Project RAPP

7/30/2019 Project RAPP

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ACKNOWLEDGEMENT

First and foremost, we wish to acknowledge our debt to `HARD

WORK IS THE KEY TO SUCCESS` who has given us knowledge and

good health. We would like to express to the chairman of our college,

Dr.P.MUTHUVEL RAJ and the principalDr.M.PALANICHAMY, for

providing better working environments and educational facilities.

We are much gratefulMr.KIRUBASHANKAR Head of the

department of the Aeronautical engineering for this encouragement

discussion, valuable comments and many innovative ideas in carrying

out this project. Without his timely help it would have been impossible

for us to complete this work.

We acknowledge in no less terms the qualified and excellent

assistance rendered by Mr.PUGAZHARASAN, Lecturer, Department of

Aeronautical Engineering. We owe a debt of gratitude for his valuable

suggestions, kind inspiration and encouragement.

We most sincerely acknowledge the staff members of Department

of Aeronautical Engineering for their constant inspiration and

suggestions.

We owe a debt gratitude to our parents and friends for their

advice and to keep our spirits high to complete this project.


2/55

Performance

SpecificationsS.No Name of the

Aircraft

Range Cruise Speed Altitude Maximun Speed

1

Boeing

EA- 18G

Growler

1458 miles 777 mph 50000 fts 1190mph

2

Boeing

F/A18/FSuper Hornet

1458 miles 777 mph 50000 fts 1190mph

3

Grumman

F14

Tomcat

1217 miles 550 mph 50000 fts 1544mph

4

McDonnell Douglas

F- 15E

Strike Eagle

1801 miles 620 mph 60000 fts 1650mph

5

Sukhoi

Su30

1864 miles 580 mph 56800 fts 1320mph

6

Sukhoi

Su30MKI

1864 miles 590 mph 56800 fts 1317mph

7

Sukhoi

Su34

1864 miles 600 mph 49200 fts 1180mph


3/55

WING

SPECIFICATIONSS.No Name of the Aircraft Wingspan Length Height Wing Area

1

Boeing

EA- 18G

Growler

44 ft 8.5 inches 60 ft 1.25 inches 16 fts 500 ft2

2

Boeing

F/A18/FSuper Hornet

44 ft 8.5 inches 60 ft 1.25 inches 16 fts 500 ft2

3

Grumman

F14

Tomcat

64 ft 62 ft 9 inches 16 fts 565 ft2

4

McDonnell Douglas

F- 15E

Strike Eagle

42.8 ft 63.8 ft 18.5 fts 608 ft2

5

Sukhoi

Su30

48.2 ft 72.97 ft 20.85 fts 667 ft2

6

Sukhoi

Su30MKI

48.2 ft 72.97 ft 20.85 fts 667 ft2

7

Sukhoi

Su34

48 ft 3 inches 72 ft 2 inches 19 fts 5

inches

770 ft2


4/55

Engine Specifications

S.No Name of the Aircraft No. of

Engines

Engine Details Thrust

Produced

1

Boeing

EA- 18G

Growler

2 2General Electric F414-GE-400 turbofans

97.9KN

2

Boeing

F/A18/F

Super Hornet

2 2 General Electric F414-GE-400 turbofans

97.9KN

3

Grumman

F14Tomcat

22 General Electric F110-GE-

400 afterburning turbofans 123.7KN

4

McDonnell Douglas

F- 15E

Strike Eagle

2 2 Pratt & Whitney F100-229afterburning turbofans

129KN

5

Sukhoi

Su30

2 2 AL-31FL low-bypassturbofans

122.58KN

6

Sukhoi

Su30MKI

2 2 Lyulka AL-31FM1turbofans

123KN

7

Sukhoi

Su34

2 2 Lyulka AL-31FP turbofanswith thrust vectoring

132KN


5/55

Cruise Speed VS Maximum Speed

Cruise Speed: 560mph

Maximum Speed: 1477 mph

0

200

400

600

800

1000

1200

1400

1600

1800

0 200 400 600 800 1000

Maximun Speed

Maximun Speed


6/55

Cruise Speed VS Maximum Takeoff Weight


Maximum Take off Weight: 86000 lbs

0

20000

40000

60000

80000

100000

120000

0 200 400 600 800 1000

Maximun Take off weight

Maximun Take off weight


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Cruise Speed VS Altitude


Altitude: 56000 fts

0

10000

20000

30000

40000

50000

60000

70000

0 200 400 600 800 1000

Altitude

Altitude


8/55

Cruise Speed VS Range


Range: 2000 miles

0

200

400

600

800

1000

1200

1400

1600

1800

2000

0 100 200 300 400 500 600 700 800 900

Range

Range


9/55

Cruise Speed VS Wing Span


Wing Span:14.7m

0

10

20

30

40

50

60

70

0 100 200 300 400 500 600 700 800 900

Wing Span

Wing Span


10/55

Cruise Speed VS Wing Area


Wing Area:62 m

0

100

200

300

400

500

600

700

800

900

0 100 200 300 400 500 600 700 800 900

Wing Area

Wing Area


11/55

Weight Estimation of Aircraft

To estimate the overall weight of the aircraft by using the

weight fraction method

Requirements:

