcrP

Pcr

W

W

D

L

CsmR

(1.10)

Using information of table1-5 and Eq. 1.10, results:

4

5

W

W=0.9441

ii. Loiter phase:

Using the assumptions listed below and Breguit formula for Eltr, calculation of fuel

fraction for this phase can be performed.

HADAFTM


Cp in loiter phase will be increased because the fuel consumption will be increased in

compare with the cruise phase. Cp was assumed to be 0.45 for loiter phase.

ηp in loiter phase will be decreased because the speed of propeller is decreased. So, it

is assumed to be 75 % for loiter phase.

Table 1-6 suggested values for L/D,CP, according to Roskam method for Loiter

L/D CP(lbs/hp/hr) Vltr(mph) Eltr (hours) P

Loiter 14 0.45 70 1 0.75

1

ln1

375)(

i

i

ltrltrP

P

ltr

ltrW

W

D

L

CVhoursE

(1.11)

Using information of table1-5 and Eq. 1.11 results:

5

6

W

W= 0.992

So for Mff:

ffMTOW

W1

1

2

W

W

2

3

W

W

3

4

W

W

4

5

W

W

5

6

W

W

6

7

W

W

7

8

W

W= 0.9170 (1.12)

For reserved fuel below equation is used:

resFWresFTOffF WWMW )1( (1.13)

Substituting the magnitudes in the main equation for WTO, results:

(1.14)

(1.15)

Solving the equation system below results:

(1.16)

(1.17)

HADAFTM


It is considerable that suggested values in tables 1-4 and 1-5 affect the takeoff gross

weight and the empty weight indirectly. For example if the Cp is increased it is

reasonable that the empty and take off gross weight will be increased because the fuel

consumption is increased. It means the plane needs more fuel for a specific value of

loiter and range. Similarly if the L/D and propeller efficiency are increased, the empty

and take off gross weight will be decreased.

WTO = 626.60 kg

WE = 335.74 kg

HADAFTM


1.5 SENSETIVITY ANALYSIS

It is evident from the way the results in previous sections were obtained, that their

outcome depends on the values selected for the various parameters in the range and

endurance equations.

Data calculated in this part, shows how the take-off gross weight of HADAF1404,

varies with parameters below.

1. Payload, WPL

2. Empty weight, WE

3. Range, R

4. Endurance, E

5. Lift-to-drag ratio, L/D

6. Specific fuel consumption, CP

7. Propeller efficiency, P

After preliminary sizing it is mandatory to conduct sensitivity studies on the

parameters 1-7 listed above.

The reasons for doing this are:

A. To find out which parameters drive the design

B. To determine which areas of technological change must be pursued, if some

new mission capability must be achieved.

C. If parameters 5, 6 or 7 are selected optimistically (or pessimistically), the

sensitivity studies provide a quick estimate of the impact of such optimism (or

pessimism) on the design.

With the help of Equations (1.1) to (1.3) and assumptions made in the previous part

the following simplifications could be done:

WE = WTO – WF – WPL – Wtfo – Wcrew

WF = (1- Mff)WTO + Wres = (1- Mff)WTO + Mres (1- Mff) WTO (1.19)

With a substitution we have:

WE = WTO {1- (1+ Mres)(1-Mff) – Mtfo)} – (WPL + Wcrew) (1.20)

The latter can be written as:

WE = CWTO – D (1.21)

HADAFTM


where:

C = {1- (1+ Mres)(1-Mff) – Mtfo)} (1.22)

and:

D = (WPL + Wcrew) (1.23)

WE can be eliminated from equation (1.3) to yield:

log10 WTO = A + B log10 (C.WTO – D) (1.24)

If the sensitivity of WTO to some parameter y is desired, it is possible to obtain the

sensitivity, by partial differentiation of WTO in equation (5.5). This results in:

DCW

y

D

y

WC

y

CWB

y

W

W TO

TOTO

TO

TO

)(

)(1

(1.25)

Since the line constants A and B vary only with airplane type, the partial derivatives

yA

and y

B

are zero. Simplifying equation (6.5) results:

DWBC

y

DBW

y

CWB

y

W

TO

TOTO

TO

)1(

))(( 2

(1.26)

The parameter y can be any one of those listed as 1-7 at the beginning of this section.

Now it‟s time to derive the sensitivities.

1.5.1 Sensitivity of Takeoff weight to Payload weight

Using equation (7.5) for sensitivity to payload weight results:

1})1({

0.0

0.1

TOTO

PL

TO

PL

PL

PL

WBCDBWW

W

W

C

W

D

Wy

(1.27)

HADAFTM


Having the assumptions below in mind the result will be:

Assumptions: A= 0.096, B = 1.069, C = 0.9165, D = 230, WTO = 626

1.5.2 Sensitivity of Takeoff weight to Empty weight

By partial differentiation of WTO with respect to WE the take-off weight to empty

weight sensitivity is expressed as:

B

AW

TO

E

TO

TO

e

BW

W

W

)(log

10 (1.28)

Having the assumptions below in mind the result will be:

Assumptions: A= 0.096, B = 1.069, C = 0.9165, D = 230, WTO = 626

53.51

E

TO

W

W

1.5.3 Sensitivity of Takeoff weight to Range, Endurance, Speed, Specific Fuel

Consumption, Propeller Efficiency and Lift-to-Drag Ratio

Withdrawing the time-consuming derivation of the equations, the following set of

relations is derived:

Table 1-7 BreguitpartialsforpropellerdrivenairplanesAdoptedfrom“AirplaneDesign”writtenbyDr.John

Roskam, part I, preliminary sizing

Range/Endurance Case Y ∂R

/∂y / ∂E

/∂y

Range R 1375

)(

DLC

yR

pp

Endurance E 1375

)(

DLVC

yE

pp

Range Cp 1375

)(

DLR

yR

p

Endurance Cp 1375

)(

DLVE

yE

p

Range ηp 12375

))((

DLRC

yR

pP

_ _

481.2

PL

TO

W

W

HADAFTM


Endurance ηp 12375

))((

DLEVC

yE

pp

Range V ---------------------------

Endurance V 1375

)(

DLEC

yE

pp

Range L/D 12375

))((

DLRC

yR

pP

Endurance L/D 12375

))((

DLEVC

yE

pp

The general equation for calculating sensitivity is:

y

yF

y

WTO

(1.29)

Where:

ffresTOTO MMDBCWWBF )(})({)( 11 12 (1.30)

For the goal airplane, it is assumed that Mres= 0. Now having the following values for

other terms the result for F is as follows:

Assumption: A= 0.096, B = 1.069, C = 0.9165, D = 230, WTO = 626, Mff=0.9170

F = 1437.55

So using equation (12.5), formulas of table 4 and value of F, calculation of airplane

gross factors due to range, endurance, speed, specific fuel consumption, propeller

efficiency and lift-to-drag ratio will be done which is listed in table5.

HADAFTM


Table 1-8results for airplane growth factors due to range, endurance, speed, specific fuel consumption,

propeller efficiency and lift-to-drag ratio

Range/Endurance Case Y ∂W

TO/∂y

Range R ∂W

TO/∂R = 0.1438 lbs/nm

Endurance E ∂W

TO/∂E = 11.50lbs /hr

Range Cp ∂W

TO/∂Cp= 184.30kg/lbs/hp/hr

Endurance Cp ∂W

TO/∂Cp= 25.56kg/lbs/hp/hr

Range ηp ∂W

TO/∂ P = -89.85lbs

Endurance ηp ∂W

TO/∂ P =-15.33lbs

Range V ---------------------------

Endurance V ∂W

TO/∂V = 0.1642 lbs/mph

Range L/D ∂W

TO/∂L/D = -5.529 lbs

Endurance L/D ∂W

TO/∂L/D = -0.821lbs

HADAFTM


1.6 Appendix:

Secondary database after the first step of data filtering was processed.

Max Gross

Wt(kg)

Standard Empty

Wt(kg) N.O.Seats type

Zodiac CH-601 HDS 544.31 267.6 2 side by side

Thorp T-211 576 354 2 side by side

The Taylorcraft BC-12 544 331 2

The Luscombe Model 8 Silvaire 8A 545 302 2 side by side

Tetras 495 284 2 side by side

ST3 UL (Calypso) 450 265 2 side by side

SportCruiser 599 330 2 side by side

Skylark 599 296 2 side by side

Seamax 598 340 2

pioneer 200 472 260 2 side by side

Parrot 599 360

K-10 SWIFT LSA 575 285 2 side by side

Jodel D18 460.4 236 2

Jabiru SP 469 234.5 2 side by side

FM250 Vampire mk1 450 265 2 row

F99 Rambo 470 285 2 side by side

DYNAMIC WT9 550 300 2 side by side

DynAero MCR-01 Sportster VLA 490 260 2 side by side

Dart 544 283 2

CT2K 480 258 2 side by side

CORBEN JUNIOR ACE 556 293 2 side by side

JK-05 560 280 2

SportCruiser 1 599 306 2

F99 LSA 594 308 2

Bushbaby Explorer 550 260 2

jabiru j-170 545 290 2 side by side

tecnam 598 331 2

EUROFOX 558 288 2 side by side

remos 598 303

HADAFTM


Climb rate

(ft/min)

max speed

(kt)

stall speed

(kts)

stall speed

(m/s)

Zodiac CH-601 HDS 1300 140 48 24.6912

Thorp T-211 <750 138 39 20.0616

The Taylorcraft BC-12 500 96 33 16.9752

The Luscombe Model 8 Silvaire 8A 900 185 0

Tetras 1300 113 30 15.432

ST3 UL (Calypso) 1000 116 35 18.004

SportCruiser 1200 139 30 15.432

Skylark 1200 156 36 18.5184

Seamax 1000 139 41 21.0904

pioneer 200 1000 130 33 16.9752

Parrot 1000 137 35 18.004

K-10 SWIFT LSA 984 119 37 19.0328

Jodel D18 650 106 40 20.576

Jabiru SP 1000 110 40 20.576

FM250 Vampire mk1 984 119 35 18.004

F99 Rambo 1200 124 33.5 17.2324

DYNAMIC WT9 1000 151 0

DynAero MCR-01 Sportster VLA 1750 172 47 24.1768

Dart 1200 170 56 28.8064

CT2K 1000 167 33 16.9752

CORBEN JUNIOR ACE 600 113 38 19.5472

JK-05 1574.8 111.24 29.5559 15.235

SportCruiser 1 1181.1 139.32 29.5559 15.235

F99 LSA 1082.675 138.24 499.7634 257.61

Bushbaby Explorer 984.25 112.86 33.85494 17.451

jabiru j-170 700 125 38 19.5472

tecnam 1200 154 37 19.0328

EUROFOX 980 110 35 18.004

remos 1050 120 38 19.5472

HADAFTM


fuel capacity

(L) Powerplants

Power

(hp)

Zodiac CH-601 HDS 72.74 Rotax 912ULS (100Hp) 100

Thorp T-211 79.5 0-200-A, 100hp @2700 rpm 100

The Taylorcraft BC-12 68 Continental A-65 flat four piston engine 65

The Luscombe Model 8 Silvaire 8A Continental A-65 flat four piston engine 65

Tetras 85 912 S (100 hp) 100

ST3 UL (Calypso) Jabiru 2200 (85 hp) 85

SportCruiser 114 Rotax 912 ULS(100 hp) 100

Skylark 90 Rotax 912S(100hp) 100

Seamax 93.37 Rotax 100

pioneer 200 54 Rotax 912 - 100 Engine incl.Airbox 100

Parrot 114 Rotax 912ULs (100 PS) or Jabiru 3300 100

K-10 SWIFT LSA 80 4 - Stroke Rotax 912ULS(100hp) 100

Jodel D18 65 Limbach EO2X L2000 80

Jabiru SP 65 Jabiru 2200cc (85hp) 85

FM250 Vampire mk1 65 4 - Cylinder 4 - Stroke Rotax 912UL 80

F99 Rambo Rotax 912 80

DYNAMIC WT9 99.93

DynAero MCR-01 Sportster VLA 75 Rotax (100 hp) 100

Dart 95 VW 2100cc HP Range 80/80-150

CT2K 110 Rotax912S 100

CORBEN JUNIOR ACE 83 Continental, Lycoming

JK-05 60 Rotax 912 80

SportCruiser 1 112 Rotax 912 ULS 100 HP 100

F99 LSA Rotax 912 S 100

Bushbaby Explorer 100 Rotax 912 ULS 100

jabiru j-170 134 Jabiru 2200 85 hp 85

tecnam Rotax 912 ULS2 Engine (100 hp) 100

EUROFOX 75 Rotax 912 80

remos 912 UL-S 100

HADAFTM


Wingspan

(m)

Wing Area

(sqm)

wing loading

(lbs/sq.ft) Clmax

Zodiac CH-601 HDS 7.01 9.1 12.25748141 1.5714

Thorp T-211 7.6 9.7 12.1687809 2.3631

The Taylorcraft BC-12 10.98 17.1 6.519272158 1.7682

The Luscombe Model 8 Silvaire 8A 10.68 13 8.591113812 #DIV/0!

Tetras 10.1 15.7 6.461031658 2.1204

ST3 UL (Calypso) 9.4 9.31 9.905106633 2.3883

SportCruiser 8.78 13.2 9.299277628 3.0519

Skylark 7.92 9.38 13.08640348 2.9825

Seamax 8.74 12.07 10.15290299 1.784

pioneer 200 7.55 10.5 9.211895911 2.4985

Parrot 9.5 11 11.15913315 2.6907

K-10 SWIFT LSA 9.1 11.8 9.98578382 2.1545

Jodel D18 7.5 9.84 9.588187959 1.77

Jabiru SP 8.02 7.89 12.18125857 2.2487

FM250 Vampire mk1 7.8 10.05 9.175775398 2.2124

F99 Rambo 9.1 10.1 9.536144135 2.5099

DYNAMIC WT9 9 10.3 10.94263183 #DIV/0!

DynAero MCR-01 Sportster VLA 6.63 5.2 19.31030169 2.582

Dart 7.01 7 15.92565056 1.5

CT2K 10.22 12.06 8.156244798 2.2122

CORBEN JUNIOR ACE 8 10.21 11.15951633 2.2827

JK-05 10.76 9.72 11.80641608 3.9756

SportCruiser 1 8.5 11.8 10.40258175 3.5029

F99 LSA 9.1 10.1 12.05206301 0.0142

Bushbaby Explorer 9.06 12.45 9.052940386 2.3234

jabiru j-170 9.6 9.29 11.8 2.4591

tecnam 8.99 12.4 9.882704761 2.1322

EUROFOX 9.2 11.5 9.943348957 2.3975

remos 9.29 10.96 11.18116232 2.2871

HADAFTM


1.7 References

1) Roskam, J., Airplane Design: Part II, Preliminary Configuration Design and

Integration of the Propulsion System.

2) Mattingly, J.D., Elements of Propulsion: Gas Turbines and Rockets

3) Anderson, J.D., Fundamentals of Aerodynamics

HADAFTM

1404 Aircraft Design Book Performance Estimation

2 PERFORMANCE ESTIMATION

2.1 INTRODUCTION

In addition, to meeting range, endurance and cruise speed objectives, airplanes are

usually designed to meet performance objectives in the following flight regimes:

a. Stall speed

b. Take-off field length

c. Landing field length

d. Cruise speed (or maximum speed)

e. Climb rate

There are some airplane design parameters which affect the performance flight

regimes listed above. These parameters are:

1. Wing area, S

2. Take-off trust, TTO or take-off power, PTO

3. Maximum required take-off lift coefficient with flaps up: CL ,max (clean)

4. Maximum required lift coefficient for take-off, CL ,max TO

5. Maximum required lift coefficient for landing, CL ,max L , or CL ,max PA

The purpose of this part is to determine a range of values of wing loading, W/S, thrust

loading, T/W, and maximum lift coefficient, CL,max, within which certain performance

requirements are met. Combination of the highest possible wing loading and the

lowest possible thrust loading (or power loading) which still meets all performance

requirements, results in an airplane with the lowest weight and the lowest cost.

2.2 SIZING TO STALL SPEED REQUIREMENTS

The mission task demands a stall speed not higher than some minimum value. As

certified by the FAR23, single-engine airplanes may not have a stall speed greater

than 61 kts at WTO.

The power-off stall speed of an airplane may be determined from:

(2.1)

HADAFTM


By specifying a minimum allowable stall speed at some altitude, Eq. (2.1) defines a

maximum allowable wing loading W/S for a given value of CL. Table2-1 presents

typical values for CL for homebuilt airplanes.

Table 2-1Typical values for maximum lift coefficient

Airplane Type CL ,max CL ,max TO CL ,max L

Homebuilts 1.2-1.8 1.2-1.8 1.2-2.0

Values which are assumed during the design process are as listed in the following

table (table2-2). It is clear that CL max is strongly influenced by wing and airfoil

design, flap type, size and center of gravity location.

Table 2-2Assumptions made for calculating the stall speed requirement meeting criteria

Eq. (2.1) and Table2-2 may be combined to yield:

a. To meet the flaps down requirement:

(

)

(2.2)

b. To meet the flaps up requirement:

(

)

(2.3)

Therefore, to meet both requirements, the take-off wing loading, (W/S) TO must be less

than 10.9783 lbs/sq.ft. Figure2-1 illustrates it. The stall speed requirement was

formulated as a power-off requirement It means that neither power loading nor thrust

loading are important in this case, as seen in figure2-1:

VS(kts) CL ,max(clean) CL ,max TO CL ,max L ρ (lbm/ft2)

45 1.6 1.8 2 0.062

HADAFTM


Figure 2-1Stall speed sizing – illustrates the acceptable region for W/S and W/P values in according to stall

speedcriteria

2.3 SIZING TO TAKE-OFF DISTANCE REQUIREMENTS

Take-off distances of airplanes are determined by the following factors:

1. Take-off weight, WTO

2. Take-off speed, VTO

3. Trust-to-weight ratio at take-off, (T/W)TO (or weight-to-power ratio, (W/P)TO

4. Aerodynamic drag coefficient, CD and ground friction coefficient, µG

5. Pilot technique

Take-off requirements are normally given in terms of take-off field length

requirements. FAR23 and FAR25 criteria can be used for doing the design process in

this part. This requirement differs widely and depends on the type of airplane. So

FAR23 requirements are chosen during sizing process because FAR23 airplanes

usually are propeller driven airplanes.

Figure2-2 presents a definition of take-off distances used in the process of sizing an

airplane to FAR23 requirements.

9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 140

20

40

60

80

100

120

140

160

180

200

Wing Loading(lbs/sq.ft)

Pow

er

Loadin

g(lbs/h

p)

Clmax Clean = 1.6

Clmax Take-off = 1.8

Clmax Landing = 2

Stall SpeedRequirements met

HADAFTM


Figure 2-2Definition of FAR 23 take-off distances

STOG (take-off ground roll) is proportional to (W/S)TO, (W/P)TO and CL,max TO :

(2.4)

Where TOP23 is take-off parameter for FAR23 airplanes:

Figure 2-3Take-off parameter vs. Take-off distances for HADAF1404 Database

Figure2-3 relates STOG to take-off parameter for a range of single engine. There is a

lot of scatter in the data. Because take off procedures vary widely and take-off thrust

depends strongly on propeller characteristics. Nevertheless, it is useful to employ the

correlation line of figure in the preliminary sizing. It is a polynomial trend line with

2nd

order which has an intercept of zero. The correlation line suggests the following

relationship:

(2.5)

0 10 20 30 40 50 60 700

100

200

300

400

500

600

Ta

ke

-off D

ista

nce

(ft)

TOP 23

y=0.0058*x2 + 6.8552*x

HADAFTM


According to figure 3.4 of first part of Roskam book (Figure2-4), STO (take-off

distance) can be related to STOG by the following relationship:

(2.6)

Figure 2-4 Correlation of Ground Distance and Total Distace for Take-off (FAR23)

But HADAF1404 database (Figure2-5) suggests the following relationship between

STO (take-off distance) and STOG:

(2.7)

Figure 2-5 Correlation of Ground Distance and Total Distace for Take-off (According to HADAF1404

Databese)

60 80 100 120 140 160 180 200 220 240 260

100

150

200

250

300

350

400

450

500

550

Take-off Distance Ground Roll (m)

Ta

ke

-off D

ista

nce

ove

r 1

5m

(m

)

Data form Database

Linear fittig for Data

Sto=1.98Stog

HADAFTM


The average take-off ground roll distance for data in HADAF1404 database is 470 ft.

So it is assumed to meet the following take off criteria:

On the other hand, it is necessary that:

(2.8)

is the density ratio of air. At sea level, it‟s 1.00; at 5,000 feet, it‟s 0.8616; and at

10,000 feet, it‟s 0.7384. Since this results the following relationship:

(

)

(

)

(2.9)

Figure 2-6Effect of take-off wing loading and maximum take-off distance on take-off power loading

Figure2-6 translates Eq. 2.9 into diagrams of (W/S)TO to (W/P)TO for given values take

of distance and CLmax,TO= 1.8 . As it is seen in figure2-6, if the takeoff distance

decrease the minimum allowable power loading will be decreased. It means that the

airplane should have more engine power to take off in a short take-off distance.

1 2 3 4 5 6 7 8 9 100

20

40

60

80

100

120


Pow

er

Loadin

g(lbs/h

p)

TOP23 take-off distance=470 ft



HADAFTM


2.4 . SIZING TO LANDING DISTANCE REQUIREMENT

Landing distances of airplanes are determined by four factors:

1. Landing weight, WL

2. Approach speed, VA

3. Deceleration method used

4. Flying qualities of the airplane

5. Pilot technique

Landing distance requirements are nearly always formulated at the design landing

weight, WL of the airplane. According to first part of Roskam book, WL is related to

WTO as.

Table 2-3Typical values for landing weight to take-off weight ratio for single engine propeller driven

Minimum Average Maximum

0.95 0.997 1

Also according to the before section (weight estimation) and Eq. 4.4 and Eq. 5.4

landing weight is related to take-off weight as below:

(2.10)

On the other hand, according to kinetic energy considerations, total landing distance

is proportional to approach speed with 2nd

order.