Crew Weight

Payload Weight

And other constant values

Need of Estimation of Weight:

The first technical step in the designing of the aircraft is to

estimate the weight. In order to design an aircraft to our desired

performance we have to know the weight of the aircraft. According

to that weight we can move to next step in designing the aircraft.

Weight Estimation:

The overall weight of the aircraft is calculated by using the

following formulae.

The overall weight is taken as W0

W0 = Wcrew + Wpayload + Wfuel + Wempty

Substituting,

Wf = (Wf / W0) * W0 ; and We = (We / W0) * W0 in the above

equation, we arrive at a equation like this,

W0 = (Wcrew + Wpayload) /(1 - (Wf/ W0) - (We / W0))


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Where,

Wf= Fuel Weight

We = Empty Weight

(Wf/ W0) = Fuel Fraction

(We / W0) = Empty Weight Fraction

Empty Weight Fraction (We / W0) :

Empty Weight fraction ranges from 0.3 to 0.7, so we choose

a compromised value of 0.50 which is for a Air superiority Fighter

type Aircraft.

Estimation of Fuel Fraction(Wf / W0) :

Before estimating the fuel fraction, we have to choose the

typical mission profile as follows

Typical Mission Profile:

Each segment of the mission profile is associated with a

weight fraction, defined as the airplane weight at the end of the

segment divided by the weight at the beginning of the segment.

0 to 1 = Engine start, warm up and taxing

1 to 2 = Take off

2 to 3 = Climb

3 to 4 = Cruse speed at 250.34 m/s

4 to 5 = Loiter maximum 20 minutes

5 to 6 = Descend for initial approach

6 to 7 = Land at airbase


13/55

The fuel fraction can be found through the following formulae

(Wf/ W0) = 1Mff

Where Mffis mass per fuel fraction

Mff= (W1 / W0) * (W2 / W1) * (W3/ W2).

This continues until the path of the mission gets completed.

The constant profile for the weight estimation during the warm up,

taxing, take off, climbing, descend and landing. In order to find the

fuel fraction we need to calculate the weight ratio for the cruise and

loiter.

At Engine start, warm up and taxing (0 1):

Beginning weight is W0 and Ending weight is W1

Therefor the weight ratio is (W1 / W0) = 0.99

At Take off (1

2):

Smillarly here (W2 / W1) = 0.97

At Climb (23):

(W3/ W2) = 0.985

At Cruise Speed (3 4):

We need to use Brequent range equation to find the weight

ratio in cruise speed

(W4/ W3) = e-(RC / V (L / D))


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Where,

R = range of the aircraft in meter

C = specific fuel consumption in kg/kg s

V = cruise speed of the aircraft in m/s (which is 250.34 m/s)

(L/D) = maximum lift to drag ratio

C = 0.5 lb/lbhr for a turbofan engine as reffered in Aircraft

Conceptual Design Book

Therefore,

C = 0.000222 kg/kg sRange = 2000 miles

= 3218688 m.

(L/D) = 12 for fighter aircraft

(W4/ W3) = e-((3218688*0.000222)/(250.34*12))

(W4/ W3) = 0.7034451

In Loiter (4 5):

(W5/ W4) = e-(EC / (L / D))

Where,

E = Endurance or Loiter time

E = 1200 sC = 0.000222 kg / kg s

(L/D) = 12

(W5/ W4) = e-((1200*0.000222) /12)

(W5/ W4) = 0.978044606

At Descend (5 6):


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(W6/ W5) = 0.993

At Landing (67):

(W7/ W6) = 0.995

Therefore,

Mff= (W1 / W0) * (W2 / W1) * (W3/ W2).

Mff= 0.99*0.97*0.985*0.7034451*0.978044606*0.993*0.995

Mff= 0.64299

(Wf/ W0) = 1 Mff

(Wf / W0) = 0.357

Estimation of Crew Weight:

The designed aircraft is a two seater fighter; therefore oly

two pilots can sit.