Like sizing for take-off distance requirement, in this part FAR23 and Far25 criteria

can be used. We choose FAR23 again because of our propeller driven airplane.

Figure2-7 presents a definition of landing distances used in the process of sizing an

airplane to FAR23 requirements. It is known that there is the following relation for

approach speed and stall speed:

(2.11)

Figure 2-7Definition of FAR 23 landing distances

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Also it is known:

(2.12)

Figure 3.13 of first part of Roskam book (Figure2-8) suggests the following

relationship between the landing ground run, SLG and the square of the stall speed,VS

landing.

(2.13)

In Eq. 2.13 the distance is in ft and the stall speed is in kts.

Figure 2-8Effect of Square of Stall Speed on Landing Ground run

Figure 2-9Effect of Square of Stall Speed on Landing Distance for HADAF1404 Database

0 500 1000 1500 20000

200

400

600

800

1000

1200

1400

Square of Stall Speed (Vs2) (Kts

2)

La

nd

ing

Dis

tan

ce

, S

l (f

t)

Data from Database

linear fitting for Data

Sl=0.516*Vs2

HADAFTM


Also Figure2-9 which is drawn according to HADAF1404 Database suggests another

relationship (Eq. 2.14) between the landing distance, SL and the square of the stall

speed.

(2.14)

In Eq. 2.13 the distance is in ft and the stall speed is in kts.

The average landing distance for data in HADAF1404 database is 700 ft. It is

required to size a landing distance of 1100 ft (335 m). So it follows that:

(2.15)

(2.16)

Finally, this translates into the following requirement:

(2.17)

Also as it mentioned above, the design landing weight is specified as:

and it follows that:

(2.18)

At last, figure2-10 Present the range of value of (W/S)TO and for a given

value of which meet the landing distance requirement.

Figure 2-10 the range of value of (W/S)TO and CL,max foragivenvalueofρ=1.225

17.5 18 18.5 19 19.5 200

20

40

60

80

100

120

140

160

180

200


Pow

er

Loadin

g(lbs/h

p)

Landing Distance(SL)= 1100 ft

HADAFTM


2.5 SIZING TO CLIMB REQUIREMENT

All airplanes must meet certain climb rate or climb gradient requirements. To size an

airplane for climb requirements, it is necessary to have an estimate for the airplane

drag polar. So a rapid method for estimating drag polar for low speed flight

conditions have been used in this section described as followed.

In a parabolic drag polar, the drag coefficient of an airplane can be written as:

(2.19)

Where A is the aspect ratio and e is the Oswald number and finally the zero-drag

coefficient can be expressed as:

where f is the equivalent area and S is

the wing area.

On the other hand, according to figures 3.21a and b of first part of Roskam book

(Figure 2-11), it is possible to relate f to wetted area Swet. The relationship between

these two parameters is : (2.20)

Figure 2-11 Effect of equivalent Skin friction on parasite and wetted areas

HADAFTM


It is considerable that the coefficients a and b themselves are a function of the

equivalent skin friction coefficient of an airplane, Cf as seen in figure2-11. The latter

is determined by the smoothness and streamlining designed into the airplane. These

coefficients can be calculated from figure2-11. The Cf is assumed to be 0.005.

Table 2-4Correlation coefficient for Parasite area vs. Wetted area

Cf A B

0.005 -2.3010 1.0000

It is so clear that the method for estimating drag boils down to the ability to predict a

realistic value for Swet. Fortunately, Swet correlates well with WTO for a wide range of

airplanes. Again, according to figure 3.22 of 1st part of Roskam book (Figure 2-12)

Swet can be related to WTO with following relationship:

(2.21)

Figure 2-12 Correlation between wetted area and take-off weight

HADAFTM


Values for c and d is obtained by correlating wetted area and take-off weight data

which is done by reference book.

So it is easily possible to abtain an initial estimate for airplane‟s wetted area without

knowing what the airplane actually looks like.

Table 2-5Coefficients A and B of wetted area eqution

Type C d

Homebuilts 1.2362 0.4319

Since an estimate for WTO was already obtained in previous book ( weight estimation),

the drag polar for the clean airplane can now be determined. So the cruise

requirement should be investigated for an airplane with WTO equal to 626kg(1380

lbs). By using the relationship between WTO and SWet,it is possible to estimate Swet as

below:

(2.22)

Then it is possible to estimate parasite area, f, as following:

(2.23)

It is assumed that the aspect ratio to be equal to 7.5. According to Figure 2-13

Oswald number is assumed equal to 0.91. So easily and can be calculated.

Since it is better to minimize CD, the wing area should be maximized. So it is assumed

that the wing area, S, which is the minimum wing area in database to be equal to 8

m2(86 ft

2). It follows:

(2.24)

Now it is possible to find the clean drag polar at low speed:

(2.25)

For take-off and landing the effects of high lift devices and the landing gear, which

are strongly dependent on their size and type, need to be accounted for. These items

are defined as . Typical values for

are given in the following table(Table2-6)

HADAFTM


Figure 2-13 Effect of aspect ratio and sweep angle on wing efficiency factor (Oswald number)

Table 2-6 Firstestimatesfor∆CD0 and the Oswald No. "e"

Configuration e

Clean 0 0.8-0.85

Take-off flaps 0.01-0.02 0.75-0.8

Landing flaps 0.055-

0.075 0.7-0.75

Landing gear 0.015-

0.025 No effect

The additional zero-lift drag coefficients due to flaps and landing gear are as follows:

due to :

Take off flaps = 0.02

Landing gear = 0.02

And finally the airplane drag polar at take-off with gear down can be represented as:

(2.26)

HADAFTM


It is time to get back to the main goal, sizing to climb requirement. The take off climb

requirements of FAR 23 can be summarized as follow:

- All airplane must have a minimum climb rate at sealevel of 300 fpm and a

steady climb angle of at least 1:12 for landplanes.

Also the balked landing climb requirement of FAR 23 can be summerized as follows:

- The steady climb angle shall be at least 1:30 with the airplane in an specific

configuration.

Loftin has been represented a method for estimating rate of climb (RC) and climb

gradient (CGR) of an airplane in reference 2 (Loftin). All airplanes in this method

should have the following criteria for sizing to rate of climb:

(2.27)

Where :

(2.28)

It is better to maximize RC, so it is evidently necessary to make

as large as

possible. Fortunately, this has been noted before and CD has been minimized.

Also Loftin represents all ingredients needed for sizing to climb gradient criteria as

below :

(2.29)

And

√ (2.30)

Where :

(2.31)

To find the best possible climb gradient, it is necessery to find the minimum value of

CGRP. This minimum value depends on the the lift coefficient and on the

corresponding lift to drag ratio. Evidently, the minimum of this parameter is usually

found at a value of CL very close to . In other hand, some margin relative to stall

speed is alwaye desired. But this margin are not specified by Federal Aviation

Regulation in detail. So it is suggested to ensure that a margin of 0.2 exists between

and

.

HADAFTM


As it is said above, in the case of FAR 23 climb requirement:

(2.32)

By assuming and and with the take off configuration,as

before calculated, the drag polar is as following :

√

√ (2.33)

Figure2-13 translates Eq. 2.33 into regions of (W/S)TO and (W/P)TO .

Figure 2-13Effect of FAR 23 rate of climb requirements on the allowable values of take off thrust to weight

ratio and take off wing loading

Climb gradient requirements are computed as below:

(2.34)

As said before, CGR=1/12 rad=0.0833 and for this case the drag polar is :

0 2 4 6 8 10 12 14 16 18 2020

25

30

35

40

45

50

55

Wing loading (lbs/sq.ft)

Po

we

r lo

ad

ing

(lb

s/h

p)

climbrequirementsmet

HADAFTM


it is assumed that with take off flaps the value of So with a

margin of 0.2, the value of will be equal to 1.4.This yields:

(2.35)

Therefore :

(2.36)

Figure2-14 translates Eq. 2.36 into regions of (W/S)TOand (W/P)TO .

Figure 2-14Effect of FAR 23 climb gradient requirements on the allowable values of take off thrust to weight

ratio and take off wing loading

Figure2-15 shows the effect of FAR 23 climb requirements on the allowable values of

take off thrust to weight ratio and take off wing loading and it also shows the region

that all climb requirements in case of FAR 23 can be met.

0 2 4 6 8 10 12 14 16 18 2010

20

30

40

50

60

70

80

90


Po

we

r L

oa

din

g (

lbs/h

p)

climbgradientrequirementsmet

HADAFTM


Figure 2-15Effect of FAR 23 climb requirements on the allowable values of take off thrust to weight ratio

and take off wing loading

In the case of the FAR, part 23 which is related to balked landing climb requirement

the gradient should be equal to 1:30. This means that CGR=0.0333 rad.

By assuming and , the drag polar and the corresponding lift to

drag ratios in this case are:

(2.37)

Therefore :

(2.38)

0 2 4 6 8 10 12 14 16 18 2020

30

40

50

60

70

80

90

100

110


Po

we

r L

oa

din

g (

lbs/h

p)

all climbrequirementsmet

HADAFTM


Figure 2-16 Effect of FAR 23.77 requirements on the allowable values of take off thrust to weight ratio and

take off wing loading

Figure 2-16 translates Eq. 2.38 into regions of wing loading and thrust loading.

2.6 SIZING TO CRUISE SPEED REQUIREMENT

Cruise speed for propeller driven airplanes is usually calculated at 75 to 80 percent

power. In that case it can be shown that the induced drag is small in comparison with

the profile drag. is assumed to be:

(2.39)

In the case of HADAF1404 Airplane if it is assumed to be at cruise

condition, according to Eq. 2.26, the induced Drag will be equal to 0.004194. So it is

reasonable to assume

.

Loftin showed that because of this fact, cruise speed turns out to be proportional to

the following factor:

(2.40)

Also from this, he found the following proportionality between Vcr and IP:

(2.41)

0 2 4 6 8 10 12 14 16 18 2020

40

60

80

100

120

140


Po

we

r L

oa

din

g (

lbs/h

p)

climbgradientrequirementsmet

HADAFTM


For a given desired cruise speed the parameter IP which is called the power index can

be estimated from figure 2-17 and figure 2-18. In fact, these figures can show how

cruise speed is related to IP for a range of example airplanes which indirectly copied

from reference 2, and for airplanes in HADAF1404 database, respectively.

Figure 2-17Correlation of airplanes speed with power index for biplanes and strutted monoplanes with fixed

gear

The direct relationship between power index, wing and thrust loading is as followed:

(2.42)

HADAFTM


Figure 2-18Correlation of airplanes speed with power index for airplanes in HADAF1404 database

HADAF 1404 must achieve a cruise speed of 185km/h (115mph) at 75 percent power

at cruise condition at take-off weight. In this case according to figure 2-18 the power

index is equal to 0.95. Also at cruise condition (10000 ft), . Therefore it is

found that:

(2.43)

0 0.2 0.4 0.6 0.8 1 1.2 1.40

20

40

60

80

100

120

140

160

Power Index, Ip

Sp

ee

d, V

, m

ph

Data from Database

linear fitting for data

V=120*Ip

HADAFTM


Figure 2-19Allowable values of wing loading and power loading to meet a given cruise speed

Figure2-19 shows the range of combinations of wing loading and power loading by

translating Eq. 2-43 for which the cruise requirements is met.

It is considerable that (W/P) in the figure 2-19 is at cruise condition (10,000 ft). It is

necessary to transfer that ratio to sea level. In this case it must be multiplied by the

power ratio for cruise power at 10,000 ft to that sea level which is typically 0.7 for

reciprocating engine without supercharging. Figure 2-20 compares these two

parameters at 10,000 ft and sea level condition.

(2.44)

0 2 4 6 8 10 12 14 16 18 200

5

10

15

20

25

30

35


Pow

er

Loadin

g(lbs/h

p)

Cruise Speed Requirement

HADAFTM


Figure 2-20

2.7 MATCHING OF ALL SIZING REQUIREMENT

Considering a series of relations between:

- Take off power loading ,

- Take off wing loading ,

- Maximum required lift coefficient ,

- And aspect ratio ,

It is now possible to determine the best combination of these quantities for the design.

The word best is used rather than optimum because the latter implies a certain

mathematical precision. What is usually done at this point is to overlay all

requirements and select the highest possible power loading and wing loading which

are consistent with all requirements. This process is also known as matching process

and this selected point is known as the design/matching point.

After calculating all requirements, figure2-15 shows how these requirements restrict

the useful range of combinations of takeoff wing loading (W/S)TO and take off power

loading (W/P)TO.

0 2 4 6 8 10 12 14 16 18 200

5

10

15

20

25

30

35


Pow

er

Loadin

g(lbs/h

p)


Cruise Speed Requirement,Take-off power

HADAFTM


Figure 2-21 Matching results

Figure 2-222-4Final Favorable Area

By examining the matching diagram, point (9.8, 11.77) seems a reasonable choice.

Because it has highest possible wing and power loading. It means that the wing

2 4 6 8 10 12 14 16 18 200

20

40

60

80

100

120

140

160

180

200


Po

we

r L

oa

din

g(lb

s/h

p)

Stall Speed Requirement,Cl max = 1.6

Stall Speed Requirement,Cl max t = 1.8

Stall Speed Requirement,Cl max l = 2

Take of Distance Requirement

Landing Distance requirement

Climb requirement

Climb Gradient Requirement

Balked Landing Requirement


2 4 6 8 10 12 14

2

4

6

8

10

12


Po

we

r L

oa

din

g(lb

s/h

p)

Stall Speed Requirement,Cl max = 1.6

Stall Speed Requirement,Cl max t = 1.8

Stall Speed Requirement,Cl max l = 2

Take of Distance Requirement

Landing Distance requirement

Climb requirement

Climb Gradient Requirement

Balked Landing Requirement


All Requirement

met

Design Point

HADAFTM


loading of the airplane is 9.6 lbs/sq.ft and the power loading of the airplane is 13.78

lbs/hp. With this choice, our airplane is now characterized by the following design

parameters:

{

}

(2.45)

The following table shows the results which are extracted from this part (performance

estimation). These results will be used in the future books. The power loading will be

used in engine book to determine the engine power required for HADAF at takeoff.

Also, the wing loading will be used in wing book to determine the required wing area.

Table 2-7Final results

Wing Area

(sq.ft)

Power

(hp)

140.51 116.99

HADAFTM


2.8 ROAD MAP

Finally, the below diagram shows the outline we stated in this book visually:

HADAFTM


2.9 APPENDIX

All data calculated in this book are computed by a code which is programmed by

MATLAB®, the Language of Technical Computing. The following program is the open

source of this code:

clear all clc grid on %1)Sizing to Stall Speed Requirements p=1.225; vstall=23.15;%(m/s) clmax=1.6;%clmax clean clmaxt=1.8;%clmax take off clmaxl=2;%clmax landing sicma=1; wingloading1=(1/2*p*(vstall^2)*clmax)*0.225/10.764 for i=1:1:201 wing_loading(i)=wingloading1; end powerloading=0:1:200;

hold on plot(wing_loading,powerloading,'g','LineWidth',2) title('') xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on hold on

wingloading2=(1/2*p*(vstall^2)*clmaxt)*0.225/10.764 for i=1:1:201 wing_loading(i)=wingloading2; end powerloading=0:1:200; plot(wing_loading,powerloading,'r','LineWidth',2)

hold on wingloading3=(1/2*p*(vstall^2)*clmaxl)*0.225/10.764 for i=1:1:201 wing_loading(i)=wingloading3; end powerloading=0:1:200; plot(wing_loading,powerloading,'k','LineWidth',2) legend('Clmax Clean = 1.6','Clmax Take-off = 1.8','Clmax Landing = 2')

%2)Sizing to Take off Distance Requirements

STO=470;%ft for l=1:3 if l==1 STOn=STO; elseif l==2 STOn=STO+0.25*STO; else STOn=STO-0.25*STO;

HADAFTM


end clear Top23 syms top23 eq1=6.855*top23+0.0058*(top23)^2-STOn; disp('STO = '),disp(eq1); m=solve(eq1);m=double(m) i=length(m); for k=1:i if m(k)>0 TOP23=m(k) end end

wingloading=1:0.1:20; powerloading=(sicma*clmaxt*TOP23)./wingloading; hold on plot(wingloading,powerloading,'b','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on hold on legend('TOP23 take-off distance=90 ft') end

%3)Sizing to Landing Distance Requirements

vapproach=1.3*vstall; SL=1100; SLG=SL/1.938; VsL=sqrt(SL/0.516); VsL=0.514*VsL

wingloading4=(1/2*p*(1.3*VsL^2)*clmaxl)*0.225/10.764 for i=1:1:201 wingloading(i)=wingloading4; end powerloading=0:1:200;

hold on plot(wingloading,powerloading,'g','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Landing Distance(SL)= 1100 ft')

%Polar Drag estimation

%inputs c=1.2362;d=0.4319;a=-2.3010;b=1; e=0.91;%oswald No. AR=7.5;%aspect ratio wto=1377.889;%pound deltacd0_take_off_flaps=0.020; deltacd0_landing_gear=0.020; Smin=86;%sq.ft

swet=10^(c+(d*log10(wto)))%log10(Swet)=c+d*log10(wto) f=10^(a+b*log10(swet))%log10(f)=a+b*log10(Swet) smin=86;%sq.ft cd0=f/smin

HADAFTM


deltacd0=deltacd0_landing_gear+deltacd0_take_off_flaps; cd0=cd0+deltacd0 k=1/(3.14*AR*e) syms cl cd=cd0+k*cl^2 disp(clmax) cd=cd0+k*clmax^2

%4)Rate of climb Requirements

%inputs rc=300;%as in the case of FAR23 Climb Requirements rcp=(1/33000)*rc; etap=0.8;

wingloading=1:0.1:20; powerloading=etap./(rcp+((wingloading).^0.5)/((19*1.345*(AR*e)^0.75)/(cd0^0

.25)));

hold on plot(wingloading,powerloading,'g','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Rate of Climb Requirement')

%Climb Gradient Requirements

%inputs etap=0.8; CGR=1/12;%Radian

clclimb=clmaxt-0.2; cd=cd0+k*clclimb^2 lift_to_drag_ratio=clclimb/cd wingloading=1:0.1:20; powerloading=18.97*etap*sqrt(sicma)./((((CGR+(cd/clclimb))/sqrt(clclimb))*(

wingloading).^0.5));

hold on plot(wingloading,powerloading,'g','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Climb Gradient Requirement')

%Balked Landing Requirements

%inputs etap=0.8; CGR=1/30;%Radian

clclimb=clmaxl-0.2; cd=cd0+k*clclimb^2 lift_to_drag_ratio=clclimb/cd wingloading=1:0.1:20; powerloading=18.97*etap*sqrt(sicma)./((((CGR+(cd/clclimb))/sqrt(clclimb))*(

wingloading).^0.5));

HADAFTM


hold on plot(wingloading,powerloading,'b','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Balked Landing Requirement')

%Cruise Speed Requirements

%inputs Ip=0.95;%power index sicma=0.7386; wingloading=1:0.1:20; z=1/(sicma*(Ip^3)) powerloading=z.*wingloading;

hold on plot(wingloading,powerloading,'b','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Cruise Speed Requirement,75 percent power') %at take-off condition powerloading=0.75*z.*wingloading; hold on plot(wingloading,powerloading,'g','LineWidth',2) xlabel('Wing Loading(lbs/sq.ft)') ylabel('Power Loading(lbs/hp)') grid on legend('Cruise Speed Requirement,Take-off power')

HADAFTM


2.10 References

1. Roskam, J., Airplane design: Part , Preliminary Sizing of Airplanes.

2. Loftin, Jr., L.K., Subsonic Aircraft: Evolution and the Matching of Size to

Performance, NASA Reference Publication 1069, 1980.

3. Federal Aviation Regulation, FAR, Part 23.

4. A. Lennon, the Basics Aircraft Design: Published by Air Age Media Inc.2002,

2005.

HADAFTM

1404 Aircraft Design Book Selection of Engine

3 SELECTION OF ENGINE

3.1 INTRODUCTION

Selection of the propulsion system involves the following three decisions:

- Selection of the propulsion system type

- Determination of the number of engines

- Disposition of these engines

3.2 SELECTION OF THE PROPULSION SYSTEM TYPE

The following propulsion system types are available for using in the airplane:

Piston/Propeller

Turbo/Propeller

Prop fan

Inducted fan

Turbojet

Turbofan

Rocket

Ramjet

The following factors play a role in selecting the type of propulsion system to be used:

i. Required cruise speed and maximum speed

Each range of velocity requires specific propulsion system. HADAF is an ultra-light

aircraft with ⁄ cruise speed. Piston/Propeller engines are the most

efficientand popular types in this range of velocity. This part will be discussed in more

details.

ii. Maximum operating altitude

Operating altitude for HADAF aircraft is 12000ft. It is evident that for this altitude

a Piston/Propeller engine is the most suitable one. This part will be investigated in

more details in the following sections.

iii. Range economy

HADAFTM


Ultra-light aircrafts are designed for private transports so reducing cost is a mater.

Again it is seen that Piston/Propeller engine is the most efficient one according to its

low costs.

iv. Installed weight

In aviation science vehicles with lower weight are desired in order to lower the

essential fuel (costs) and at the same time exceeding flight range. The

Piston/Propeller engine meets this condition too.

v. Reliability and maintainability

Probability of failure is one the most important issues to be concerned. The

propulsion system must be safe enough. In a simple assessment, the number of moving

parts in the engine is considered as the criterion for evaluating the engine safety. The

less the number of the moving parts, the more reliable the engine would be. According

to this, Jets are the safest ones. Although Piston/propellers are not well ranked from

this point of view, their safety is getting better and better recently.

vi. Fuel amount needed

As mentioned in the installed weight part, the propulsion system must work with the

minimum amount of fuel in order to lower the aircraft weight. Also from the biological

aspect more fuels causes more pollution.

vii. Fuel cost

Generally ultra-light aircrafts are the cheapest ones. So a cheap fuel is preferred for

this sort of vehicles.

viii. Fuel availability

Most aviation fuels available for aircrafts, are kinds of petroleum spirit which are

used in engines with spark plugs (i.e. piston engines and Wankel rotaries) or fuel for

jet turbine engines which is also used in diesel aircraft engines. HADAF is an ultra-

light aircraft and must be used in small or private airports so the fuel must be

available in these kinds of places.

ix. Market demands

The propulsion system must be available and easy to repair and overhaul. Some types

of engines, like Piston/Propeller, are not fully supported in Iran.

HADAFTM


x. Timely certification

For selection of engine type, the mission specification should be checked for any

definition of the type of powerplant.Then, a preliminary speed(Mach) versus altitude

envelope should be drawn for the airplane and after that the speed-altitude envelope

of the airplane should be compared with Figure3-1 and the type of powerplant

providing the best overall match, must be chosen.

In the present design the maximum flight altitude is 18000 feet and its operating

magnitude is 12000 feet, maximum flight velocity was designed to be 51.7 meter per

second. From performance book it is known that Mach number is:

{

⁄

⁄

(3-1)

Now, referring to Figure 3-1, Piston/propeller engine is the most suitable selection

for this airplane.

HADAFTM


Figure 3-1Suitable Propulsion system indifferent velocity-altitude areas – Source: AIRPLANE DESIGN, Dr.

Jan Roskam, Part II, 124

3.3 DETERMINATION OF THE NUMBER OF ENGINES

The number of engines used in an airplane is often specified in mission specification.

The number of engines is determined by dividing the required take off power by an

integer: usually 1,2,3,4.

For selecting number of engines, the following points are necessary:

This airplane is classified in ultra-light airplanes class, so, the weight of the

airplane shouldn’t exceed the permissible range.

Reducing number of engines decreases the airplane expense.

HADAFTM


Consideration of space limitation for cockpit design affects the number of

engines.

Maximum required power is low enough to use a four stroke piston motor.

Due to these points, one engine is selected for this airplane.

3.4 DISPOSITION OF ENGINE

When the propeller is located in front of the gravity center, the installation is called

"tractor installation". When the propeller is located behind the gravity center, the

installation is referred to as "pusher installation".

Tractor installations tend to be destabilizing while pusher installations tend to be

stabilizing in both static longitudinal and directional stability.

In design of HADAF1404, tractor installation is selected for engine position. Reasons

are described below:

i. Pusher aircrafts are structurally more complicated than their equivalent

tractor types, especially when it is desired to mount the empennage behind the

rear mounted propeller. This would lead to increase in drag and loss of

empennage effectiveness.

ii. Due to the fact that center of gravity is usually located further behind on

longitudinal axis than most tractor airplanes, the pushers can be more prone to

flat spin, especially if they are loaded improperly.

iii. Normally the engine of a pusher exhausts forward of the propeller, and in this

case the exhaust may contribute to corrosion or other damage to the propeller.

This is usually minimal, and may be mainly visible in the form of soot stains on

the blades.

iv. Since the engine exhaust flows through the propellers, Propeller noise might

increase. This effect may be particularly pronounced when using turboprop

engines due to the large volume of exhaust they produce. Similarly, vibrations

may be induced by the propeller passing through the wing downwash, causing

it to move asymmetrically through air of differing energies and directions.

v. The propeller increases airflow around an air-cooled engine in the tractor

configuration, but does not provide the same benefit to an engine mounted in

the pusher configuration. Some aviation engines experience cooling problems

when used as pushers.

HADAFTM


3.5 ENGINE BRANDS

According to the power requirement which is about 90 Hp, the engines below may be

suitable for the aircraft:

i. Rotax 912 - 100 hp

Complete specification of engine is listed in the following table:

Table 3-1 Rotax 912, detailed specifications

Aircraft Engine Rotax 912 ULS or S

Displacement 1352.0 cc (82.6 cu.in.)

Bore 84.0 mm (3.31")

Stroke 61 mm (2.40")

Compression Ratio 10.5:1

Ignition Timing 4˚ up to 1000rpm above 26˚

Power Rating 95hp @ 5500 rpm continuous, 100 hp @ 5800 rpm intermittent (5min)

Maximum torque 128 N.m @ 5000rpm

Engine weight 56.6 kg (124.8 lb.)

Fuel Consumption 26 l/hr (6.7 US gal/hr) @ 5500 rpm

Fuel Premium grade leaded gas, according to DIN 1600, ONORM C 1103 EURO

SUPER ROZ 95 unleaded, according to DIN 51603, ONORM C1101

Lubrication system Dry sump lub. with trochoid pump, cam shaft driven, oil return by BLOW-BY

gas.3 liters (.8 US gal.), SAE 20W50 or SAE 30 high performance

automotive oil API, S6, Mobil 1, 15W50, NO AVIATION OIL

Cooling system Liquid-cooled cylinder heads, air cooled cylinder

Cooling liquid Conventional (mix ratio 50:50) or water free

HADAFTM


Figure 3-2 Rotax 912, detailed-sized 2D sketches

Figure 3-3 Power to Rpm curve for Rotax 912 ULS

HADAFTM


Rotax 914 – 115 hp


Table 3-2 Rotax 914, detailed specifications

Figure 3-4 Power to Rpm curve for Rotax 914 UL

Aircraft Engine Rotax® 914UL DCDI or 914F DCDI

Displacement 1211.2 cm3 (73.91 cu. In.)

Bore 79.5 mm (3.13 in.)

Stroke 4 Strokes - 61 mm (2.4 in.)

Compression Ratio 9:1

Ramp Weight 153.5 lbs (70kg) complete including exhaust, carburetor,

electronic dual ignition, electric starter and

External Alternator 40A/12V

Ignition Timing 4˚ up to 1000 RPM 1/min above 26˚/22˚

Cylinders 4 cylinders. with opposed cylinders

Power Rating 100 hp @ 5500 rpm continuous, 115hp @ 5800 rpm intermittent

Fuel Consumption at 75% power* 26 l/hr (6.87 US gal/hr)

Maximum torque 144 Nm (106 ft. lb.) @ 4900rpm

Fuel Min. MON 85 RON 95*. min AKI 91*

Oil API SF or SG

Lubrication system dry sump forced lubrication with separate 3l (.8 gal US) oil tank

Cooling system 50% BASF Glysanthin Anitcorrosion 50% Water

HADAFTM


Figure 3-5 Rotax 914, detailed-sized 2D sketches

HADAFTM


ii. Jabiru 3300 – 120 hp


Table 3-3Jabiru3300, detailed specifications

Aircraft Engine Jabiru 3300cc 120hp

Displacement 3300 cc (201.378cu.in.)

Bore 97.5 mm (3.838")

Stroke 4 Stroke - 3300cc (200 cubic inches)74 mm (2.913")

Compression Ratio 8:1

Directional Rotation of Prop Shaft Clockwise –One Central Camshaft - Pilot's view Tractor applications - Direct Propeller Drive

- 6 Bearing Crankshaft

Ramp Weight 178 lbs (81kg) complete including exhaust, carburetor, starter motor,

alternator and ignition system

Ignition Timing 25˚ BTDC fixed timing

Firing order 1 - 4 - 5 - 2 - 3 - 6

Cylinders 6 Horizontally Opposed

Power Rating 107 hp @ 2750 rpm continuous, 120 hp @ 3300 rpm intermittent

Fuel Consumption at 75% power* 26 l/hr (6.87 US gal/hr)

Fuel Mechanical Fuel Pump - AVGAS 100/130

Oil Wet Sump Lubrication - Aero shell W100 or equivalent

Oil Capacity 3.51 (3.69 quarts)

Spark Plugs NGK D9EA - Automotive

Electrical specifications Electric Starter -Integrated AC Generator

Cooling system Ram Air Cooled

Naturally Aspirated - 1 Pressure Compensating Carburetor

Over Head Valves (OHV)

HADAFTM


Figure 3-6 Power to Rpm curve for Jabiru 3300

Figure 3-7Jabiru3300, detailed-sized 2D sketches

HADAFTM


iii. Simonini 105 hp


Table 3-4Simonini, detailed specifications

Figure 3-8 Power & Torque to Rpm curve for Simonini VICTOR 2 PLUS

Aircraft Engine Simonini VICTOR 2 PLUS

Displacement 764 cc (46.62cu.in.)

Bore 80 x 2 mm (3.15")

Stroke 76 x 2 mm (2.99")

Compression Ratio 9.5 : 1

Ramp Weight 114.64 lbs (52kg) complete including exhaust, carburetor, starter motor,

alternator and ignition system

Power Rating 82.3hp @ 5400 rpm continuous, 102 hp @ 6.200 rpm intermittent

Fuel Consumption at 5400rpm 9 liters/hour ( 2.38 US gal/hr)

Electrical specifications Double Ducati electronic ignition with alternator to recharge battery in fly

Aluminum cylinders with Nikasil ceramic coating

HADAFTM


Figure 3-9 Simonini VICTOR 2 PLUS, detailed-sized 2D sketches

HADAFTM


Chart blew compares the major particular of these engines:

Table 3-5 Comparison of four engines brands

Weight

kg

Size cm

(length ×

thickness

× height)

Power

hp

RPM

max Cooling Liquid other

Simonini

(victor2plus) 52 648×410×490 102 6200rpm

low fuel burn:

2.5- 3 gph

@70%power

$4800

Rotax 912

(S or ULS)

62

100

with

Rotax

airbox&

exhaust

system

5min

5800rpm

Liquid-cooled

cylinder heads, air

cooled cylinder

50% BASF

Glysanthin

Anitcorrosion

50% Water

26 l/hr (6.7 US

gal/hr) @ 5500

rpm

Rotax

914

(F or UL)

64

+4Exhaust

System+2truss

assembly

561×540 115 5min

5800rpm

Liquid-cooled

cylinder heads, air

cooled cylinder

50% BASF

Glysanthin

Anitcorrosion

50% Water

6-7gph

@100% power

$19,370

Jabiru 3300

81

complete

including

625×380×445 120 3300 rpm

Ram Air Cooled

5.0 US Gal/Hour

(gph)

$17,500

HADAFTM


A brief comparison between engines shows that :

- Simonini has the minimum weight

- Number of air crashes due to engine failure plays an important role in engine

selection. Since Jabiru has the minimum failure, we can consider it as the safest

choice we have got.

- Engine availability is another important factor. Rotax 912 is the most popular

engine in Iran. Jabiru is in the second step.

- Another important factor in engine selection is economy. Jabiru has the

minimum cost through all.

- Rapid and convenience in overhauling also is important. Due to the popularity

Jabiru engines are of the easiest engines to overhaul.

- Fuel consumption is another important factor. Simonini is the best choice from

this aspect. Jabiru is the next one.

According to factors mentioned and other engine preferences like configuration and

weight Jabiru 3300 is selected.

The CADs of Jabiro 3300 are as following:

Figure 3-10 the CAD Models of Jabiru 3300, Modeled by HADAF group

HADAFTM


Figure 3-11 Jabiru 3300 assembled in HADAF1404

Figure 3-12Jabiru 3300 assembled in HADAF1404, TOP VIEW

Note : Engine is completely tested by the manufacturer,so it is not necessary to

analyze it again with any software.

HADAFTM


3.6 PROPELLER DESIGN

By assuming an initial value for Pb base on the database, the propeller diameter can

be derived. This initial guess can be modified in the next attemps. The equation bellow

is related in AIRPLANE-DESIGN, by Dr. Jan Roskam:

*√

+ (3-2)

Where DP is diameter of propeller, np is nomber of prop blades, Pb is power laoding

per blade and Pmax is maximum power per engine.

Some design data for homebuilt Airplanes can be found in Table 3-3. The term Pb is

assumed to be in a range of 1.7 to 2, corresponds to 2blades, 115 hp engine, though

the computed engine power is 100-110Hp.

Hence by substituting np=2 and Pb = 1.7-2 in the following equation,

and the answers below are derived :

for 110 hp:

5.91<Dp<6.41 ft (3-3)

for 100 hp :

5.6 <Dp<6.11 ft (3-4)

So final range of diameter of prop blades would be as below:

5.6<Dp<6.41 ft (3-5)

HADAFTM


Table blew lists maximum engine power, propeller diameter,number of propeller

blades and Pb which is the blade power loading for home built aircrafts:

Table 3-6 Design data for homebuilt airplanes– Source: AIRPLANE DESIGN, Dr. Jan Roskam, Part II, 129

HADAFTM


3.7 DESIGN CHART (Abstract):

Select engines

Suitable & Available PROPELLER DESIGN

Power required

HADAFTM


3 .8 References

1. Roskam, J., Airplane Design: Part II, Preliminary Configuration Design and


2. Engines catalogs

HADAFTM

1404 Aircraft Design Book The General Arrangement and…

4 The General Arrangement and Fuselage Design

4.1 INTRODUCTION

The purpose of this book is to determine a suitable and efficient configuration for our

plane and Design of cockpit and fuselage layouts to meet the mission requirements in

terms of payload and other different variables like manufacturability, market and

customer favors.

The design of the fuselage is based on payload, aerodynamic, and structural

requirements. The overall dimensions of the fuselage affect the drag through several

factors. Fuselages with smaller fineness ratios have less wetted area to enclose a given

volume, but more wetted area when the diameter and length of the cabin are fixed. The

higher Reynolds number and increased tail length generally lead to improved

aerodynamics for long, thin fuselages, at the expense of structural weight. Selection of the

best layout requires a detailed study of these trade-offs, but to start the design process,

something needs to be chosen. This is done by selecting a case not so different from

existing aircrafts with similar requirements, for which such a detailed study has been

presumably done.

The following sections are divided in two main parts: The General arrangement and

Fuselage Design.

4.2 Outline of configuration possibilities

Before general arrangement, the sketch of a new design can be drawn on paper. The

choice will be made according to the relative location of the main components including

wing, fuselage, tail surfaces and landing gear. A specific configuration is often inspired

by a trend or line of evolution, which may have its origin somewhere in the past. It may

be that previous experience with aircraft in a similar category has established a tradition,

which cannot be easily discarded. A successful choice of the configuration does not mean

that no major changes will be required as development proceeds. For a given mission it

can be several possible various solutions, each with its own particular merits.

Unfortunately, for various reasons, few examples of design evolution have been

published. It is therefore difficult to draw general conclusions from which

recommendations can deduce. The next few sections will be devoted to discussion of the

general arrangement.

The aspects followed in configuration design are considered as:

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i. Overall configuration

ii. Engine type and disposition

iii. Wing configuration

iv. Empennage configuration

v. Landing gear type and disposition

vi. Determination of the Center Of Vision (COV)

4.2.1 Overall configuration

Configuration is a term used frequently in aviation. The configuration will determine the

overall aerodynamic performance of the aircraft. When pilots use the term

“configuration”, they are usually referring to the choice of flap and gear setting. In other

words it might be described the configuration as "gear down with flaps 30 degrees" Or

pilots may say „we are clean,‟ which means gear up and flaps up. In aerodynamics,

configuration means the same thing and more. In addition to the gear and flap settings,

which are very important aerodynamically, an aerodynamicist is very interested in the

relative position of the tail and wing

The first stage of the design process is to determine the overall configuration of the

aircraft. This is called „conceptual design‟ stage. Four different configurations are

available for a designer to select which three possible configurations of them are:

i. Conventional (tail at rear)

ii. Canard (tail at front)

iii. Tailless (has no tail)

A basic Conventional configuration will define as one having the following layout

characteristics:

i. A cantilever monoplane wing

ii. Separate vertical and horizontal tail surfaces

iii. A discrete fuselage used to provide volume and continuity airframe

iv. A retractable tricycle landing gear

Conventional configuration simply means that the elevators are at the rear, in other

words behind the center of gravity and the engine is at the front but that is not a defining

characteristic of a conventional design. In a conventional design, the elevators are

usually mounted on a horizontal surface, called "horizontal stabilizer" or "horizontal tail

plane." An exception is the V-tail. All modern aircrafts have a conventional

configuration. Designers have a lot of experience with this configuration. It is a relatively

simple and cheap design. The blended wing is the most advanced design and gives the

best performance out of the four configurations. However, it is also the most expensive

design. The joined wing and three surface configurations give better performance than

the conventional design, but also cost slightly more.

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In conventional aircraft design, designer first decides on an aircraft configuration, and

then estimates its characteristics. With the inverse design methodology, it is the other

way: the characteristics will be determined first, and then the right configuration to

match the requirements will be found. The disadvantage of the inverse design is that it

requires another step - analyzing the overall performance of the resulting aircraft

configuration from its characteristics but with multi-objective design optimization,

designer can optimize the configuration by simply specifying the multiple performance

aspects they would like to improve, such as aerodynamic drag and sonic booms.

.

In conventional aircraft design, designer first decides on an aircraft configuration, and

then estimates its characteristics. With the inverse design methodology, it is the other

way: the characteristics will be determined first, and then the right configuration to

match the requirements will be found. The disadvantage of the inverse design is that it

requires another step - analyzing the overall performance of the resulting aircraft

configuration from its characteristics but with multi-objective design optimization,

designer can optimize the configuration by simply specifying the multiple performance

aspects they would like to improve, such as aerodynamic drag and sonic booms.

4.2.2 Engine type and disposition

The arrangement of engines influences the aircraft in many important ways. Safety,

structural weight, flutter, drag, control, maximum lift, propulsive efficiency,

maintainability, and aircraft growth potential are all affected. Engines may place in the

wings, on the wings, above the wings, or suspended on pylons below the wings. They may

mount on the aft fuselage, on top of the fuselage, or on the sides of the fuselage. Wherever

the nacelles are, the detailed spacing with respect to the wing, tail, fuselage, or other

nacelles is crucial.

When aircraft becomes smaller, it is difficult to place engines under the wing and still

maintains adequate wing nacelle and nacelle-ground clearances. This is one reason for

the aft-engine arrangements.

At this part according to the results obtained in the “Engine” book, which are taken

again at below, one tractor engine, is selected that logically will install and bury in the

nose of the airplane like the common used designations in the available database.

The results obtained in the “Engine” book:

i. Pusher aircrafts are structurally more complicated than equivalent tractor types,

especially because of efforts to mount the empennage behind the rear propeller. This

results in increased drag .

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ii. Due to center of gravity often being further behind on longitudinal axis than on most

tractor airplanes, the pushers can be more proneflat spin, especially if loaded

improperly.

iii. Normally the engine of a pusher, exhausts in front of the propeller and in this case the

exhaust may cause corrosion or other damage to the propeller. This is usually minimal,

and may be mainly visible in the form of soot stains on the blades.

iv. Propeller noise might increase since the engine exhaust flows through the props. This

effect may particularly pronounce when using turboprop engines due to the large

volume of exhaust they produce. Similarly, vibration may induce by the propeller,

passing through the wing’s downwash; causing it to move asymmetrically through air

of differing energies and directions.

v. The propeller increases airflow around an air-cooled engine in the tractor

configuration, but does not provide this same benefit to an engine mounted in the

pusher configuration. Some aviation engines experienced cooling problems when used

as pushers.

4.2.3 Wing configuration

4.2.3.1 Wing location:

It will be obvious that the location of the wing relative to the fuselage is to large extent

determined by the operational requirements. Although the aerodynamic and structural

differences are important, they can be only decided as factors when the choice between

high, low and mid wing is not dictated by considerations of maximum operation

flexibility.

Forewing location there are several choices shown in figures below but three of them are

mostly known as conventional and operative sorts, which will be discussed in next

paragraphs.

Figure 4-1 Location of the wing

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i. Mid wing:

Figure 4-2Example of mid-wing plane

This layout is used when minimum drag in high-speed flight is of paramount importance

Thus the divergence of the airflow over the wing root at high angles of attack is

minimized. For such reasons many mid-wing layouts are found in fighter and trainer

aircrafts. It is not generally feasible to adopt such a scheme for transport aircraft, and a

few mid-wing monoplanes are found in this category. In this case, the cabin floor, which

is located just above the wing center section, is the position relatively high in the fuselage

cross section.

Also with this layout, it is difficult to avoid considerable shift of the center of gravity for

different loading conditions unless serious loading restriction are accepted. Since in a

two-seat airplane, minimum drag is not as important as high-speed flight and mid wing

use for a high-speed flight, it is better to discuss low or high wing structures.

Also in this design the maximum area in cabin is needed which mid-wing airplanes are

opposed to this demand.

ii. High wing & low wing (comparison):

Figure 4-3Example of high-wing plane

Figure 4-4Example of low-wing plane

In the case of a low-wing aircraft of comparable size with high wing, the main deck floor

is higher above the apron. This makes such an aircraft dependent on special loading and

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boarding equipment, which is unacceptable for military aircraft. But in the case of most

passenger aircraft the height of the cabin floor above the ground is of less importance.

In smaller high-wing propeller aircrafts, it may be possible to retract the main gear into

the engine nacelles or in tail booms but in the case of very large aircraft, doing so would

make it too tall and too heavy. This will unavoidably lead to mounting the gear to the

fuselage, but strengthening of the fuselage structure required for the transmission of the

landing impact loads will result in a weight increase. This is only in offset by the saving

in weight in comparison with a low-wing design, due to the shorter landing gear struts.

In this design, the target is a fixed landing gear -which will discuss in landing gear book-

so application of low wing will cause difficulties in landing gear design. In addition,

braced-wing monoplanes are generally high-wing designs, which cause little

interference.

In the case of a STOL airplane, close proximity of the wing to the ground in takeoff and

landing may cause pronounced and generally undesirable ground effect. Moreover, if a

low wing was adopted, the required clearance of the large, fully deflected trailing edge

flaps and -in the case of propeller-driven STOL aircraft- large propellers, would entail a

very tall and heavy landing gear. In this case, a high-wing design generally has more to

recommend it.

4.2.3.2 Effect of the wing location on the general arrangement:

According to descriptions above, design can continue by paying attention to some points:

i. Interior arrangement

ii. Safety

iii. Performance & flying qualities

iv. Structural aspects

In a high wing aircraft, the fuselage section under the floor is generally flattened .Also

there is a wide area in cabin.

From another view, since the impact is not too heavy, damage and fire in high wing

aircraft will be limited so these kinds of aircrafts have more safety than others do.