Therefore the weight is consider as 200 kgs

Wcrew = 200 kgs

Estimation of payload:

This aircraft is a Air Superiority fighter whch can carry lots

of weapons.

Overall we can carry 5500 kgs of pay load

Wpayload = 5000 kgs


16/55

Overall Weight Estimation:

W0 = (Wcrew + Wpayload) /(1 - (Wf / W0) - (We / W0))

W0 = 3636.36 kgs

Fuel Estimation:

Wf= 0.357 * W0

Wf= 12981.8 kgs

Empty Weight Estimation:

We = 0.50*W0

We = 18181.8 kgs

Conclusion:

Thus the overall weight of the aircraft is calculated and this gives

the ides to the estimation of powerplant.

Overall Weight = 36363.6 kgs

Weight of the fuel = 12981.8 kgs

Empty Weight of the aircraft = 18181.8 kgs


17/55


18/55

Power Plant Selection

To select the powerplant which required for the aircraft and

to notice the charateristics of the powerplant

Requirements:

Thrust is required to specify what type of engine is going to

be used

Over all weight is required to calculate the required thrust

Thrust and drag ratio which is required to calculate the thrust

Thrust to Weight ratio:

Assume the Thrust toWeight ratio of the typical fighter as 1

Overall Weight:

The overall weight is estimated by weight fraction method,

therefore the overall weight is about 36363.6 kgs

Thrust Required:

T/W = 1

T = W0 * 1 * 9.8 KN

T = 356.73 KN


19/55

Powerplant Selection:

The thrust required is obtained as 356.73KN. The selection of

powerplant is based on the type of the aircraft. As this aircraft is a

fighter which carries weapons and some avionics instruments toreflect radar system, so placing of the propeller engine or

turboprop engine is impossible. Because the path variation cause

sever damage. So there lies choice between the turbojet and the

turbofan engine. Since it is mutirole fighter it have to attack the

enemis from the low altitude also. The thrust at low altitude is very

less for the turbojet engine. Therefore using Turbofan engine is

suitable for this aircraft.

The Required Thrust is 356.73KN. So we may choose an

engine which have its maximum thrust more than our required

thrust. In such a way we have selected the Pratt and Whitney F135

afterburning turbofan.

So in this case two engines are required to produce the

desired thrust.

Conclusion:

Engine : Pratt and Whitney F135

Engine Type : Afterburning Turbofan

No.of Engines : 2

Thrust Required : 356.73KN


20/55

Engine Selection

Primary Function Fighter Aircrat

Propulsion Two Prat& Whitney F135 afterburning

turbofan

Type Afterburning Turbofan F-35 B also partially

turboshaft

Length 220 inches (5.59 meters)

Diameter 51 inches (1.29 meters)

Dry weight 1701 kgs (3750 lbs)

Compressor Axial 3 stage low-pressure compressor, 6

stage high pressure compressor

Combustors Short annular combustor

Turbine Single stage high pressure turbine, two stage

low pressure turbine

Thrust produced 191.35 KN (Maximum) and 124.6 KN

(intermediate)

Specific Fuel Consumption 0.886 lb/(hr*lbf) or 25 g/KNs (without

afterburner)


21/55

Advantages:

Light Weight starter

Electronic igniter

Remote oil filter

Air condition provision

Optical magnets

Fixed pitch and chord propeller

Disadvantages:

Slip stream component induces drag

Slip stream may disturb the free air flow over the wing


22/55

Wing Design

To estimate the basic values for the wing design and to

choose the type of wing that is to be used in the aircraft.

Type of the wing:Since it is a multirole fighter the delta wing is to be used. The

taper ratio for the delta wing is 0.3 for fighter aircraft

Wing Area:

The area of the wing is 667 ft2. It is selected from thegraoh that is used in the selection parameters.

S = 667 ft2

S = 62 m2

Wing Span:

Wing Span is also selected from the graph as 48.2 ft

.b = 14.7 m

Aspect Ratio:

A.R = b2/S

A.R = 3.49

Lift Coefficient:

At sea level,

W0=36363.6

=1.225


23/55

V=250.34m/s

S=62 m2

W0=

V

2

SCL

36363.6=

(1.225250.34

262)CL

CL sea= 0.0076

SWEPT ANGLE:

MACH CONE,

sin =1 / M

=Sin-1

(1/2)

=30

Swept angle = 25 degree

CHORD:b C =S

C=62/14.7

C = 4.217 m

Aerodynamic center = 0.44.217

=1.69m

TAPER RATIO:=0.3

=Ct/ Cr

Cr=2S / b(1+)