The principal difference between characteristics of high and low wing layouts during

takeoff and landing is the ground effect, which decreases by increasing the wing height

relative to ground. Ground effect will generally cause a reduction in vortex-induced drag,

resulting in a decreased distance. In a high wing layout, the minimum ground effect

happens.

Also high wing structure can bring a high stability for performance of target.

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Decisions:

According to descriptions above, design in this part can complete by choosing the

location of the wing. After a comparison between high and low wing some results

obtained:

1. In a high-wing layout, the minimum ground effects happened.

2. In a high-wing aircraft, the fuselage section bellow the floor is generally flattened.

3. In a high-wing aircraft, the cabin is more spacious in comparison with other types of

wings.

4. High wing structure causes high static stability for targeted performance.

5. In high-wing aircrafts, structure design has a simple way in comparison with other types

of wings.

6. High-wing aircrafts have more safety than others do.

4.2.3.3 RESULTS:

Because of the reasons above, project limits and necessity of simple design and

manufacturing, it seems that a high wing layout is the best choice for a two seat aircraft

like Hadaf1404.

4.2.4 Empennage configuration

Figure 4-5Tail types

The tail surfaces design is strongly dependent on the overall airplane details. Location of

the surfaces is affected by the engine disposition, especially in propeller cases and the

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arrangement of the wing with tail. The fact that how tail installation is done, affects on

the empennage structural design and therefore on overall structure of the airplane. Since

we have already determined our wing and engine location, now we can survey their

effects on the tail and prepare a logical discussion for tail architecture. At first, all of the

principal configurations will be included and then the best choice will be chosen. As it is

shown in the figure above, all tails can be classified in three groups. The first group

consists of crucifix tail and T-tail in which the layout includes a vertical and horizontal

stabilizer. In T-tail model, we should pay attention that flutter may happen and from

structural sight it needs a strengthened fin in root, since there is a large bending moment

caused by stabilizer. The crucifix tail is stiffer in these cases but T-tail is more reliable in

Spin. The operative area of the T-tail during spin is equal with total fin area but in

crucifix, the wake caused by stabilizer should be mounted.

Figure 4-6Effect of the rudder during a spin

A geometrical approach to analyze the wake of stabilizers is to draw a tangent line with

60 degrees inclination. The desired tail to pass this analysis is one that this line sweeps

less than one third of the fin efficient area.

The twin vertical tail is a beneficial theme in structural design. Since the center of

pressure of the fin is lowered during the deflections of the rudder. It causes less moment

on the tail root but the operation of this theme, is not desired during spin.

The V-tail is not a popular and common tail. Since the moving surfaces should act as both

of the rudder and elevator, it needs a complicated control system. Since there is not

engine efflux on the tail, according to engine disposition, it acts more efficient in tail

plane design. The distance of the engine and tail surfaces is the highest amount in

HADAF configuration than ones in database.

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Hadaf is in the group of heavy UL aircrafts and this limit causes to choose the best tail in

performance and structural weight. In order to provide the best performance, we should

choose one of the configurations from A-1, A-2 and A-3. T-tails need more advance

structures, which seems to employ additional weight. So A-1 and A-2 should be surveyed.

Since in configuration A-2 the elevators will receive more efflux from the wing

downwash, it is strongly suggested to use A-1, which has more reliability in this case. In

order to provide security standards of the tail during spin, elevators must move either

back/down or even both such in Remos GX.

4.2.5 1.5. Landing gear type and disposition

Since we have a heavy plane in UL class, choosing an appropriate landing gear is very

important. This importance will be shown when the brake system is going to install and

also in landing process. The requirements that should be met by landing gears are to

strengthen properly over the loads exerted on it while the landing impulses happen; also

transferring less impulses into cockpit.

Various configurations for undercarriage have been adopted up until now, but each of

them was designed for special purposes. Only three of these need be discussed in the

present context.

i. Tandem undercarriage

ii. Tail-wheel undercarriage

iii. Nose-wheel undercarriage

4.2.5.1 Tandem undercarriage

Here the main wheels are arranged practically in the plane of symmetry of the aircraft

and the front and the rear wheels landing impact forces of the same magnitude. use of the

tandem gear is justified when much emphasis has to be placed on the following

advantages:

- Both main legs are placed at nearly equal distances ahead of and behind the center of

gravity, thus locally creating space for payload close to it.

- The wheels may be retracted inside the fuselage without interrupting the wing structure.

The increase if any in fuselage weight will depend on other factors.

Against these we have to set following disadvantages:

- Outrigger wheels will be required to stabilize the aircraft on the ground and these may

increase the all-up weight by approximately 1#.however by using two pairs of main legs

instead of single ones, a certain amount of track may be obtained, resulting in a reduction

of the load on the outriggers.

- The pilots must be carefully maintain the proper touchdown attitude in order to avoid

overstraining the outriggers. It may sometimes be a possible to locate the rear legs close

to the center of gravity of the aircraft, and so reduce this disadvantage, but that means

losing the opportunity to have an unobstructed space.

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- A large tail download is required to rotate the aircraft.it will therefore be desirable to

choose the attitude of the aircraft at rest so that it will fly itself off, but this may lead

either to an increase in drag during the takeoff roll or to a high liftoff speed.

The arguments against the tandem gear are of such a nature that its adoption should only

be considered when no other solution meets the case.

4.2.5.2 Tail-wheel undercarriage

Figure 4-7Tail-wheel undercarriage

Though this type of undercarriage was in aerial use during the first three decades of

aviation, it must now regard as obsolete for most designs. Its advantages could

nevertheless be mentioned.

i. The tail-wheel is small, light and simple design.

ii. The location of the main gear legs makes attachment to the wings an easy matter.

iii. A three-point landing may bring the aircraft to stall condition. The aerodynamic drag

will provide a force, which is particularly in need when the airfield is unsuitable for

full application of brakes (e.g. wet grass).

iv. When brakes are applied, the vertical load on the main gear will increase, thereby

reducing the risk of skidding.

The reason why the tail-wheel undercarriage has been almost completely superseded by

the nose-wheel or tricycle gear is that it also possesses the following drawbacks:

i. Violent braking tends to tip aircraft onto its nose.

ii. The braking force acts ahead of the center of gravity and thus has a destabilizing

effect when the aircraft is moving at an angle of yaw relative to its track. This may

cause a ground loop.

iii. In a two-point landing, a tail-down moment will be created by the impact force on

the main landing gear, resulting in an increase in lift, which makes the aircraft

bounce.

iv. The attitude of the wing makes taxying difficult in a strong wind.

v. In the case of transport aircraft the inclined cabin floor will be uncomfortable for

the passengers and inconvenient for loading and unloading.

vi. In the tail-down attitude the inclination of the fuselage will limit the pilot’s view

over the nose of the aircraft.

vii. During the initial take-off run drag is high until the tail can be raised.

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viii. In some designs it is possible to circumvent some of these disadvantages at least

partly. Interconnection of the tail-wheel and the rudder control the aircraft on the

ground.

4.2.5.3 Nose-wheel undercarriage

Figure 4-8Nose-wheel undercarriage

The merits and drawbacks of the nose-wheel gear are roughly the opposite of those of the

tail-wheel type. The principal advantages are:

i. The braking forces act behind the center of gravity and have a stabilizing effect, thus

enabling the pilot to make full use of the brakes.

ii. With the aircraft on the fuselage and consequently the cabin floor are practically level.

iii. The pilot’s view is good.

iv. The nose-wheel is a safeguard against the aircraft turning over and so protects the

propeller (s) when used.

v. During the initial part of takeoff the drag is low.

vi. In a two-point landing, the main gear creates a nose-down pitching moment. The steady

increase in landing speeds of modern aircraft has accentuated these advantages, so that

they carry more weight than the following these advantages:

- The nose unit must take 20 to 30thof the aircraft’s weight in a steady braked

condition and it is therefore relatively heavy.

- The landing gear will probably have to be fitted at a location where special

structural provisions will be required. In the case of a retractable nose-gear on light

aircraft, it may also prove difficult to find stowage space inside the external

contours of aircraft.

Although there is still a measure of choice during the preliminary design stage, this

constitutes one of the most difficult problems to be solved. Summing up, we may state that

the nose-wheel undercarriage has gained favor because it greatly facilitates the landing

maneuver and enables the brakes to be used more efficiently.

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4.2.6 Determination of the Center Of Vision (COV)

Regarding to the development of the UL airplanes in last two decades, after when

Roskam method was published, a database is provided from the seat modes and

dimensions in order to determine the pilot‟s Center of Vision (COV).

The Database of seat dimensions is provided below:

Table 4-1database of seat dimension

1 2 3 4 α Θ

Land Africa 65 38 46 14 16 8

Jabiru 47 38 45 12 18 20

Savannah 54 42 42 10 30 7

Euro Fox 47 40 44 12 21 12

Remos GX 45 35 46 8 24 21

The dimensions announced here are equivalent with some of stated notations in the figure

below. Number 1 is the vertical component of R, number 2 is horizontal component of P,

number 3 is the width of the seat and number 4 is equivalent of C. The angles indicate the

position of the seat in Cartesian coordination. (α) is equal with D and (θ) is equal with E

– 90.

Figure 4-9standard seat dimensions in Roskam methods

According to FAR 25 standards, the distance between COV and the seat joint must be

utmost 80 cm and it is assumed to install the pilot seat in at least 20 cm above the cabin.

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Using dimensions measured in the following tabulation, the location of COV was

determined.

Table 4-2Roskam to locate COV

A B C D E F G H I J K L M N O P Q R

37 30.25 5 21 101 29.75 10.0 14.5 19 6 9 11.5 36 5 9.25 15 7 25

39 30.758 5 19 101 30.25 9.75 13.75 19 6 9 13.75 35 5 9.25 15 7 25

41 31.5 5 16 101 31.0 9.75 13.5 19 6 9 15.5 34.5 5 9.25 15 7 25

43 31.75 5 16 101 31.25 10.0 13.25 19 6 9 17.5 34.5 5 9.25 15 7 25

Figure 4-10 Definition of joint angle

According to the recent researches, the standard values for human skeleton position

angles is obtained as shown in table-3 which discusses the angles of figure-9. Comparing

these values with ones we obtained from practical measurement and ones announced by

Roskam, leads to a process to choose definite values for these angles and length. Now we

can calculate the COV of the pilot regarding to Hadaf primary configuration and

dimensions, which will be modified later. These body geometrical values are also used to

locate the definite position of the stick and rudder pedals.

Using the data above and calculating for a standard body gesture of seated pilot, the

COV of Hadaf locates 112 cm above the cabin and 110 cm cross the firewall.

Table 4-3Comfort range and most comfort value of joint angles

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Figure 4-11COV location

Location of the COV and also the downward vision angle are strongly dependent on the

cabin interior configuration and also the nose style which should be surveyed in CAD

files. The minimum angle of vision under the horizontal line of pilot's eye should be 15

degrees and in order to meet visibility requirements during landing phase, this session

must be concentrated. This angle for our plane is 15.97 degrees.

4.3 Outline of fuselage design

The preliminary general arrangement of the aircraft is closely tied up with the fuselage,

the main dimensions of which should be laid down in some detail. In fact, the fuselage

represents such an important item in the total concept that its design might well be

started before the overall configuration is settled.

i. The fuselage is the most suitable part for housing the cockpit, the most function at

location generally being in the nose.

ii. The fuselage may be regarded as the central member structural member to which

the other main parts are joined(wings, tail unit and in some cases the engines) on

the one hand, and the aircraft on the other. In some aircraft a number of these

duties are assigned to tail booms.

iii. Most of the aircraft systems are generally housed in the fuselage, which

sometimes also carries the engines, fuel and/ or the retractable undercarriage.

Many of the requirements laid down in relation to the fuselage limit the designer‟s range

of choice.

The aspects followed in this part are considered as:

1. The selection of cabin cross-section dimensions.

2. Determination of fuselage length and shape.

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4.3.1 Cross-Section Design

It is often reasonable to start the fuselage layout with a specification of the cross-section:

its shape and dimensions.

Most fuselage cross-sections are relatively circular in shape. This is done for two

reasons:

i. By eliminating corners, the flow will not separate at moderate angles of attack or sideslip.

ii. When the fuselage is pressurized, a circular fuselage can resist the loads with tension

stresses, rather than the more severe bending loads that arise on non-circular shapes.

Here based on knowledge of unpressurized cabins and as dictated by cost constraints and

volumetric efficiency, our fuselage has a basic relatively rectangular section with

eliminating corners that is shown at figure4-12.

Figure 4-12Main cross-section

4.3.2 Fuselage Diameter

The fuselage consists of three basic sections: the engine section, the cabin section, and

the sheet-metal tail cone section.

Figure 4-13.Three basic sections of fuselage

The first question to answer at this design stage is related to seating type, side-by-side or

tandem seating; here based on knowledge of existing aircraft data, we chose side-by-side

type. Then the dimensions are set so that pilot and standard cargo containers may be

accommodated. Typical dimensions for two-seat aircraft cabin, which our initial layout

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based on from FAR 23 standards was taken at “Determination of the Center Of Vision”

part.

4.3.3 The sheet-metal tail cone section

The fuselage cone is normally a smooth transition from the maximum fuselage cross

section to the end of the fuselage. When the „fineness ratio‟ of this cone is too low, there

will be a large base drag penalty although the fuselage weight may be reduced. When the

„fineness ratio‟ of this cone is too large, there will be a large friction drag penalty as well

as a large weight penalty. It is obvious that a long fuselage cone tends to increase the tail

moment arm thereby reducing required tail area and vice versa.

The decision on the fuselage cone fineness ratio is there for one that involves a number of

trade-offs.

4.3.3.1 Database discussion

In books 1 and 2 a database of similar airplanes was used to estimate some of our

specifications. In this book also we do it, but we should note that database planes were

classified in both high wing and low wing configurations. In this book we should survey

ones correspond our configuration‟s outline. So we have to filter our database again to

contain the appropriate cases. In this database new parameters of airplanes should be

determined. These parameters are LFC, LF, DFC and θFC as discussed in reference book.

These data are prepared in tabulation as below:

Table 4-4LFC, LF, DFC andθFC parameters of the database

Name LF LFC θFC DF LF/DF LFC/DF

Parrot 6.4 4.6 9.6 1.18 5.4 3.9

F99Rambo 5.1 3.1 11.7 1.1 4.7 2.8

Tecnam 6.4 3.6 5.1 1.275 5.1 2.8

CTSW 6.2 4.1 6.0 1.324 4.6 3.1

Remos 6.4 4.6 8.4 1.25 5.1 3.7

According to this database, we can estimate the similar values for our plane. Table 2

shows geometric fuselage parameters used on our airplane HADAF1404:

Table 4-5.HADAF1404 Geometric Fuselage Parameters

LF LFC θFC DF LF/DF LFC/DF

6.1 4.0 8.1 1.2 5.0 3.2

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These parameters orientation are shown in figure4-14:

Figure 4-14LFC, LF, DFC and θFC parameters on the aircraft

4.3.4 Fuselage Shape

The fuselage shape must be such that separation and shock waves are avoided when

possible. This requires that the nose and tail cone fineness ratios be sufficiently large so

that excessive flow accelerations are avoided. Figure 4-15 shows the limit on nose

fineness ratio set by the requirement for low wave drag on the nose.

For our plane, HADAF1404

:

Diameter = 1198.6mm

Nose Length = 949mm

Nose Fineness = 0.79

From the figure:

Drag Divergence Mach Number = 0.73

Figure 4-15effect of nose fineness on drag divergence Mach number

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The tail cone taper is chosen based on similar considerations and generally falls in the

range of1.8 to 2.0.

Figure 4-16 after body drag of a fuselage tail, when added to a cylindrical shape

For our plane:

Diameter = 1198.6mm

Tail Cone Length = 2846mm

Tailcone Fineness = 2.37

From the figure:

Drag ratio = 0.25

A frequently used value for the length/diameter ratio is 1.5 to 2. A lower value may be

used on fighters provided that this lightens the door and door support structure to such

an extent that it outweighs the extra drag.

Several rules result from these analyses: The transition from nose to constant section,

and constant section to tail cone should be smooth - free of discontinuities in slope

(kinks). The tail cone slopes should resemble those shown in the examples. That is, the

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slope must change smoothly and the trailing edge should not be blunt. The closure angle

near the aft end should not be too large (half angle less than 14°-20°).

4.3.5 Hadaf configuration

Using data obtained from database and also some comparison between other cases out of

database, a graphical design was drawn. The properties of this configuration, called

Hadaf, are set with mission profile and also requirements that obtained in book 2. Some

individual values in parameters discussed are because of mission defined for Hadaf that

requires special design in cockpit. The ability of containing three passengers needs more

spacious cockpit, which affects the overall configuration. The figures of side and front

view of Hadaf are presented in below:

Figure 4-17

Figure 4-18

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4.4 Designing diagram

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4.5 Appendix

First render

Main cross section

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Other cross sections:

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Engine section

Cabin section

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The sheet-metal tail cone section

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The sheet-metal tail cone section

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4.6 References:

5. E.Toreenbeek: “syntheses of subsonic airplane design”, Delft university press,

Rotterdam, 1976.

6. J. Roskam: “Airplane design”, 1985.

7. Kroo: “Engineering aircraft design”, Stanford University press, 2001.

8. “Designworkbook”,airline.http://www.futureflight.org/downloads/designlogbo

ok.pdf

9. http://selair.selkirk.ca/Training/Aerodynamics/configurations.html.

10. http://www.jaxa.jp/article/special/aviation/oonuki01_e.html

http://www.futureflight.org/downloads/designlogbook.pdf

http://www.futureflight.org/downloads/designlogbook.pdf

http://selair.selkirk.ca/Training/Aerodynamics/configurations.html

http://www.jaxa.jp/article/special/aviation/oonuki01_e.html

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1404 Aircraft Design Book Wing Sizing

5 WING SIZING

5.1 INTRODUCTION

The purpose of this section is designing the aircraft wings in the way that not only

bear the aircraft weight (which is the main duty of the wing for sure), but also have

the desired aerodynamic specifications. The wings must produce enough lift while not

generating so much drag. In order to gain this, the lift to drag (L/D) ratio has to

maximize.

When it comes to design a wing for an aircraft the designer must decide about two

important parameters.

1. The wing section profile (airfoil)

2. The wing planform (wing layout)

All the design parameters are hidden in the two above. So the foregoing book is being

presented, by simply fully define the two parameter mentioned above. In the next

paragraph the more detailed design parameters will be introduced.

5.1.1 DECIDE 1DECIDE ON THE OVERAL WING/FUSELAGE ARRANGMENT

As a result of previous books it is known that the overall configuration of the airplane

is conventional which means that the tail is situated on the aft of the aircraft and the

wing configuration is high. High wing configuration means the wings are connected

to the fuselage on the top. Although the high wing configuration does not have the

least interface drag, but its superior lateral stability was the main reason that it was

selected as the default configuration for wings. Moreover the high wing configuration

helps the passengers to enjoy the broader landscape with more ease.

5.2 MORE DETAIL DESIGN PARAMETER

The aircraft wing designing is finished when the 10 items below are known, for

certain. All the parameters below are defining the planform design except the 4th

item

which is about the wing cross-section.

1. Size (s)

2. Aspect ratio (A)

3. Sweep angle

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4. Airfoil profile

5. Thickness ratio (t/c)

6. Taper ratio ( )

7. Incident angle

8. Twist angle

9. Dihedral angle ( )

10. Lateral control surface size and layout

There is no need to note that items 1 and 2 are already known from the previous

books.(i.e. performance) The definition of these 10 parameters exists almost in every

aerodynamic books, but here some of them will introduce in brief.(The reader is

referred to: '' Airplane Design. By Dr. Jan Roskam'' which is one of the greatest books

in this field.)

Aspect ratio:

The ratio of span and the average chord is called, "aspect ratio". In this definition the

average chord is the ratio of wing area and the span. As this parameter increases the

airplane glides in the sky more easily and the sliding angle decreases. But increasing

this parameter is usually bounded by structural problems and also the induced drag

that is the drag due to the lift. For a rectangular wing, there will be:

Sweep angle

The sweep angle( ) is usually measured as the angle between the line of the 25%

chord and a perpendicular to the root chord .sweep angles of the leading edge and of

the trailing edge are also presented with other parameters, since they are of interest

for many applications. The sweep of a wing causes definite changes in the maximum

lift, in the stall characteristics, and in the effects of compressibility.

Taper ratio:

Considering the wing planform to have straight lines for the leading and trailing

edges, the taper ratio, is the ratio of the tip chord to the root chord.

The taper ratio affects the lift distribution and the structural weigh of the wing. A

rectangular wing has the taper ratio 1.00 while a pointed tip delta wing has a taper

ratio of 0.0.

Dihedral angle:

The dihedral angle is the angle between a horizontal plane containing the root chord

and a plane midway between the upper and lower surfaces of the wing .if the wing lies

below the horizontal plane, it is termed an anhedral angle. Generally, the dihedral

angle affects the lateral stability characteristics of the airplane.

Below the subject with a database of the parameters mentioned above in the ultra-

light airplanes is brought up.

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Table 5-1Database of ultra-light wing parameters

Type

Dihedral

Angle

deg

Incidence

Angle

deg

Aspect

Ratio(AR)

Sweep

Angle

deg

Taper

Ratio

Max Speed

Kts

Wing Type

PIK-21 Duruble 0 0 3.8 0 1 NA ctl/low

RD-03C PIEL 6.5 3 7 0 0.51 182 ctl/mid

CP-750 5.7 4.2 5.9 0 0.55 183 ctl/low

CP-90 POTTIER 5.7 3 5.4 0 0.44 171 ctl/low

P-50R 4.4 NA 5.1 2 0.54 167 ctl/low

P-70S O-O 0 2 4.8 0 1 129 ctl/mid

Aerosport 2.5 NA 5.7 0 1 76 ctl/low

Micro-Imp 4 4 4.7 0 1 260 ctl/high

SA-III Sequoia 3 1.5 5.7 0 1 165 ctl/low

300 Ord Hume 3 3.5/1.5 6.9 0 0.55 243 ctl/low

OH-4B Procter 3 5 5 1 95 brcd/parasol

Petrel 5 0 6.6 0 1 113 ctl/low

Bede BD-s 0 3 3.9 0 1 238 ctl/low

5.3 AIRFOIL PROFILE DESIGN

As the only part of a conventional airplane that produces lift are its wings. So in order

to have this plane safely in sky the wings must generate lift as much as the weight.

The weight is known, so the required lift is known. Therefore the Cl for the airfoil is

known for every velocity magnitude according to the popular relation shown below:

It is clear that the critical state for the airfoil design is when the airplane is moving in

the lowest speed in such condition the airfoil must produce enough lift to suppress the

weight. But since the velocity has its minimum value the denominator would be very

small and the Cl would be high. This is the key to find out how much Cl does the plane

need at the minimum speed? (i.e. the Cl max)

For the present case this figure was estimated about 1.5 which can be produced by

variety kinds of airfoil. So which one should be selected?

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The answer is, an airfoil which produces appropriate lift at zero angle of attack is

desired. In other words if the incidence angle considered equal to zero as it is in the

present case, the airplane in the cruise condition, must produce enough lift to

overcome the weight. This is very important key point, because if the lift is more than

weight in the cruise condition the airplane would ascend and the cruise speed

decreases and also if the lift is less than weight, then the airplane would descend and

the cruise speed would increase. So the Cl of the airfoil at zero angle of attack must be

around special distinct value. This will confine the range of the airfoil which can be

selected for HADAF aircraft.

Moreover than mentioned above it need to maximize the (L/D) ratio, while not letting

the drag to expand. To do so, a new airfoil ''HADAF 1404'' by the airfoil shape

optimization method was designed. The numerical investigations show that this airfoil

has great aerodynamic behaviors but the results of the numerical analysis must be

verified with the experiment. Having finished the experimental tests the airfoil will be

introduced in details.

Design team tried to find the best airfoil from the existing databases that meets the

criteria mentioned above. Firstly NACA 4415 was tried. But as you can see in the

figure 3-1 it did not produce enough lift in the cruise condition. So, the NACA 5413,

5314, GOE 533 and NACA 5215-62was tried as next efforts. The results are given in

figure 5-1 in brief. The numerical investigations were done using commercial CFD

package, FLUENT. Some special cases were solved in order to verify our CFD

results.

Figure 5-1Sample airfoils wing lift and desired comparison

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As can be seen in figure 5-1 NACA 5215-62 (2) is in good agreement with desired

cruise condition. It is good to say the suffix (2) is stand for some slight planform

modifications used to modify the cruise lift conditions. In other words the differences

between the last two columns in figure 5-1 are their area and planform configuration.

Although NACA 5215-62 (2) is the best wing section, matching the desired conditions,

there is another important point here in finalizing the decision. The airplane may also

cruise at some lower speeds than 185km/hr (say 160 km/h). So a wing that produces

more lift in cruise condition has to be selected. This slightly more lift force, can be

trimmed by the pilot through the flight. In this way the aircraft can cruise gently in a

range of speeds between 155km/hr and 185km/hr.

According to what discussed above, the best wing section matching our HADAF 1404

would be NACA 5314.

The aerodynamic characteristics of some of the wing sections mentioned above

presented in figure3-2. These data were gotten from the Design foil software package.

Figure 5-2Cl to AOA of sample airfoils at Design foil

The result of the design-foil software was verified with the FLUENT software. The

grid that was used in this analysis was illustrated in figure3-3.

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

-10 -5 0 5 10 15 20

NACA 5314

Naca 4415

NACA 5415

NACA 5215-62

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The solver was selected to be pressure base due to the low subsonic cruise speed and

the SIMPLE scheme was chosen to discretize the governing equations. The Cl diagram

was the criteria for detecting the convergence. The aerodynamic characteristic of the

airfoil has been illustrated in figure3-4.

It is remarkable that Spalart-Almaras was used, as the viscous turbulent model due to

its accuracy and reliability.

As can be seen in figure3-4, the Cl for the airfoil was estimated 0.54 in FLUENT

software, while it was previously calculated 0.62 by design-foil software. This

difference is obviously because of the lack of enough accuracy in design-foil software.

In this commercial software the program usually estimates the aerodynamic

characteristics by some quick method like panel method, which are good to give one a

sense about the general characteristics but they can‟t be used for predicting the exact

values.

Figure 5-3Analysis used grid of airfoil

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Table 5-2 The aerodynamic characteristic of the airfoil

5.4 WING PLANFORM DESIGN

In this part, the 9 other parameters mentioned in part 2 will introduce to reader.

The first parameter is the size or wing surface which is known from the previous parts

and is 13.074 (sq.m). As it is discussed in aircraft performance analysis, the larger

wings, or lower wing loading, help the aircraft to have shorter take-off / landing field

length. Approximate equations, suggested by FAR 23 & 25, claim that the take-off /

landing field length is directly proportional to the wing loading. As wing loading is

increased the take-off and landing length would be increased. On the other hand a

high wing loading helps the aircraft to ride through the turbulences with more ease

and it also makes the aircraft lighter. The previous discussion can be summarized in

the following table:

Table 5-3 Comparison of performance for wing loading

High W/S Low W/S

Stall Speed High Low

Field length (landing and take of) Long Short

Max. L/D High Low

Ride quality in turbulence Good Poor

weight Low High

The next items is the aspect ratio which was believed to be 7.65.Again it emphasizes

that these two important parameters are drawn out from the configuration book,

where the matching diagram was developed.

The next items are not known yet and need to speak about them in more details.

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5.4.1 SWEEPANGLE

Generally speaking, these are the options for the sweep angle:

Zero or negligible sweep angle

Aft sweep ( also called : positive sweep)

Forward sweep (also called: negative sweep)

Variable sweep

Oblique sweep

The last two alternatives are suitable for the flight missions involve cruising in broad

range of the Mach numbers and for high "g" maneuvering. In the variable sweep

configuration a big weight penalty exists, due to the high weight wing pivot structures

which are necessary for altering the sweep angle.

The sweep angle helps us mostly in high Mach numbers especially in the transonic

regimes. But how?

When the wing has a sweep angle, the free stream velocity coming toward the leading

edge, decomposes into two components. One is tangent to the wing and the other is

normal. The normal component is the velocity vector which determines the critical

Mach number. So, the critical Mach number increases. This means that the airplane

can navigate faster than the speed of sound while the flow regime on the wings is

subsonic. The tangential component flows along the wing span. This flow is called the

span-wise flow.

Taking a glance at the database, it is obvious that nearly no U.L. aircraft has been

designed with the swept back or forward wings. This is mainly due to the low Mach

number in their flight regime. Considering that swept wings helps us in the mach

numbers greater than 0.5, especially in the transonic flows; it was concluded that a

swept wing for an U.L. with the cruise Mach 0.1 is somehow wasting money and

energy.

In the case of the sweep aft wings, in low Reynolds numbers(as in this case) the

boundary layer growth is enhanced and this growing boundary layer tends to

approach the wing tip in the aft swept wings because the span-wise flow, moves from

root to chord in sweep back wings. The wing tip is a crucial section and the ailerons

are positioned there. So boundary layer growth in this zone can be very harmful and

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the separation in this zone means losing of the control on our ailerons. However

these things won't happen in higher Reynolds numbers.

Although the sweep forward wings do not have the problem above, but other problems

like: manufacturing difficulties, stability problems, fluttering and etc. have left this

configuration on the papers. But now the developed manufacturing technologies and

also integrated control systems tempt the designers to reconsider this configuration

again.

In HADAF 1404, the sweep angle is set to zero degree.

5.4.2 THICKNESS RATIO (t/c)

As the wing section is defined the t/c ratio is known definitely. This t/c is very

important in determining the drag and lift coefficients of the airfoil. It's 14% for

NACA 5314.

This parameter is also responsible in determining the critical Mach number.

Thickness ratio and the sweep angle are usually gathered in some charts to determine

the critical Mach number. See figures 5.5 and 5.4.

As you see in the figure, the thickness ratio has a diverse effect on the critical Mach

number and by increasing the t/c the critical Mach number decreases. While the

critical Mach number increases with the increase of the sweep angle, which verifies

our previous statements about the effect of the sweep angle.

As the t/c increases the increase in the speed of the free stream velocity during the

pass over the upper edge of the airfoil increases, causing the critical mach number to

decrease.

Figure 5-4 (“AirplaneDesign”,JohnRoskam,PART2,page150)

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Figure 5-5 (“AirplaneDesign”,JohnRoskam,PART2, page 150)

5.4.3 TAPER RATIO

Taper ratio is the ratio of the wing tip chord over the wing root chord as illustrated

below.

Figure 5-6Tip and root cord layout

CordRoot

CordTip (5.2)

The effect of the taper ratio is smoothing the variations of the lift along the span. The

elliptic wing as an ideal wing can produce a constant lift along the span, while a

rectangular wing loses lift along the span in the way that at the tip does not produce

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any lift. Having the above explanation in mind it can be deduced that the more

designer approach to the elliptic layout, the more constant Cl along the span.

According to the conclusion discussed above, the design team decided to arrange the

trim line in a manner to get near to an elliptic wing. Below the method is illustrated.

Figure 5-7The Wing Layout

Finally the was chosen for HADAF1404.

5.4.4 TWIST ANGLE

A good design is a compromise of the optimums and the cost. Here using a twist angle

will cost too much while the influence is not that much. Again looking at the database

will ensure us that this decision is logical.

In this case twist angle is set to zero.

5.4.5 INCIDENT ANGLE

When the aircraft is moving on the ground the wings make an angle with the horizon

that is the incident angle. When the aircraft is flying at the cruise speed and it is

horizontal, the wings must create a special Cl to overcome the weight. But this Cl is

produced by the wing at special angle of attack. That angle is called the incident

angle. For the present case this angle is 0 degree. Because the airfoil is designed in

the way that it produces enough lift even in 0 degree.

Trim lines are shown by dotted lines

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5.4.6 DIHEDRAL ANGLE

The dihedral angle is shown in the picture below. This angle will have a good

influence in the aircraft roll equilibrium. Generally speaking, this angle is essential to

guarantee the roll stability in low wing aircrafts, but it is also used in high-wing

configurations as a stability booster. In these cases the stability increases in expense

of increase in rolling radius. In the next picture you can see how the dihedral angle

enhances the aircraft equilibrium.

Figure 5-8Dihedral Angle

Figure 5-9Exaggerated Dihedral Angle

In this project, HADAF 1404, the dihedral angle was set to be 2 degree.

5.4.7 Wing test:

At this point the wing cross-section and its planform have designed. But the question

is: How reliable the wings are? Do they really keep the aircraft in the air? How much

drag do they produce?

To answer the questions above, there were two choices: one is to test the wing

prototype in a wind tunnel and the second solution is numerical modeling. The

numerical investigation method due to its inexpensiveness and the ease of it was

chosen. The wings were modeled in a rectangular domain and were exported into the

FLUENT software. The grid is illustrated in figure5.10.

In the figure4-7-1you can see the section view of the grid in the Z direction and

similarly in figure 4-7-2 a section view in the X direction was shown. In figure4-7-3

the whole geometry and the grid was shown.

The lift forces cancel each

other

The lift forces on the right wing keen to

cancel the roll

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Next, the mesh independency was checked. (I.e. to assure that the results won't change

drastically if the mesh was further finer). This step in grid generation is so important,

that without this step you can't be hopeful to gain logical results from C.F.D. soft-

wares.

Figure 5-10section view of the grid in the Z direction

Figure 5-11section view of grid in X direction

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Figure 5-12whole geometry and the grid view of wing

5.4.7.1 Numerical Investigations:

As explained above, the geometry and the grid was created using Gambit software,

and exported to the commercial C.F.D. software, FLUENT.

As the wing is going to be used in subsonic regimes, the software options were

changed in the way that is suitable for this purpose.

In the definition steps, the pressure base solver with Spalart-Almaras as the viscous

turbulent model was defined. The constant density option for the air used which is

blown toward the wings, but the value 0.904 Kg/m3 was chosen for density due to the

flight altitude (which is almost 10000 ft). In the boundary condition panel , the

velocity of the air is set to the value 51.389 m/s which is equivalent with 185km/hr (

cruise speed). The angle of attack is set to zero because incident angle is zero

according to the design results.

The SIMPLE algorithm was used as the discretization scheme, due to its generality.

In the reference values panel, the wing geometry data should be entered in order to

get the correct lift and drag coefficients.

Setting up the solver according to the hints mentioned above the solution process is

now ready to start.

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5 .4 .7 .2 Results and discussion:

After nearly 1000 iteration passed the results were obtained as follow:

Table 5-4solution results of wing at fluent

Since the forces shown above are for just one wing and the whole lift and drag forces

are nearly twice as much. So it can be said that the wings produce enough lift to

maintain the aircraft level in the air while cruising at 0 degree angle of attack with

the speed of 185 km/hr.

It is remarkable that the lift force produced by wings, as can be seen, is 16 percent

more than the aircraft weight. This extra lift could be suppressed via elevators, and is

essential to cruise at lower speeds. So that the aircraft can cruise in a range of

velocities, by simply trimming the elevator.

5.4.8 LATERAL CONTROL SURFACES

According to the following table which extracted from database the related flap and

aileron area to wing area ratio can be estimated for aileron and flap. Aileron area to

wing area ratio (Sa/Sw) is equal to 0.09 and flap area to wing area ratio (Sf /Sw) is

equal to 0.11. On the other hand, for calculating the dimension of the aileron and flap

the related chord to mean wing chord ratio must be determined. As it seen in the

following table Ca/Cw is equal to 22% and Cf /Cw is equal to 22.3%. The mean chord

of HADAF wing is equal to 1.317 m. Another important factor which affects the

design of control surfaces is the wing area. It is equal to 13.073 m2. Now all

parameters are available to design control surfaces.

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Table 5-5Database of lateral control surfaces parameters

Sa/Sw Sf/Sw Ca/Cw Cf/Cw

CTSW 0.081777 0.128984 21.50113 21.50113

dynamic 0.061648 0.140686

Jabiru 170 0.059266 0.141359 20.74127 20.74127

pioneer 200 0.134182 0.146072 28.01959 28.01959

Rambo 0.080054 0.080054 16.76621 16.76621

sport cruiser 0.091938 0.082014 18.44568 15.96967

Tecnam 0.104255 0.116092 21.91037 21.91037

Zodiac 0.111784 0.111784 28.54085 28.54085

AVERAGE 0.090613 0.118381 22.27501 21.9213

Finally the geometry of wing with control surfaces is as below:

Figure 5-13wing geometry with control

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5.4.9 VERIFYING CLEAN AIRPLANE MAXIMUM LIFT COEFFICIENT

AND SIZING THE HIGH LIFT DEVICES

The scope of this section is to present a methodology for determining:

1- Whether or not the wing geometry selected before is consistent with the required

value of clean airplane

2- Verifying the type and size of high lift devices, determined in section 4-8.

5.4.9.1 VERIFYING THE MAXIMUM CLEAN LIFT COEFFICIENT

During the past sections the following parameters have been derived:

Clean : =1.4

Take-off : =1.6

Landing : =1.7

The lift coefficient, produced by wing planform, , must be consistent with the

required value of clean airplane Roskam suggests that:

It It should be noted that, the subscript L is devoted to the wing, while l is selected for

a 2d airfoil. The coefficient was suggested to be 1.05, for long-coupled aircrafts and

1.1, for short-coupled aircrafts. Roskam defines long-coupled and short-coupled

aircrafts as follows:

Short-coupled: if Lh/C < 3.0

Long-coupled: if Lh/C < 5.0

Since HADAF 1404, is a short coupled aircraft, the value of 1.1 is selected. This

coefficient, actually takes the trim effect into account. So, the value of , which is

required, is 1.54. In calculating this value, the

, considered to be 1.4.

To verify whether or not the wing can produce its required value of , the

following approximation may be used:

(

)

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Where:

The values of for wing sections can be derived from the figures below, suggested

by the main reference. Note that, it is necessary to compute the Reynolds number for

wing root and tip, separately.

Figure 5-14 Effect of thickness ratio and Reynolds number on section maximum lift coefficient

Fig.5-13 Effect of thickness ratio and Reynolds number on section maximum lift coefficient

The for NACA 5314 with t/c 14, and average Reynolds is 1.6. Since in

HADAF aircraft, , and therefore the would be equal to

1.472 from the eq.4.4. This value for is the wing planform lift coefficient,

which must be consistent within 5 percent by the required lift coefficient

.

The value above is consistent with the value of within 5 percent. So, the

wing planform design and the airfoil selection is verified.

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5.4.9.2 High lift devices design verification

The next step is to determine the maximum lift coefficient increment, which is

produced by high lift devices.

Take-off: (

)

Landing: (

)

The factor 1.05 in the above equations, is for accounting the additional trim penalties

incurred by the use of flaps.

When the flaps are down, the incremental section coefficient can be found as:

(

)

Where is defined in figure 5.14and where is found from:

(

)

Figure 5-15Definition of flapped wing area

The factor accounts for the effect of sweep angle. The sweep angle in HADAF

aircraft is assumed to be 0 degree. For straight, tapered wings the ratio can be

computed from:

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Where the span stations and are defined in fig. 5.14 The incremental section lift

coefficient due to flaps, is related to its counterpart , as defined in figure

5.15.

In preliminary design, it is conservative to use:

(

)

Figure 5-16Relationbetween∆Cl and∆Cl,maxL

Where the factor k is found from figure 5.17, the magnitude of incremental section lift

coefficient due to flaps, depends on the following factors:

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Figure 5-17Effect of flap chord ratio and flap type on k=

The flap-to-chord ratio Cf/c of the flaps

The type of flaps used

The flap deflection angle used.

In case of plain flaps we have:

Where and k

/ may be found from figure 7.5 and 7.6 respectively.

Comparing the and

we can deduce that the high lift devices produce more

lift increment, than what is required from the design criteria.

5.5 DECIDE ON THE OVERALL STRUCTURAL WING CONFIGURATION

The choices here are between cantilever wings and braced or strutted wings. Braced

wings are used primarily on relatively low speed airplanes. But the interference drag

due to struts and wing weight is generally unfavorable. But here using a cantilever

wing will cost much while the influence is not that much. So the braced wing was

chosen as structural configuration.

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5.6 COMPUTE THE WING FUEL VOLUME

It will be assumed here that the wing fuel is carried in what called a 'wet wing' which

means there are no separate fuel tanks.

Torenbeek suggest the following equation for estimating wing fuel volume in

preliminary design:

rtW

WWWWWrWF

ctct

ctbSV

)//()/(

})1/()1{()/)(/(54.0 225.02

litVWF 109

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5.7 ROAD MAP

Finally, the below diagram shows the outline of the design process:

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5.8 References:

1. Airplane design by: Dr. Jan Roskam: part .2

2. Introduction to Aerodynamics by:Dr. Anderson

3. Aerodynamics for engineers by: Bertin

4. Wing modeling tutorials

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1404 Aircraft Design Book Preliminary Tail Sizing

6 Preliminary Tail Sizing

6.1 Introduction

The purpose of this section is to present a step-by-step method for deciding on the

size and disposition of the empennage as well as of the longitudinal and directional

control surfaces. This section presents the basic methods for designing the aircraft

tail and analyzing its stability and its role on the aircraft performance.

Figure 6-1Roadmap of Tail Design Process

The aerodynamic design of the tail-plane is based on many specific requirements

regarding its functions. The tail in an aircraft provides equilibrium in steady flights

(trim), damps the disturbances to help the aircraft maintain its stability, generates the

essential aerodynamic forces for maneuvering the aircraft. The control forces

involved, must be acceptable to the pilots, whether the airplane is in trimmed or out-

of-trim conditions.

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6.2 Empennage Functions

In fact, tails are little wings. Many of the previous discussions, concerning wings, can

also be applied to tail surfaces. The major difference between a wing and a tail is

that, while the wing is designed to operate normally at only a fraction of its lift

potential. Any time in flight that a tail comes close to its maximum lift potential, and

hence its stall angle, something is very wrong!

Tail provides trim, stability, and control to the aircraft. Trim refers to the generation

of a lift force that, by acting through some tail moment arm about the center of

gravity, balances some other moment produced by the aircraft.

The other major function of the tail is control. The tail must be sized in the way that

provides adequate control power at all critical conditions. These critical conditions

for the horizontal tail or canard typically include nose-wheel liftoff, low-speed flight

with flaps down, and transonic maneuvering. For the vertical tail, critical conditions

typically include engine-out flight at low speeds, maximum roll rate, and spin

recovery.

It must be noted that control power depends upon the size and type of the moving

surfaces as well as the overall size of the tail itself. For example, several airlines use

double-hinged rudders to provide more engine-out control power without increasing

the size of the vertical tail beyond what is required for dutch-roll damping.

The aircraft stability and control requirements are usually considered in three flight

regimes. Roll response (motion about the x axis) is conventionally provided by

ailerons but for some aircraft's layouts this is not feasible so differential motion of the

horizontal tail-plane is used. The pitch response, often termed longitudinal stability

(motion about the y axis), dictates the size of horizontal stabilizers (conventional tail,

or canard, or both!). The yaw and sideslip response, termed lateral stability (motion

about the z axis), dictates the vertical stabilizers.

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Figure 6-2Three different axes on aircraft

Each of the three types of motion are not independent since motion about one axis

causes an effect on the others but for simplification in the initial project stage and for

conventional layouts it is acceptable to consider them both de-coupled. The aircraft

layout will have a considerable effect on its stability and control, so intense care is

needed if some unusual arrangements are proposed. The following list identifies some

of the layout considerations for each flight condition.

6.2.1 Pitch

For the horizontal tail, trim primarily refers to the balancing of the moment created

by the wing. An aft horizontal tail typically has a negative incidence angle of about 2-

3 degrees to balance the wing pitching moment. As the wing pitching moment varies

under different flight conditions, the horizontal tail incidence is usually adjustable

through a range of about 3 degrees up and down.

The following points have to be taken into account when sizing the tail-plane.