24/55

= 262 /14.7(1+0.30)

Cr=6.4887m

Ct= Cr

= 0.36.4887

Ct=1.95m

Mean Aerodynamic Chord:

Chord is Defined as the distance between the leading and the

trailing edge of the aerofoil. It is calculated as follows,

C = ((2*croot)/3) *((1++2)/(1+))

C = 4.625m

Distance at which the chord loacating:

Y = (b/6)*((1+2*)/1+)

Y = 3.015m


25/55

Conclusion:

Wing Type = Delta wing

Aspect ratio = 3.49Wing Area = 62 m

2

Wing Span = 14.7m

Camber at the root = 6.4887m

Camber at the tip =1.95m

Mean aerodynamic Chord =4.625mLocation of chord = 3.015m


26/55

Aerofoil Selection

To estimate the related values for designing of the aerofoil.

Then the required aerofoil selected and thet is used in the aircraft

construction.

Vstall Estimation:

Vapproach = 1.2 * Vstall

Assume the Vapproach as 250 knots for typical fighter.

Vstall = (250 * 0.3098)/1.2

Vstall = 64.541 m/s

Lift Coefficient estimation:

The maximum lift coefficient is estimated as follows ,CLmax = (2*w0)*(1/v2*s)

Where,

W0 = overall weight of the aircraft.

. = density at the altitude 17068 m

S = surface area of the wing

CLmax = (2*36363.6)/(62*1.4*250.34*250.34)

CLmax = 0.014

Leading edge flap = 0.3

Plain flap = 0.9

Total cLmax = 0.014+0.3+0.9

CLmax = 1.214


27/55

Reynolds number

The Reynolds number is needed to select the aerofoil and it is

calculated as follows

Re no = (*v*c)/

Where,

. = density at 17000 m

V = cruise velocity

C = mean aerodynamic chord

= viscous coefficient

= 0*(T/273)0.75

where

T = thrust required.

0 = 1.734*10-5

N.S/m2

= (1.734*10

-5

)*(356.73/273)

0.75

= 2.119*10-5

re no = (0.175*250.3424*4.625)/ 2.119*10-5

Re no = 9.56*106

Conclusion :

The values required for the selection of the aerofoil is

calculated and with the corresponding values aerofoil is selected.

Vstall = 64.541 m/s. : CLmax = 1.214

Reynolds no = 9.56*106


28/55

CONTROL SURFACE SIZING

To size the control surface of our aircraft and this control surfaces

are reason for the controlled flight.

Aileron:

Span = 90% of the wing span

= 0.9*14.7

= 13.23 m.

Chord =20-25%of wing chord

= 20% of wing chord

= 0.2*4.625

= 0.925 m.

Hinge axis = 5% of aileron chord

= 0.05*0.925

= 0.04625 m.

Rudder:

Span = 90% of the vertical tail span

= 0.9*1.65

= 1.485 m.

Chord = 20-25%of the vertical tail chord= 20%of the vertical tail chord

= 0.2*0.7837

= 0.15 m

Hinge axis = 5% of the rudder chord

= 0.05* 0.15

= 0.007


29/55

Elevator:

Span =90% of the horizontal tail span

= 0.9*5.011

= 4.5 m.

Chord = 25-50% of the horizontal tail chord

= 20% of the horizontal tail chord

= 0.2*2.387

= 0.4774 m.

Hinge axis = 5% of elevator chord

= 0.05*0.4774

= 0.02387 m.

Conclusion:

The control surfaces are the main reason for the three movements.

The three movements are pitching movement, yawing movement,

rolling movement. Thus the control surfaces are designed and

sized.


30/55

DRAG ESTIMATIONTo calculate the amount of drag that is formed in our aircraft.

Drag:Drag is a aerodynamic force which is parallel to the relative

wind. Drag forces is occurred in the aircraft by different ways. So

there are many types of drags are there.

Parasite drag:Drag that is caused due to the payload that is carried in the

aircraft.

Form drag:Form drag is formed due to the shapes that is used in the

aircraft while constructing it.

Skin friction drag :

The skin friction drag is due to the surface

discontinuities and some rashes in the surface.

Zero lift drag coefficient:Parasite drag coefficient cfe = 0.0035 (constant)

Zero lift drag coefficient cdo = cfe*(swet/sref)

Cdo = 4*0.0035

Cdo = 0.014

(swet/srefis considered as 4 from the other fighters value)

Ostwald efficiency factor e = 0.8K = 1/(*A.R*e)

= 1/(3.14*3.49*0.8)

K = 0.114

Drag coefficient :

Cd = cdo + k*cLMAX2

= 0.014+0.173*1.2142


31/55

Cd = 0.269

Drag at cruise speed :

D = 0.5*vcruise2

*s*cd

*= 0.5*250.3424*250.3424*62*0.269

D = 93.538 KN

CLmax/cd= 4.12 = L/D

L =d*4.12

L = 385.37 KN.