- The tail-plane has to cope with the required center of gravity (c.g.) travel in

the en-route flight regime. A typical c.g. range would be from a forward c.g.

of 10% of the mean aerodynamic chord to a rear c.g. of 35% of the mean

aerodynamic chord. It must have enough power to provide the necessary trim

and control.

- The forward c.g. position with a typical tricycle undercarriage and with take-

off flap will give the highest load on the nose wheel. The tail-plane needs to be

sized to provide enough force to lift the nose at the required rotation speed.

- The tail-plane must be large enough to be able to trim the aircraft on the

approach with full landing flap at the worst c.g. position and at the same time

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provide enough power to control and flare the aircraft at touchdown. During

the flare maneuver ground effect can have a significant influence.

- The wing section profile, wing plan-form and position relative to the aircraft

center of gravity will affect aircraft stability due to the pitching moments

created by the lift, drag and aerodynamic pressure distribution. Higher wing

camber, thicker section and flapped wings have intrinsically larger

aerodynamic moment which will demand a larger tail to balance the aircraft.

- Engine installation has an influence due to the vertical offset of the thrust line

from the aircraft center of gravity and the pitching inertia contribution from

the engine mass position.

- The position of the horizontal tail, in the vertical plane, relative to the wing

and fuselage (and rear engines) influences the effectiveness of the tail to

produce the balancing force. A high (T) tail positioned is the most effective

configuration because of being away from the fuselage interferences. The tail

should be positioned relative to the jet efflux so that effects from throttle

changes are avoided or kept to a minimum. The position of the horizontal tail

relative to the wing in side view will also determine the aircraft's ability to enter a

deep stall. This will be discussed more in section 6.3.1.1.

6.2.2 Yaw

Aircraft yawing is accomplished through the movements of rudder. For the vertical

tail, the generation of a trim force is normally not required because the aircraft is

usually left-right symmetric and does not create any unbalanced yawing moment. The

vertical tail of a multi-engine aircraft must be capable of providing a sufficient trim

force in the event of an engine failure.

The following points need to be taken into account when sizing the fin.

- The fin size must be such as to cope with the required c.g. travel in the en-

route flight regimes.

- In the event of an engine failure particularly for engines mounted on the wing,

the fin must be capable of generating a sufficient side force to balance the

resulting de-stabilizing moment.

- The cross-wind requirement in the landing configuration can often size the

fin.

The engine failure case, especially in the take-off condition, is usually the critical

sizing criterion for the fin, particularly for aircrafts with wing-mounted engines.

6.2.3 Roll

The primary controls here are usually wing mounted ailerons or spoilers or a

combination of both. An alternative would be to use differential controls on the

horizontal stabilizers. These have mainly been used on fighter aircraft where the

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inertia in roll is so much lower than on the larger civil transports. The roll controls

have to be sized to produce rates of roll and acceleration in roll to meet the

appropriate requirements.

6.3 Empennage Sizing

Tail surfaces are used to both stabilize the aircraft and provide control moments

needed for maneuver and trim. Because these surfaces add wetted area and structural

weight they are often sized to be as small as possible. Although in some cases this is

not optimal, the tail is generally sized based on the required control power as

described in other sections of this chapter. However, before this analysis can be

started, several configuration decisions are needed to be made. This section discusses

some of the considerations involved in tail configuration selection.

For designing aircraft empennage, it‟s necessary to use step-by-step method. This

method consists of three parts:

1- Decision on the empennage configuration to be used

2- Determination of the empennage disposition

3- Determination of the empennage size

6.3.1 Empennage Configuration

Generally, four types of configurations are common to be used:

1- Conventional configurations

2- Canard configurations

3- Three-surface configurations

4- Butterfly empennage configurations

For HADAF1404 aircraft, only conventional configuration is considered and

described. The basic configuration for the tail surfaces in conventional configuration

consists of a horizontal fixed tail-plane, stabilizer, and a vertical fixed fin, each

having a hinged rear flap acting as an elevator for pitch control and a rudder for yaw

control respectively. A dorsal fairing is often incorporated into the base of the fin to

preclude the possibility to fin stall.

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Figure 6-3Aft tail Variations

A large variety of tail shapes have been employed on aircraft over the past century.

These include configurations often denoted by the letters whose shapes they resemble

in front view: T, V, H, +, Y and inverted V. The selection of the particular

configuration involves considerations of a complex of design variables, but here are a

few of the reasons these geometries have been used.

The conventional configuration with a low horizontal tail is a natural choice since

roots of both horizontal and vertical surfaces are conveniently attached directly to the

fuselage. In this design, the effectiveness of the vertical tail is large because

interference with the fuselage and horizontal tail increase its effective aspect ratio.

Large areas of the tails are affected by the converging fuselage flow, which can

reduce the local dynamic pressure.

Figure 6-4Scottish Aviation Bulldog

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A T-tail is often chosen to move the horizontal tail away from engine exhaust and to

reduce aerodynamic interference. The vertical tail is quite effective, being 'end-plated'

on one side by the fuselage and on the other by the horizontal tail. By mounting the

horizontal tail at the end of a swept vertical, the tail length of the horizontal can be

increased. This is especially important for short-coupled designs such as business jets.

The disadvantages of this arrangement include higher vertical fin loads, potential

flutter difficulties, and problems associated with deep-stall.

Figure 6-5Lockheed C 5A

One can mount the horizontal tail part-way up the vertical surface to obtain a

cruciform tail. In this arrangement the vertical tail does not benefit from the end-

plating effects obtained either with conventional or T-tails, however, the structural

issues with T-tails are mostly avoided and the configuration may be necessary to

avoid certain undesirable interference effects, particularly near stall.

Figure 6-6Raytheon Hawker 800XP

V-tails combine functions of horizontal and vertical tails. They are sometimes chosen

because of their increased ground clearance, reduced number of surface intersections,

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or novel look, but require mixing of rudder and elevator controls and often exhibit

reduced control authority in combined yaw and pitch maneuvers.

H-tails use the vertical surfaces as endplates for the horizontal tail, increasing its

effective aspect ratio. The vertical surfaces can be selected shorter, since they enjoy

some of the induced drag savings associated with biplanes. H-tails are sometimes

used on propeller aircraft to reduce the yawing moment associated with propeller

slipstream impingement on the vertical tail. More complex control linkages and

reduced ground clearance discourage their more widespread use.

Y-shaped tails have been used on aircraft such as the LearFan, when the downward

projecting vertical surface can serve to protect a pusher propeller from ground strikes

or can reduce the 1/rev interference that would be more severe with a conventional

arrangement and a 2 or 4-bladed prop. Inverted V-tails have some of the same

features and problems with ground clearance, while producing a favorable rolling

moments with yaw control input.

6.3.1.1 Variations on the basic arrangement

i. Variable incidence tail-plane

The forward section of the horizontal surface is capable of rotation through a range

of angles of attack. In this way, it may be used to adjust the pitch trim, especially in

case of deployment of the high lift devices introduces significant pitching moment

increments.

ii. All moving or "flying" tail-plane

In this concept the whole surface is used as the primary pitch control with the

elevator. Such an arrangement offers significant advantages at transonic and

supersonic speeds when the effectiveness of conventional trailing edge controls is

much reduced and fuselage bending can result in unfavorable loads on a fixed tail-

plane. Some combat aircrafts use differential movement of the two sides of the

horizontal surface to provide roll control.

iii. Vertical position of the horizontal tail

As a general rule, the horizontal tail should not be placed directly in the propeller

slipstream. But it is observed that many airplanes in fact do have the horizontal tail in

the slipstream. The reasons against this arrangement are:

a) The slipstream will usually cause the tail to buffet which leads to structure-

borne cabin noise. Tail buffet can also lead to early structural fatigue.

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b) Rapid power increases or decreases called for by the pilot can result in

undesirably large trim changes.

Single engine propeller driven airplanes usually do have the empennage mounted in

the slipstream. This does enhance elevator effectiveness and rudder effectiveness

during the take-off roll. On the other hand, it also causes considerable tail buffet

during the take-off roll in some airplanes.

There is not usually a problem with a vertical tail mounted in the slipstream at the aft

end of a fuselage.

The horizontal tail is within the wing downwash field which has the effect of reducing

the effectiveness as a stabilizer. The degree of this reduction is a function of the

vertical location of the tail relative to the wing and the effect may be reduced by

significant upward movement in tail location (Figures 6.6, 6.7, and 6.8). In general a

horizontal tail mounted at the top of the fin can be smaller than would otherwise be

the case. Unfortunately such an arrangement accompanies some disadvantages. There

is a mass penalty on the fin due to higher loading and aero-elastic effects, and there is

also the possibility of a deep stall. Essentially this may occur when the aircraft pitches

nose up rapidly and reaches an attitude such that the tail-plane is virtually ineffective

as a stabilizer in the conventional sense. In this situation, the aircraft is in a stable

stalled condition from which it may be difficult or impossible to recover.

Although means are available to resolve this difficulty it is suggested that a high-

mounted 'T' tail should only be used when it is really necessary, as may be the case of

a high-mounted swept back wing configuration or when an engine intake is placed at

the bottom of the fin. A possible alternative for the 'T' tail which does not suffer from

the deep stall problem is to mount the tail-plane very low. Unfortunately in most cases

this is not an option because of tail down ground clearance limitations, but it is

worthy of consideration on smaller aircraft, such as combat types, especially when the

wing is positioned high on the fuselage, Figure 6.9.

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Figure 6-7British Aircraft Corporation TSR2

Figure 6.8 shows the effect of the deep stall on the aircraft. Notice that the tail-

plane lies in a region where the airflow has little net velocity in the longitudinal

direction. This degrades the effectiveness of the tail-plane to produce a lift force to

take the aircraft out of the stall attitude. This produces a stable flight condition with

little forward speed but with a steady vertical descent. Without the ability to recover

from this condition the aircraft will eventually crash. This unsafe situation can be

avoided by positioning the tail-plane outside (usually below) the area of the stalled

wing wake. Wind tunnel tests would be used to verify the safe effectiveness of tail

position in the wing stall attitude.

Figure 6-8Avoidanceofaircraft“deepstall”condition

Figure 6.9 illustrates the boundaries of the acceptable locations for a horizontal tail

to avoid this deep stall. It must be noted that low tails are best for stall recovery. It

must also be noticed that a tail approximately in line with the wing is acceptable for a

subsonic aircraft, but may cause problems at supersonic speeds due to the wake of the

wing.

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Figure 6-9Aft tail positioning

6.3.2 Empennage Disposition

After selection of suitable configuration for tail, now, it comes to decide about the

location of empennage components on the aircraft. In other words, empennage

moment arms (Xh,Xv) and some other parameters must be known.(These parameters

are shown in the figure 6.10).

To achieve a good disposition, these parameters are determined by analyzing the

plans of several ultra- light aircrafts. The database analysis results are used for

HADAF aircraft. The samples of these determinations will be presented in next

sections.

To keep the aircraft weight and drag as low as possible it is desirable to keep the

empennage area as small as possible. This in turn can be achieved by locating the

empennage components in the way that they have the largest possible moment arm

relative to the critical center of gravity.

It must be noticed that in some airplanes (carrier based airplanes are on example)

severe restrictions are place on the allowable length, height and width.

6.3.3 Empennage Size

Sizing the empennage for a conventional configuration means deciding on the

magnitude of Sh and Sv. For a first „cut‟ at the size of either the vertical or the

horizontal tail, the so-called V-method is often used. The tail volume coefficients are

defined as in equations 6-1 and 6-2.

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Vh= Xh.Sh /Sc (6-1)

Vv= Xv.Sv/Sb (6-2)

Having determined which type of aircraft best fits the aircraft being designed, suitable

values for Vh and Vv are selected. This can be done by averaging or comparison to

specific types. Having selected the volume coefficients, since the moment arms (Xh and

Xv) are known, the tail areas can be computed.

In deciding which value for Vv to use, care must be taken that the lateral disposition of

the engines is not too dissimilar. It must be noticed that vertical tail sizes are often

dictated by the engine-out condition.

6.3.4 Final Calculations

In order to determine the moment arms and volume coefficients which best fit

HADAF1404 specifications, some wing geometry data were needed to be known; such

as wing areas, wing span and wing chord.

To find the most suitable coefficients for HADAF aircraft, these coefficients for some

of the aircrafts in the database computed, as shown in table 1. Then, the coefficients

which will be used in design of HADAF1404 vertical and horizontal tails can be

determined by using the average values or choosing between them.

Table 6-1specifications of horizontal and vertical tails of database aircrafts

Name

Wing

Span

(m)

Wing

Area

(sqm)

Wing

Chord

(m)

HT

Area

(sqm)

VT

Area

(sqm)

Xh

(m)

Xv

(m) Vh Vv

CTSW 10.22 12.06 1.11 2.17 0.91 3.7 3.75 0.59977 0.02768

Dynamics 9 10.3 1.39 1.53 0.9 3.5 3.6 0.37403 0.03495

Jabiru 9.6 9.29 1 1.67 0.9 3.45 3.29 0.62018 0.03320

Parrot 9.5 11 1.43 2.17 1.18 4.3 3.7 0.59319 0.04178

Pioneer 7.55 10.5 1.44 2.12 0.9 3.43 3.45 0.48092 0.03916

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Rambo 9.1 10.1 1.2 1.88 1.04 3.28 3.69 0.50877 0.04175

Remos 9.29 10.96 1.2 1.65 0.9 4 3.91 0.50182 0.03456

Sport

Cruiser 8.5 11.8 1.61 2.25 1.1 3.95 3.74 0.46781 0.04101

Tecnam 8.99 12.4 1.35 1.75 1.3 4 3.71 0.41816 0.04326

Zodiac 8.23 9.1 1.77 1.76 0.8 2.95 3.6 0.32234 0.03845

According to the volume coefficients computed as shown in table 6-1, the ones best fit

target aircraft HADAF1404 desires are chosen as Vv=0.027 and Vh=0.45 by choosing

the average values.

Having a wing area of 13.074sqm and a wing span of 10.8 m and the moment arms as

Xh=4.2 m and Xv=3.6 m, the area of target aircraft HADAF1404 is calculated as

Sv= 0.94 sqm

Sh= 2 sqm

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Figure 6-10 Parameters needed in V-method

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6.4 Planform Geometry of Empennage

This step involves evaluating the following parameters:

1- Aspect Ratio

2- Sweep Angle

3- Taper Ratio

4- Thickness Ratio

5- Airfoil

6- Dihedral

7- Incidence Angle

Table 6.2 provides some guidance in making these choices for Homebuilts. The

selection of these items follows some of the same reasoning used in selecting these

items for the wing. However, most of those reasons will be mentioned here.

Table 6-2 Planform Design Parameters for Homebuilts

Dihedral

Angle

Incidence

Angle

Aspect

Ratio

Sweep

Angle

Taper

Ratio

Horizontal

Tail +5 – -10

0 fixed to

variable

1.8 –

4.5 0 – 20 0.29 – 1.0

Vertical

Tail 90 0

0.4 –

1.4 0 – 47 0.26 – 0.71

The design of the horizontal tail for optimum performance, stability and control is

concentrated on it efficiency in producing the required lift and pitching moment. In

the process of lateral-directional stability analysis, the horizontal tail design must

consider the boundaries of both angle of attack and sideslip.

This requires a detailed investigation of the design parameters of horizontal tail,

especially, the aspect ratio on the stability of the aircraft in lateral-directional motion.

6.4.1 Aspect Ratio

The ratio of span and the average chord is called, "aspect ratio". This factor is of

direct influence because of its effect on the lift-curve slope. For manual control

systems the c.g. range satisfying the stick-force requirements will be widened, or the

required tail-plane size may be reduced. For aircraft with a fixed stabilizer the

forward c.g. limitation required to cope with the stall is favorably affected with

increasing aspect ratio. If out-of-trim conditions are the predominant factor, a high

aspect ratio is not always desirable.

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According to Ref. 4:

i. The horizontal tail aspect ratio has significant effect on damping in roll. For

the range of the aspect ratio considered in the analysis, the aircraft still stable,

is an indication for heavy damped total design.

ii. For spiral mode, there is an effect of the horizontal tail aspect ratio with a

good convergent, stable level for the range considered.

iii. The rolling convergence mode and Routh discriminant are improved with the

increase in horizontal tail aspect ratio.

iv. There are no effects of the horizontal tail aspect ratio as a design parameter on

the lateral-directional stability derivatives, damping, frequency of Dutch roll

mode, lateral numerical parameters of all modes and the characteristic

equation coefficient, E.

These results are sufficient for the designer for further developments of the lateral-

directional stability characteristics of this aircraft according to the requirements.

From the database gathered in last step, aspect ratio of each tail was calculated by

the formulas below and is shown in table 2. In these formulas, b is the span of each

tail. Like before, suitable value of these ratios for target airplane can be selected by

averaging or comparison to specific types.

ARv= (bv2)/Sv (6-3)

ARh= (bh2)/Sh (6-4)

Table 6-3Aspect Ratio of horizontal and vertical tails of database aircrafts

Name HT Area

(sqm)

VTArea

(sqm) Bh (m) Bv (m) AR HT AR VT

CTSW 2.17 0.91 2.3 0.9 2.437788018 0.89010989

Dynamics 1.53 0.9 2.4 0.9 3.764705882 0.9

Jabiru 1.67 0.9 2.5 1.1 3.74251497 1.344444444

Parrot 2.17 1.18 2.7 1.1 3.359447005 1.025423729

Pioneer 2.12 0.9 2.4 1.05 2.716981132 1.225

Rambo 1.88 1.04 2.6 1.3 3.595744681 1.884615385

Remos 1.65 0.9 2.5 0.9 3.787878788 0.9

Sport

Cruiser 2.25 1.1 2.9 1.1 3.737777778 1.1

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Tecnam 1.75 1.3 2.8 1.5 4.48 1.730769231

Zodiac 1.76 0.8 2.2 0.9 2.75 1.0125

According to the range of aspect ratios of horizontal and vertical tails of these

aircrafts, aspect ratio of each tail of HADAF 1404 was chosen as below by choosing

the average amounts.

ARv= 1.1

ARh=3.4

These ratios can be reasonable as the ratio for horizontal tail of homebuilt airplanes

can be chosen from 1.8 to 4.5 and from 0.4 to 1.4 for the vertical tail. So the spans of

each tail will be as below.

bv= 1 m

bh=2.6 m

6.4.2 Sweep Angle

The sweep angle ( ) is usually measured as the angle between the line of the 25%

chord and a perpendicular to the root chord.

In selecting sweep angle combinations for tail aft configurations it is important to

ensure that the critical Mach number for the tails is higher that of the wing. An

increment of is usually sufficient.

For horizontal tail, positive sweep is occasionally used on low-speed aircraft to

increase the tail-plane moment arm and the stalling angle of attack, although the

result is a decrease in the lift-curve slope. Up to about 25 degrees of sweepback there

is still an advantage. The sweepback angle may be determined by the condition of a

straight elevator hinge line, which is sometimes imposed in interest of structural

simplicity. For vertical tail of subsonic aircrafts, sweep angle is usually chosen

between 25 and 45 degrees.

If the wing is very highly swept, the horizontal tail sweep is not increased this much

because of the effect on lift curve slope. For supersonic aircrafts, higher sweep angles

may be used if the leading edge Mach number is intended to be subsonic.

So for each tail of HADAF 1404, sweep angle was chosen as moderate values as:

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v=35°

h=4°

6.4.3 Taper Ratio

Considering the empennage planform to have straight lines for the leading and

trailing edges, the taper ratio, is the ratio of the tip chord to the root.

Figure 6-11Root and tip cord

CordRoot

CordTip

Tail-plane taper has a slightly favorable influence on the aerodynamic

characteristics. A moderate taper is usually chosen to save structural weight. For

vertical tails this ratio usually differs from 0.26 to 0.71. But for horizontal tails of

homebuilt aircrafts, it differs from 0.45 to 1.So for each tail of HADAF 1404, taper

ratio was chosen as moderate values as:

λv=0.5

λh=0.85

6.4.4 Thickness Ratio

When we know the wing section we definitely know the t/c ratio of it. This t/c is very

important in determining the drag and lift coefficients of the airfoil. For tails which

usually use symmetrical airfoil it‟s about 9 to 18 percent.

This parameter is also responsible in determining the critical Mach number.

Thickness ratio and the sweep angle are usually gathered in some charts to determine

the critical Mach number.

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As we can easily see, the thickness ratio has a diverse effect on the critical Mach

number, and as we increase the t/c the critical Mach number decreases. While the

critical Mach number increases with the increase of the sweep angle. This verifies our

previous statements about the effect of the sweep angle.

As the t/c increases the increase in the speed of the free stream velocity during the

pass over the upper edge of the airfoil increases, causing the critical Mach number to

decrease.

The tail surfaces should have lower thickness and/or higher sweep than the wing

(about 5 degrees usually) to prevent strong shocks on the tail in normal cruise. Tail

t/c values are often lower than that of the wing since t/c of the tail has a less

significant effect on weight.

6.4.5 Dihedral Angle

The dihedral angle of the horizontal tail is the angle between a horizontal plane

containing the root chord and a plane midway between the upper and lower surfaces

of the horizontal tail. This angle will have a good influence in the aircraft roll

equilibrium.

Figure 6-12Dihedral angle

The position of the tail-plane relative to the propeller slipstream or jet efflux may

make it desirable to shift it slightly in an upward direction. This may be achieved by

using a certain degree of dihedral.

Since enough stability is produced by the designed wing, any positive dihedral angle

in horizontal tail will cause more difficult in controlling the airplane in rolling.

As for vertical tails of the homebuilt, dihedral angle is assumed as 90 degrees, it is

assumed for HADAF 1404 that this number is 90 degrees. And for horizontal tail of

HADAF 1404, this angle is considered to be 0 degree.

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6.4.6 Incidence Angle

When the aircraft is moving on the ground the horizontal tail makes an angle with the

horizon that is the incident angle. This will cause positive or negative lift in order to

obtain stability. For the present case this angle is 0 degree. Because the airfoil of the

wings is designed in the way that it produces enough lift so that no lift is needed to be

captured from horizontal tail.

For all types of aircrafts, incidence angle for vertical tail is considered as 0 degree in

order not to produce side forces.

6.4.7 Airfoil Shape

The basic requirements are that the airfoil section should have a high Cl and a large

range of usable angles of attack.

On all aircraft, the vertical stabilizer and rudder create a symmetric airfoil. This

combination produces no side force when the rudder is aligned with the stabilizer and

allows either left or right forces, depending on the deflection of the rudder.

Frequent use is made of approximately symmetrical airfoils with a thickness ratio of 9

to 12 percent and a large nose radius. Typical of such airfoils are NACA 0011/0015.

For HADAF 1404 airplane, NACA 0012 can be reasonable and best fit. So it is used

in our design.

Figure 6-13Shape of NACA 0012 airfoil

Using analyzing software like DesignFOIL, lift coefficient and drag coefficient

diagrams can be achieved and shown as below.

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Figure 6-14diagram of lift and drag coefficients for NACA0012 airfoil

6.5 Control Surfaces Sizing

After all these, it's the time to decide on the sizes of the control surfaces. In an aircraft

empennage, these contain rudder and elevator.

6.5.1 Elevator

The elevator is the small moving section at the rear of the stabilizer that is attached to

the fixed sections by hinges. Because the elevator moves, it varies the amount of force

generated by the tail surface and is used to generate and control the pitching motion

of the aircraft. There is an elevator attached to each side of the fuselage. The

elevators work in pairs; when the right elevator goes up, the left elevator also goes

up. Figure 6.15 shows what happens when the pilot deflects the elevator.

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Figure 6.15 elevator function in an airplane

The elevator is used to control the position of the nose of the aircraft and the angle of

attack of the wing. Changing the inclination of the wing to the local flight path

changes the amount of lift which the wing generates. This, in turn, causes the aircraft

to climb or dive. During takeoff, the elevators are used to bring the nose of the

aircraft up to begin the climb out. During a banked turn, elevator inputs can increase

the lift and cause a tighter turn. That is why elevator performance is so important for

fighter aircraft.

The elevators work by changing the effective shape of the airfoil of the horizontal

stabilizer. Changing the angle of deflection at the rear of an airfoil changes the

amount of lift generated by the foil. With greater downward deflection of the trailing

edge, lift increases. With greater upward deflection of the trailing edge, lift decreases

and can even become negative as shown on this slide. The lift force (F) is applied at

center of pressure of the horizontal stabilizer which is some distance (L) from the

aircraft center of gravity. This creates a torque on the aircraft and the aircraft rotates

about its center of gravity. The pilot can use this ability to make the airplane loop. Or,

since many aircraft loop naturally, the deflection can be used to trim or balance the

aircraft, thus preventing a loop. If the pilot reverses the elevator deflection to down,

the aircraft pitches in the opposite direction.

On many fighter planes, in order to meet their high maneuvering requirements, the

stabilizer and elevator are combined into one large moving surface called a

stabilator. The change in force is then created by changing the inclination of the

entire surface, not by changing its effective shape as is done with an elevator.

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6.5.2 Rudder

At the rear of the fuselage of most aircraft one finds a vertical stabilizer and a rudder.

The rudder is the small moving section at the rear of the stabilizer that is attached to

the fixed sections by hinges. Because the rudder moves, it varies the amount of force

generated by the tail surface and is used to generate and control the yawing motion of

the aircraft. Figure 6.16 shows what happens when the pilot deflects the rudder, a

hinged section at the rear of the vertical stabilizer.

Figure 6.16 rudder function in an airplane

The rudder is used to control the position of the nose of the aircraft. Interestingly, it is

NOT used to turn the aircraft in flight. Aircraft turns are caused by banking the

aircraft to one side using either ailerons or spoilers. The banking creates an

unbalanced side force component of the large wing lift force which causes the

aircraft's flight path to curve. The rudder input insures that the aircraft is properly

aligned to the curved flight path during the maneuver. Otherwise, the aircraft would

encounter additional drag or even a possible adverse yaw condition in which, due to

increased drag from the control surfaces, the nose would move farther off the flight

path.

The rudder works by changing the effective shape of the airfoil of the vertical

stabilizer. Changing the angle of deflection at the rear of an airfoil will change the

amount of lift generated by the foil. With increased deflection, the lift will increase in

the opposite direction. The rudder and vertical stabilizer are mounted so that they will

produce forces from side to side, not up and down. The side force (F) is applied

through the center of pressure of the vertical stabilizer which is some distance (L)

from the aircraft center of gravity. This creates a torque on the aircraft and the

aircraft rotates about its center of gravity. With greater rudder deflection to the left as

viewed from the back of the aircraft, the force increases to the right. If the pilot

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reverses the rudder deflection to the right, the aircraft will yaw in the opposite

direction.

Some fighter planes have two vertical stabilizers and rudders because of the need to

control the plane with multiple, very powerful engines.

6.5.3 Size of Elevator and Rudder

Elevators and rudders generally begin at the side of the fuselage and extend to the tip

% of the tail span. High-speed aircrafts sometimes use rudders of large chord which

only extend to about 50% of the span. This avoids a rudder effectiveness problem

similar to “aileron reversal” .

Control surfaces are usually tapered in chord by the same ratio as the wing or tail

surface so that the control surface maintains a constant percent chord. Rudders and

elevators are typically about 25-50% of the tail chord.

According to the range of ratio of elevator area to horizontal tail area and rudder

area to vertical tail area of these aircrafts, these ratios for HADAF 1404 was chosen

as below by using average values.

Table 6-4 Ratio of Control Surfaces Area to Tail Area

Name HT Area (sqm) VT Area (sqm) Se/Sh Sr/Sv

CTSW 2.17 0.91 0.2 0.57

Dynamics 1.53 0.9 0.32 0.36

Jabiru 1.67 0.9 0.5 0.33

Parrot 2.17 1.18 0.4 0.5

Pioneer 2.12 0.9 0.5 0.5

Rambo 1.88 1.04 0.15 0.45

Remos 1.65 0.9 0.25 0.46

Sport

Cruiser 2.25 1.1 0.15 0.5

Tecnam 1.75 1.3 0.2 0.45

Zodiac 1.76 0.8 0.4 N/A

Se/Sh= 0.35

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Sr/Sv= 0.5

Using these ratios, the area of rudder and elevators become clear as:

Se = 0.7 sqm

Sr= 0.47 sqm

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6.6 Final Results

As described in last steps, all dimensions and specifications of empennage became

clear. These specifications are listed below.

Table 6-5 Specifications of Each Tail of HADAF1404

Horizontal Tail Vertical Tail

Area 2 0.94

Aspect Ratio 3.4 1.1

C/4 Sweep Angle 4 35

Taper Ratio 0.85 0.5

Thickness Ratio 12 % 12 %

Dihedral Angle 0 90

Incidence Angle 0 0

Airfoil NACA 0012 NACA 0012

Control Surface Ratio 0.35 0.5

Having all dimensions of each tail determined, it's the time to model the whole

empennage in modeling software. One of the most applicable of them is Solidworks

software.

SolidWorks is a 3D mechanical CAD (computer-aided design) program that runs on

Microsoft Windows and was developed by Dassault Systèmes SolidWorks Corp., a

subsidiary of Dassault Systèmes. SolidWorks is currently used by over 1.3 million

engineers and designers at more than 130,000 companies worldwide.

In order to achieve 3D and 2D models of horizontal and vertical tails of HADAF1404

aircraft, airfoil sections must be imported to Solidworks. This can be done by

DesignFOIL. Having this done, those models can be achieved and shown as below.

HADAFTM


Figure 6-153D model of HADAF1404 Vertical Tail in Solidworks

Figure 6-162D model of HADAF1404 Vertical Tail in Solidworks

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Figure 6-173D model of HADAF1404 Horizontal Tail in Solidworks

Figure 6-182D model of HADAF1404 Horizontal Tail in Solidworks

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6.6 References

1. Roskam, J., Airplane Design: Part II, Preliminary Configuration Design and


2. Kroo, I., Aircraft Design: Synthesis and Analysis.

3. Torenbeek, E., Synthesis of Subsonic Airplane Design.

4. Yass, M.A.R, Effect of Airplane Tail Aspect Ratio on Lateral-Directional

Stability.

5. Howe, D, Aircraft Conceptual Design Synthesis

6. Raymer, D.P, Aircraft Design: A Conceptual Approach

7. Jenkinson, Simpkin, Rhodes, Civil Jet Aircraft Design

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7 Landing Gear

7.1 INTRODUCTION

The purpose of this book, is to provide a step-by-step method for designing the

landing gear system. The design of the landing gear, which is considered as “the

essential intermediary between the airplane and ground”, is one of the most

fundamental aspects of aircraft design.[3]

Today, not only much has been learned about all aspects of landing gear design, but

new materials have also become available to help the designer provide the most

efficient shock absorption, in the smallest space, with the lowest weight and cost.

Landing gear location and length are determined by the c.g. location ,tail-down angle

requirements to suit takeoff and landing attitudes, tip over, and general airframe

configuration.

The objectives in the preliminary design phase can be summarized as follows:

i. In the concept formulation phase, the landing gear location and the number

and size of the wheels is determined. The former is, at this time, a function of

center-of-gravity location and general structural arrangement. The number

and size of wheels depends upon the weight of the aircraft and braking

requirements.

ii. 2) In the project definition phase, the general configuration of the aircraft has

been decided and the preliminary design activities become more transparent

and are presented in more details. Proposal preparation usually occurs at the

end of this phase and a concerted effort must be made to provide as much

detail and credibility as possible. The objective of the proposal is to sell the

product; to do that, the customer must be convinced that every facet of the

proposed aircraft is what he wants and that it is better than any competitor's

product and the need for detail and analysis to dispel any argument

concerning its capability.

Figure 7.1.1 illustrates the preliminary design activity and the factors to be

recognized .For instance, in one project, the flotation requirement was established

after an analysis had been made of many landing gear configurations and flotation

was then related to cost.

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Figure 7.1.1: Preliminary Road Map[1]

Marketing Requirements

Gear Location and Type Concept Formulation

Request for Proposal

Project Definition Gear Layout

Tires,Wheels,Brakes

Flotation Analysis

Basic Kinematics

Steering Concept

Special Features

Tradeoff Studies Concept Freeze

Proposal

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7.2 FIXED / RETRACTABLE LANDING GEAR

Generally, if the cruise velocity of the airplane is more than 150kts, retractable gears

will be advised and usually, in the lower speeds, fixed gears will be suitable. For

ultra-light aircrafts, fixed gears don‟t produce much drag force in front of air flow

because their speed is always lower than 150kts.[4]

In addition, economic limitations and cost considerations, cause the fixed one to be

better than it's opponent in this case.

On the other hand, a retractable landing gear eliminates the drag of the legs and

wheels, which can be rather high (in fact several times larger than the drag of the tail-

plane). But the installation of a retractable landing gear also adds weight to the

model, which has to be compensated by a larger lifting force, which causes more

induced drag.

The results show that at high speeds the retractable landing gear always reduces the

required power by 2% to 3%.[7]

Only at considerably low speeds and high loads, the lighter fixed undercarriage

model is advantageous.

As the control of flaps, retractable gears maybe controlled mechanically,

hydraulically, or electrically. After takeoff, the landing gear folds into the fuselage,

where it is stored during flight until shortly before the landing. Usually, because of

having an advanced technology, retractable landing gears are not economical (Table

7.2.1 compares fixed and retractable landing gear characteristics).

Table 7.2.1: Fixed and Retractable Characteristics

Gear Type Fixed Retractable

Aerodynamic drag High Minimal

weight Low High

Complexity and cost Low High

Maintenance cost Insignificant Significant

In this section it is described the most important reasons to select non-retractable

(fixed) gears for this aircraft.

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7.3 LANDING GEAR CONFIGURATION TYPES

In order to select a suitable landing gear configurations the choices here are as

follows:

1. Tail wheel or Tail-dragger

2. Conventional (nose-wheel or tricycle)

3. Tandem

4. Outrigger

5. Beaching gear

Wheeled undercarriages normally come in two types: "tail-dragger" undercarriage,

where there are two main wheels towards the front of the aircraft and a single, much

smaller, wheel at the rear(i.e. skid); or tricycle undercarriage where there are two

main wheels (or wheel assemblies) under the wings and a third smaller wheel in the

nose.

7.3.1 Tail-wheel(Tail-dragger)[2]

Figure 7.3.1.1: Taildragger Configuration

The tail-dragger arrangement was common during the early propeller era, as it

allows more room for propeller clearance. Tail-draggers are considered harder to

land and take-off because the arrangement is unstable, that is, a small deviation from

straight-line travel is naturally amplified by the greater drag of the mainwheel which

has moved farther away from the plane's center of gravity due to the deviation, and

usually require special pilot training. The Concorde, for instance, had a retractable

tail "bumper" wheel. Delta wing aircrafts need a high angle of attack (AOA) when

taking off. Some aircraft with retractable conventional landing gear have a fixed tail

wheel, which generate minimal drag (since most of the airflow past the tail wheel has

been blanketed by the fuselage) and even improve yaw stability in some cases.

The tail-dragger configuration does have advantages. The rear wheel means the plane

naturally sits in a nose-up attitude when on the ground; this is useful for operations

on unpaved surfaces like gravel, where sands blast, could damage the propeller. The

tail-wheel also transmits loads to the airframe in a way that is less likely to cause

airframe damage over time operating on rough fields. The simpler main gear and

small tail-wheel results in both lighter weight and less complexity in the case of using

http://en.wikipedia.org/wiki/Conventional_landing_gear

http://en.wikipedia.org/wiki/Tricycle_undercarriage

http://en.wikipedia.org/wiki/Concorde

http://en.wikipedia.org/wiki/Delta_wing

http://en.wikipedia.org/wiki/Yaw_angle

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retractable. Likewise, a fixed-gear tail-dragger exhibits less interference drag and

form drag in flight than a fixed-gear aircraft with tricycle gear. Tail wheels are

smaller and less expensive to buy and maintain and manual-handling of a tail-wheel

aircraft on the ground is easier.

Its advantages can be listed in brief as follows

The tail-wheel is small, light and of simple design.

When brakes are applied the vertical loads on the main gear will increase, therefore

reducing the risk of skidding.

The main reasons why the tail-wheel undercarriage has been almost completely

superseded by the nose-wheel or tricycle gear is that it also suffers the following

drawbacks:

Violent brakes tend to tip the aircraft onto its nose.

The braking force, acts ahead of the center of gravity and thus has a destabilizing

effect when the aircraft is moving at an angle of yaw relative to its track. This may

cause the ground loop.

In a two point landing, a tail-down moment will be created by the impact force on the

main landing gear, resulting in an increase in lift, which makes the aircraft bounce.

The inclined cabin floor will be uncomfortable for the passengers and inconvenient

for loading and unloading.

In the tail-down attitude the inclination of the fuselage will limit the pilot‟s view over

the nose of the aircraft.

During the initial takeoff run, the drag is considerably high until the tail can be

raised.

7.3.2 Nose wheel (Tricycle)

Most modern aircrafts have tricycle undercarriages. Tricycle gear describes an

aircraft undercarriage, or landing gear, arranged in a tricycle fashion.

Figure 7.3.2.1: Nose wheel Configuration

http://en.wikipedia.org/wiki/Interference_drag

http://en.wikipedia.org/wiki/Form_drag

http://en.wikipedia.org/wiki/Aircraft

http://en.wikipedia.org/wiki/Undercarriage

http://en.wikipedia.org/wiki/Tricycle

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The tricycle arrangement has one wheel in the front, called the nose wheel, and two or

more main wheels slightly aft of the center of gravity. Because of the ease of operating

of a tricycle gear aircraft on the ground, the configuration is the most widely used on

aircraft.

Tricycle gear is essentially the reverse of conventional landing gear or tail-dragger.

Tricycle gear aircrafts have the advantage of being much more difficult to be made

'nose up', which is significantly likely to happen for a tail-dragger in the case of

hitting a bump or heavy brakes applied. Tricycle gear planes are also easier to handle

on the ground and reduce the possibility of a ground loop. This is due to the main

gear being behind the center of mass. Tricycle gear also provides an advantage of

better vision for pilot, as the nose of the aircraft is level and, unlike in the aircrafts

with conventional landing gear, does not block the view ahead. Tricycle gear aircrafts

are easier to land because the attitude required to land on the main gear is the same

as that required in the flare, and they are less vulnerable to crosswinds. As a result,

the majority of modern aircraft are fitted with tricycle gear.[2]

Generally, the merits and drawbacks of the nose-wheel gear are the opposite of those

of the tail-wheel type. These advantages are:[6]

The braking forces are located behind the C.G and have the

stabilizing effect. Thus let the pilot to use the brakes with no

limitation.

With the aircraft on the ground, the fuselage and the cabin floor are

practically level.

The pilot enjoys a broad sight ahead, with more ease and comfort.

The nose-wheel is a safeguard against the aircraft turning over and

so protects the propeller when used.

During the initial part of the takeoff, the drag does not increase

excessively.

In a two-point landing the main gear creates the nose-down pitching

moment, which tends to restrain the aircraft and keep it from the

bounce.

The steady increase in landing speeds of modern aircraft has accentuated these advantages,

so that they carry more weight than the following disadvantages.

The nose unit must take 20% to 30% of the aircraft‟s weight in a steady brake

condition and it is therefore relatively heavy.

The landing gear will probably have to be fitted at a location where special structural

provisions will be required. In the case of retractable nose-gear on light aircraft it

may also be very difficult to find a stowage space inside the external contours of the

aircraft.

http://en.wikipedia.org/wiki/Center_of_gravity

http://en.wikipedia.org/wiki/Conventional_landing_gear

http://en.wikipedia.org/wiki/Ground_loop_%28aviation%29

http://en.wikipedia.org/wiki/Landing

http://en.wikipedia.org/wiki/Aircraft_attitude

http://en.wikipedia.org/wiki/Flare

http://en.wikipedia.org/wiki/Crosswind

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7.3.3 Tandem

An unusual undercarriage configuration which has two main wheels in line, astern

under the fuselage (called a tandem layout) and a smaller wheel near the tip of each

wing. Tandem landing gears are sets of four tires aligned and set under the fuselage

of the aircraft. The wheels are attached using fortified metal beams and struts off the

fuselage of the aircraft. Usually, tandem and outrigger gears are combined.

Here the main wheels are arranged practically in the plane of symmetry of the

aircraft and the front and rear wheels absorb landing impact forces of the same

magnitude. Use of the tandem gear is justified when much emphasis has to be placed

on the following advantages:[5]

Both main legs are placed at nearly equal distances ahead of and behind the center of

gravity thus locally creating space for payload close to C.G.

The wheels may be retracted inside the fuselage without interrupting the wing

structure. The increase in fuselage weight, if any, will depend on other factors.

Against these we have to set the following disadvantages:

Outrigger (Tandem) wheels will be required to stabilize the aircraft on the

ground. However, by using two pairs of main leg instead of single ones, a

certain amount of track may be obtained, resulting in a reduction of the load

on the outriggers.

The pilot must carefully maintain the proper touchdown attitude in order to

avoid overstraining the gear. Care has to also to be taken to limit the angle of

bank during the landing to avoid overstraining the outriggers. It may

sometimes be possible to locate the rear legs close to the center of gravity to

the aircraft, and so reduce this drawback, but that also means losing the

opportunity to have an unobstructed space.

At the table 7.2 tricycle and tail-dragger and tandem configurations are

compared:

Table 7-1Comparing between Tricycle , Taildragger and Tandem

Gear Type Tricycle Tail dragger Tandem

Ground loop behavior Stable Unstable Stable

Visibility over the nose Good Poor Good

Floor attitude on the ground Level Not level Level

Weight Medium Low High

Steering after touchdown Good Poor Good

http://en.wikipedia.org/wiki/Tandem

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Steering while taxing Good Poor Good

Take-off rotation Good Good Good

Take-off procedure Easy Needs skill Easy

So, as noted above, it is suggested to use tricycle gear configuration for designing this

aircraft. Useful operation conditions, economic consideration, safety conditions and

other factors are considered to select this kind of landing gear configuration.

7.4 DISPOSITION OF LANDING GEAR AND STRUT

The positioning of the landing gear is based primarily on stability considerations

during taxiing, liftoff and touchdown, i.e., the aircraft should be in no danger of

turning over on its side once it is on the ground. Compliance with this requirement

can be determined by examining the takeoff/landing performance characteristics and

the relationship between the location of the landing gear and the aircraft c.g. For

instance it is very important for an aircraft that, the touchdown loads exert in a

manner which tip the aircraft, nose down. Otherwise the aircraft would bounce.

There are two geometric criteria which need to be considered in deciding the

disposition of landing gear struts:

1. Tip-over criteria

2. Ground clearance criteria

7 .4 .1 Tip-Over Criteria:

i. Longitudinal Tip over criteria:[4]

The main landing gear must be behind the aft C.G location. The 15 degree angle

shown in figure 7.4.1.1 represents the usual relation between main gear and aft C.G.

Figure 7.4.1.1: Longitudinal Tip over Criteria

ii. Lateral Tip over criteria:[4]

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The lateral tip over is dictated by ψ angle. Figure 7.4.1.2 shows the ψ angle in the

tricycle:

Figure 7.4.1.2: Stable configuration of Landing Gear System

7 .4 .2 Ground Clearance Criteria:[2]

The Figures 7.4.2.1.a and 7.4.2.1.b show the required ground clearance angles. The

lateral ground clearance angle applies to tricycle and tail-dragger but the

longitudinal ground clearance angle applies to tricycle only.

Figure 7.4.2.1.a

Figure 7.4.2.1.b

The available pitch angle (θ), (based on the landing gear limitations), at liftoff and

touchdown must be equal, or preferably exceed, the requirements imposed by

performance or flight characteristics. A geometric limitation to the pitch angle is

detrimental to the liftoff speed and hence to the takeoff field length. Similarly, a

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geometric limitation to the roll angle (φ) could result in undesirable operational limit

under cross-wind landing condition.

A geometric limitation to the pitch angle will be detrimental to the liftoff speed and to

the takeoff distance. A geometrical roll angle limitation may result in an undesirable

operational limit in the case of crosswind landing.

The geometric limits may be reproduced in the θ -φ diagram. The various boundaries

define the point where the rear fuselage tail, the wingtip, engine, trailing edge flaps

and any other parts of aircraft, just touch the ground plane.

Figure 7.4.2.2: θ-φDiagram

For a given aircraft geometry and gear height (hg), the limit for the takeoff/landing

pitch angle follows directly from Figure7.4.2.2. The roll angle at which the tip of the

wing just touches the ground is calculated using the expression.

(7.4.2.1)

In this case, Γ is taken as the dihedral angle, s is the wing span, t is the wheel track,

and Λ is the wing sweep. Similar conditions may be deduced for other parts of the

aircraft, except that Γ, Λ and s in Equation (7.4.2.1) must be replaced with

appropriate values. For example, the permissible roll angle associated with nacelle-

to-ground clearance is determined with the following values: Γ measured from the

horizon to the bottom of the nacelle in the front view, Λ measured from the chosen

landing gear location to the engine in the top view.

Table 7.4.2.1: Specification of Struts

name Strut length Installation Angle of

struts

Distance between

fuselage and ground

CTSW 0.54 45 0.54

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Dynamics 0.57 45 0.56

Jabiru 0.54 43 0.63

Parrot 0.64 43 0.5

Pioneer 200 0.56 45 0.63

Rambo 0.58 45 0.62

Remos 0.66 48 0.5

Sport Cruiser 0.47 40 0.61

Tecnam 0.53 45 0.66

Zodiak 0.51 45 0.52

The takeoff rotation angle is prescribed in preliminary design, and then estimated.

The final values for θ and φ are found as the detailed performance characteristics of

the aircraft become available. The pitch angle at liftoff (θLOF) is calculated using the

expression:

(

√

) (7.4.2.2)

Where αLOF is the highest angle of attack anticipated for normal operational use, VLOF

is the liftoff speed, g is the gravitational acceleration CL,LOF is the lift coefficient, and

dCL/dα is the lift-curve slope. As shown in Figure7.4.2.1.b, the dimension of l1 and l2

are defined by the line connecting the tire-ground contact point upon touchdown and

the location of the tail bumper. [2]

The detailed aerodynamic data required for equation (7.4.2.2) is not always available

at the conceptual design stage. In most aircraft the aft-body and/or tail bumper is

designed such that the aircraft cannot rotate by more than a specified number of

degrees at liftoff. Typically, the value is between 12 and 15 degrees.

Table 7.4.2.2: Required Geometric Specifications

name base track Ln Lm θ φ ψ

CTSW 1.4 1.7 1.1 0.3 13 19 55

Dynamics 1.4 1.9 1.2 0.2 13 20 50

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Jabiru 1.4 1.9 1 0.4 13 19 55

Parrot 1.65 2.04 1.1 0.55 17 20 55

Pioneer

200 1.385 1.8 1.11 0.275 13 20 56

Rambo 1.4 1.573 1.12 0.28 16.25 19 64

Remos 1.542 2.14 1.26 0.282 14.9 17 63

Sport

Cruiser 1.35 1.86 1.05 0.3 16 20 50

Tecnam 1.6 1.46 1.15 0.45 16 23 52

Zodiak 1.3 2.13 1.05 0.25 13 14 55

7.5 COMPUTING THE MAXIMUM STATIC LOAD[4]

The following equations can be used to compute the maximum static load per strut:

Nose wheel strut

(7.5.1)

Main gear strut

(7.5.2)

Pn Pm

Lm

Ln

C.G

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7.6 Selection of Tires[2]

The choice of tire type depends on some special factors such as maximum allowable

static loading, maximum allowable runway speed and airfield conditions. Figure

below shows examples of the recommended tires.

i. Grooved tires: This pattern normally used is provided with ribs. This is necessary in

order to obtain good adhesion on wet runways and minimize the effects of cutting

action of stones and flints in the runway surface.

ii. Chinned tires: Aircrafts with engines located in the wing roots or at the sides of the

rear fuselage use tires provided with a chine.

iii. Anti-shimmy tire: They are mainly used on light aircraft with a single, castoring nose

wheel. Shimmy is an oscillatory, combined lateral-yaw motion of the landing gear

caused by the interaction between dynamic tire behavior and landing gear structural

dynamics. The motion typically has a frequency in the range of 10 to 30 Hz. The

amplitude may grow to a level of annoying vibrations affecting the comfort and

visibility of the pilot, or can even result in severe structural damage and landing gear

collapse. Shimmy can occur on both nose and main landing gears, although the latter

case is more rare. Most publications on shimmy found in the open literature typically

deal with twin-wheeled cantilevered landing gears, which apparently are more

susceptible to shimmy vibrations compared to other landing gear configurations.[8]

iv. Smooth Contour: This type was designed for airplanes with non-retractable landing

gears (this type considered obsolete).

v. Low Pressure: This type is comparable to smooth contour and has beads of smaller

diameter, larger volume and lower pressure.

vi. Extra High Pressure: It has high load capacity and narrow width. It is almost

universal on military aircrafts.

vii. Low Profile High Pressure: It is useful for very high take-off speeds.

According to the some safety considerations about strength of tires and the prevention

of dangerous effects of Shimmy phenomenon on the nose gear, the anti-shimmy tire is

suggested for HADAF1404

aircraft.

HADAFTM


The choice of the main gears and nose gear size is generally made on the basis of the

static loading. The load is determined by the weight of aircraft, the number of legs

and wheels. The aircraft is considered to be taxing without braking, at low speed, and

hence the wheels load follows from the static equilibrium (refer to the part 7.5).

By using table 7.6.1, tires dimensions are determined. At the first, it must be specified

the load ratios (main gear static load to take-off weight and nose gear static load to

take- off weight). If the static loading factor (ns) is equal to 2 then following ratios will

be determined:

Table 7.6.1: Standard Specifications of Tires[4]

Type Main Gears Nose Gear

Homebuilt

2Pm /WTO Dt bt PSI Pn /WTO Dt bt PSI

0.8 13 5 25 0.17 9 3.4 25

0.78 12 5 45 0.22 12 5 45

0.87 16 6 45 0.13 16 6 45

So, according to the table, following result is determined:

Table 7.6.2: Final SpecificationsofHADAF’sTires

Specifications Nose gear Main gears

Outer Diameter (in) 12 12

Width (in) 5 5

Pressure (psi) 45 45

HADAFTM


7.7 Landing Gear Database:

Gear weight is about 4.0% of the take-off weight. This is the total landing gear weight

including structure, actuating system, and the rolling assembly consisting of wheels,

brakes, and tires. The rolling assembly is approximately 39% of the total gear weight:

(7.7.1)

Figure 7.7.1: Landing gear weight according to maximum takeoff weight[5]

For HADAF 1404 aircraft the takeoff weight is 626.6 kg thus landing gear‟s C.G is

calculated according to the figure below:

Table 7.7.1: Landing gear Database

name Pn Pm WTO Wnose Wmain Xc.g Yc.g

CTSW 1009.028 1849.88 480 6.4 12.8 1.45 0.64

Dynamics 770.785 2312.35 550 7.33 14.66 1.85 1

Jabiru 1527.55 1909.45 545 7.26 14.53 1.9 1.03

Parrot 1958.73 1958.73 599 8 16 2.04 0.82

Pioneer

200 919.37 1855.47 472 6.3 12.6 1.71 1.01

Rambo 931.95 1863.9 475 6.33 12.66 1.576 0.985

Remos 1072.84 2396.77 598 8 16 1.8 1.02

Sport

Cruiser 1308 2286.66 600 8 16 1.7 1

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Tecnam 1655.43 2115.28 600 8 16 1.63 1.1

Zodiak 1122.49 2357.23 595 7.94 15.86 1.74 0.9

By using the following landing gear data base ,for the nearest aircraft to HADAF, landing gear C.G

information is specified. these information will be used to specify the main aircraft‟s C.G. Thus, the

initial information that is used in C.G modification loop for first step is :

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Table 7.7.2: Preliminary Specifications of HADAF's Landing Gear System

base length 1.45 m

track length 1.85 m

Ln 1.13 m

Lm 0.32 m

θ 16

φ 20

ψ 50

Pn 1331.45

Pm 2350.85

strut length 0.56

Installation angle of struts 45

WTO 626.6 kg

Wnose 8.3 kg

Wmain 16.6 kg

Xc.g 1.8 m

Yc.g 1 m

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7.8 Final Drawing of Landing Gear System

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7.9 Table of Final Results

Title Specification

Landing Gear Type Fixed

Configuration Type Tricycle

Number of Main Wheels 2

Main Gear Position 2.12 m

Nose Gear Position 0.67 m

Track Length 1.85 m

Base Length 1.45 m

Strut Length 0.56 m

Installation Angle of Strut 45

Distance Between Fuselage and

Ground 0.6 m

Landing gear C.G (x component) 1.8 m

Landing gear C.G (y component) 1 m

Outer Diameter of Tires 12 in

Tires Width 5 in

Pressure of Tires 45 psi

Tires Type Anti-Shimmy Tire

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7.10 ROAD MAP

Landing Gear Design

Conceptual Design

Fixed or Retractable

Configuration Type

Number of Wheels

Landing Gear Position

Main Gear Position

Nose Gear Position

Selection of Tires

Detail Design

Strut-Wheel Interface

Shock Absorber

Brakes Consideration

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7.11 REFERENDES

1) Currey,N.S.,Landing Gear Design Handbook, Lockheed Georgia Company , Marietta,

Georgia,30063,1982.

2) Torenbeek, E., Synthesis of Subsonic Airplane Design, Kluwer Boston Inc.,

Hingham, Maine, 1982.

3) Conway, B.G., Landing Gear Design, Chapman Ball,London, England, 1958.

4) Roskam. J., Airplane Flight Dynamics and Automatic Flight Controls, 1981, Roskam

Aviation and Engineering Corp•• Rt 4, Box 274. Ottawa. Kansas. 66067.

5) http://www.flightsimbooks.com/flightsimhandbook/

6) http://www.mh-aerotools.de/airfoils/pylon_retracts.htm

7) Shimmy of Aircraft Main Landing Gears, I.J.M.Besselin

HADAFTM

1404 Aircraft Design Book Weight and Balance Analysis

8 WEIGHT AND BALANCE ANALYSIS

8.1 INTRODUCTION

The purpose of this book is to determine the coordinates of center of gravity of

HADAF1404 ultra-light airplane and place it in the right location for different loading

scenarios.

The precise location of the aircraft C.G is essential in the positioning of the landing

gear, as well as for other MDO applications, e.g., flight mechanics, stability and

control, and performance. So we put this design procedure to the end of conceptual

design process. The road map of C.G location calculation process according to

Roskam airplane design method is shown in the following diagram. Each part will be

introduced precisely in the rest of report.

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Figure 8-1 Roadmap of Center of Gravity calculation process

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8.2 COMPONENT WEIGHT BREAKDOWN

According to the previously drawn roadmap, the first step of center of gravity location

calculation is to breaking down the airplane components. Our airplane weight

breakdown is as shown in the following table.

Table 8-1HADAF1404 Airplane Weight breakdown

Aircraft Component

1 Fuselage Group

2 Wing Group

3 Empennage Group

4 Engine Group

5 Landing Gear Group

6 Fixed Equipment Group

7 Fuel Group

8 Passenger Group

The other weight components that are stated in the Roskam airplane design method

such as fuel, crew, passenger, etc. can be neglected for our ultra-light aircrafts. It is

supposed that fuel is placed symmetrically in the wing and its weight is added to the

weight of wing. Passengers‟ location is in symmetric form too. So these parts are

omitted from the table.

8.3 PRELIMINARY ARRANGEMENT DRAWING OF AIRPLANE and EACH

COMPONENT C.G LOCATION

According to the roadmap this part is second and third steps of designing process in

which a schematic illustration of side view of our airplane is drawn and different

component breakdown center of gravity in this drawing is denoted. The drawing will

help us to have a visual background in our mind during the rest of the process and

make it easy to choose a coordinate system to measure all distances according to it.

Our airplane preliminary arrangement drawing is shown in figure 2.

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Figure 8-2 Preliminary arrangement drawing of airplane and components C.G location

Figure 8-3top view of airplane

Figure 8-4front view of airplane

1

2

3

4

5

6

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The coordinate system is shown in Figure 8-2 and all distances are measured

according to it. Zero reference point is selected in front of airplane so that all

coordinates are positive. By suggestion of Roskam method, the pick zero point was

placed in well left and well below of nose to assure about sign errors of results.

8.4 CATEGORIZING THE x, y, z COORDINATE OF C.G OF EACH

COMPONENT

Now in the fourth step of C.G locating process we should categorize the x, y, z

coordinate of each components C.G. The C.G location data of fuselage, wing and

landing gear should be import from the result of their designing reports. Engine C.G

location is found by modeling the motor according to its dimensions which is

published in the catalogue and software calculating. Engine modeling process and

C.G calculation is done in the CATIATM

modeling software. But for fixed equipment

C.G an internet based search is done and some data is gathered according to other

ultra-light specifications published in their catalogues.

In the rest of this part, calculation process of different weight components center of

gravity will be explained. An important hint about the data given in the rest is that

because of assumption of symmetry, each part‟s center of gravity has coefficients just

in the x and y directions. ( Figures 8-3 & 8-4)

8.4.1 FUSELAGE GROUP

According to Roskam recommended formula for fuselage C.G calculation a

preliminary C.G estimation is done but for more accuracy a 3D computer based

calculation is performed too. Both calculated C.Gs are evaluated and the final result

is as reported below.

Table 8-2Fuselage Weight and center of gravity data

WFuselage (kg) XFuselage(m) YFuselage(m)

Fuselage Group 100 1.91 1.4

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8.4.2 WING GROUP

Two methods were chosen to find the exact position of center of gravity of wing in the

chosen coordinate system. First method was suggested formula of Roskam method and

the second was computer-analysis-based method. Using CATIATM

modeling software

and assigning the real material to the model, the exact position of center of gravity can

be calculated. The results are as listed in table 8-3.

Table 8-3Wing Group Weight and center of gravity data

WWing (kg) XWing (m) Ywing(m)

Wing Group 50 2.64 1.82

8.4.3 EMPENNAGE GROUP

Similar to the Wing group C.G calculation the center of gravity of empennage group

was calculated and the results are categorized in table 8-4.

Table 8-4Empennage Group Weight and center of gravity data

WWing (kg) XWing (m) Ywing(m)

Horizontal Tail 10 6.607 1.28

Vertical Tail 10 5.660 1.63

8.4.4 ENGINE GROUP

Based on catalogues of Jabiru engine and its drawings an absolute model was created

and a computer-based C.G calculation is done. The results are as listed in table 8-5.

Table 8-5Engine Group Weight and center of gravity data

WEngine (kg) XEngine (m) YEngine(m)

Engine Group 81 0.7 1.2

8.4.5 LANDING GEAR GROUP

The process of calculating the center of gravity of landing gear and the overall

airplane C.G should be done simultaneously because C.G data of whole airplane is

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one of the important parameters for calculating the dimensions of landing gear.On the

other hand the center of gravity location of landing gear is one of the inputs of overall

C.G location calculation process. So an iteration-based process is taken into

existence. The iterations were done with Microsoft Excel. A program was written to

calculate the landing gear dimensions and C.G location and another one prepared for

calculating the overall center of gravity location. In the first step a location for overall

airplane C.G is assumed and given to the landing gear C.G calculation program as an

input. And the output of this section is given to the whole C.G calculation program and

finally a C.G location is achieved. Evaluating the final result and the preliminary

assumption a new C.G location is assumed. This process continues until the

assumption and final result converges. The final C.G location for landing gear is as

listed in table 6.

Table 8-6Engine Group Weight and center of gravity data

WLanding Gear (kg) XLanding Gear (m) YLanding Gear (m)

Landing Gear Group 20 1.71 0.3

8.4.6 FIXED EQUIPMENTS GROUP

The assignment of component C.G range is based on the geometry, planform, and the

type of components involved. In the case of the primary components, e.g., fuselage,

wing and empennage, the location of these items remains relatively unchanged once

the conceptis frozen. Consequently, the C.G range is expected to be centered near the

volumetric centerof the component and is unlikely to shift too much. For ease of

identification, the primary components can be referred to as the constrained items.

As for secondary components, e.g., equipment and operational items, the location

ofeach component varies from one aircraft concept to another, depending on the

philosophy and preference of the airframe manufacturer. Note that as long as the

stowage and functionality constraints are not violated, these components can be

assigned to any available space throughout the aircraft due to their compactness.

Consequently, the corresponding c.grange is defined by the forward and aft

boundaries of the stowage space within which the item is located. Accordingly, these

components are termed the unconstrained items.

For means of calculating the overall C.G location a primary assumption was took into

account but in the upcoming sections to meet the most aft and forward boundaries it

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may be changed. This assumption made according to the other Homebuilt aircrafts

equipment placement which can be found in the technical brochures. Our primary

assumption is as listed below.

Table 8-7Equipment Group Weight and center of gravity data

WEquipment (kg) XEquipment (m) YEquipment (m)

EQUIPMENT GROUP 150 1.6 1.3

8.4.7 FUEL GROUP

The C.G location of the fuel varies as a function of time as the fuel is being consumed

during the duration of the mission. Given the added freedom in terms of the loading

pattern, these components are also classified as unconstrained items. The maximum

capacity of the fuel tank is 70 liters and its center of gravity can be determined to be

coincident of wing‟s C.G location because it is planned to place the fuel tank in the

wing. So the fuel center of gravity location data is as listed in table 8.

Table 8-8Fuel Group Weight and center of gravity data

WFuel (kg) XFuel (m) YFuel (m)

FUEL GROUP 55 2.64 1.82

8.4.8 PASSENGERS GROUP

HADAF 1404 is a 2 seat ultra-light airplane. The maximum passenger weight is 200

kg and the data of its center of gravity is categorized in table 9.

Table 8-9Passanger Group Weight and center of gravity data

WPassenger (kg) XPassenger (m) YPassenger (m)

PASSENGER

GROUP 200 2 1.35

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8.5 CALCULATING THE xc.g&yc.gOF AIRPLANE

Now it is the time to calculate the overall C.G location of airplane for different

loading scenarios. Because of simplicity of our ultra-light aircraft and the symmetry of

configuration it is just needed to perform the calculating process for just four loading

combination and in two directions. The x component of C.G is the most important

because the longitudinal stability of our aircraft is depending directly to its location

during the flight. So excursion diagram is drawing and discussing just for x component

of C.G.

Four loading combinations are:

1. Empty Weight (C.G1)

∑

∑

2. Empty Weight+Fuel (C.G2)

∑

∑

∑

∑

3. Empty Weight+Fuel+Passenger or Takeoff Weight (C.G3)

∑

∑

∑

∑

4. Empty Weight+Passenger (C.G4)

∑

∑

∑

∑

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When the airplane is standing on the ground the Empty weight C.G is the airplane C.G and

when the fuel is added, C.G2 is the main C.G. But when the passengers sit in the airplane and

the plane is ready for take-off C.G3 is the airplane center of gravity. After finishing the

mission and simultaneously the fuel, C.G4 is the main C.G. These four combinations

constitute the loading scenario which the airplane experiences. Table 10 is the weight and

balance calculation summery.

Table 8-10Weight and Balance Calculation Summery

No. Type of Component

1 Fuselage 100 1.91 191 1.4 140

2 Wing 50 2.64 132 1.82 91

3 Engine 81 0.7 56.7 1.2 97.2

4 Vertical Tail 10 5.66 56.6 1.63 16.3

5 Horizontal Tail 10 6.607 66.07 1.28 12.8

6 Landing gear 20 1.71 34.2 0.3 6

7 Fixed Equipment 150 1.6 240 1.3 195

Empty Weight (C.G1) 421 1.844 1.326

8 Fuel 55 2.64 145.2 18.2 100.1

Empty Weight + Fuel (C.G2) 476 1.936 1.383

9 Passenger 200 2 400 1.35 270

Empty Weight + Passenger (C.G3) 621 1.894 1.33

Take-off Weight (C.G4) 676 1.95 1.37

8.6 WEIGHT C.G EXCURSION DIAGRAM

Excursion diagram shows the sensitivity of center of gravity location to the weight in

different segments of loading scenario. In this diagram four calculated c.g points are

located. The vertical axis is Weight and the horizontal axis is the location of center of

gravity in terms of fuselage station. It is important to identify in this diagram the

loading sequences as well as critical weights such as WE and WTO. The most aft and

the most forward C.G locations can be found as a result of this diagram. In the

following section according to this diagram it is decided that the place of which

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component should be changed. Figure 3 shows the excursion diagram of HADAF

1404.For some airplanes it may be important to also draw C.G excursion diagrams

which reflect the vertical and lateral C.G situations. But it doesn‟t make any sense and

the x direction diagram is enough. The x-position of wing aerodynamic center is 2.27

meter. As it can be seen in Table 10 the most forward and the most aft c.g position is

ahead of aerodynamic center. So as we like the aircraft is always nose down.

Figure 8-5 Excursion diagram

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8.7 C.G EXCURSION DIAGRAM ARGUMENT

In this part the feasibility of using the configuration will be discussed. At first we

compare our airplane resulting C.G range with the airplanes in the same category.

The Roskam suggested C.G range for homebuilt airplanes is about 5 in. some

important principle should be discussed in this section.

1. As it is obvious from the excursion diagram, the c.g range is ahead of the wing mean

aerodynamic center and it is satisfactory because it is good for airplane to be

somehow noise down.

2. The ideal C.G arrangement is one for which the OWE C.G, the fuel C.G and the

payload C.G are in the same vertical location. But it is not possible to achieve the

ideal configuration. The existing C.G configuration is the closest to the ideal case.

HADAFTM


8.8 REFRENCES

1- Roskam, J., Airplane design: PartII , Preliminary Configuration Design of

Airplane.

2- Federal Aviation Regulation, FAR, Part 23

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