Conclusion:

Thus the drag estimation is done and the drag values are

estimated in terms of their coefficients.


32/55

Drag at cruise speed D = 93.538 KN.

PERFORMANCE CALCULATION

The performance parameters are must to be calculated toknow about the performance based specification of the designed

aircraft.

Rate of climb:

The rate at which the aircraft climbs into the sky. It is

calculated as follows,

R/C = ((T-D)*V)/w

Where,

T = thrust required

D = drag estimated

V = vstall

W = overall weight.

= ((348.04173.39)64.541)/31677

R/C = 57.04

Take off calculation :

Vtakeoff = 1.1*vstall

Vtakeoff = 71 m/s.

Landing calculation :

V approach = 250 knots as referred in the book.

Glide angle :

The angle at which the aircraft can glide without using the power

in that time.


33/55

Tan = 1/(L/D)

=476

Maximum range(theoretical) :

Maximum distance covered by the aircraft with the full fuel and

the full weight.

R = h/tan

R = 201168 m.

As calculated theoretically with altitude is taken as 16500 m.

Time to climb:The approximate time taken to climb into the sky.

T = 0h dh/(r/c)

Where,

H = altitude

R/C =rate of climb

T = 2.4 minutes approximately.

Flight path radius:

The maximum radius at which the aircraft turns while manuering

and performing the stunts.

R = (6.96*vstall2)/g

= (6.96*64.5412)/9.81

R = 2955.36 m


34/55

STABILITY:

AERODYNAMIC CENTER (AC):

It is of crucial importance that the aircraft's Centre of Gravity (CG) is located

at the right point, so that a stable and controllable flight can be achieved.In order to achieve a good longitudinal stability, the CG should be ahead of the

Neutral Point (NP), which is the Aerodynamic Centre of the whole aircraft.

NP is the position through which all the net lift increments act for a change in

angle of attack.

The major contributors are the main wing, stabilizer surfaces and fuselage. Thebigger the stabilizer area in relationship to the wing area and the longer the tail

moment arm relative to the wing chord, the farther aft the NP will be and the

farther aft the CG may be, provided it's kept ahead of the NP for stability.

The angle of the fuselage to the direction of flight affects its drag, but has littleeffect on the pitch trim unless both the projected area of the fuselage and its

angle to the direction of flight are quite large. A tail-heavy aircraft will be moreunstable and susceptible to stall at low speed e. g. during the landing approach. Anose-heavy aircraft will be more difficult to takeoff from the ground and to gain

altitude and will tend to drop its nose when the throttle is reduced. It also requires

higher speed in order to land safely.

The angle between the wing chord line and the stabilizer chord line is called

The Longitudinal Dihedral (LD) for a given centre of gravity, there is a LD anglethat results in a certain trimmed flight speed and pitch attitude. If the LD angle is

increased the plane will take on a more nose up pitch attitude, whereas with adecreased LD angle the plane will take on a more nose down pitch attitude.


35/55

There is also the Angle of Incidence, which is the angle of a flying surface

related to a common reference line drawn by the designer along the fuselage.The designer might want this reference line to be level when the plane is flying

at level flight or when the fuselage is in its lowest drag position. The purpose of

the reference line is to make it easier to set up the relationships among thethrust, the wing and the stabilizer incidence angles. Thus, the Longitudinal

Dihedral and the Angle of Incidence are interdependent.

ngitudinal stability is also improved if the stabilizer is situatedso that it lies outside the influence of the main wing downwash. StabilizersAre therefore often staggered and mounted at a different height in orderto improve their stabilizing effectiveness.

It has been found both experimentally and theoretically that, if the

aerodynamic force is applied at a location 1/4 from the leading edge of

a rectangular wing at subsonic speed, the magnitude of theaerodynamic moment remains nearly constant even when the angle of

attack changes. This location is called the wing's Aerodynamic Centre

AC. (At supersonic speed, the aerodynamic centreis near 1/2 of thechord).

In order to obtain a good Longitudinal Stability the Centre of Gravity CG

Should be close to the main wings' Aerodynamic Centre AC. For wings with otherthan rectangular form (such as triangular, trapezoidal, compound, etc.) we have tofind the Mean Aerodynamic Chord - MAC, which is the average for the whole wing.

The MAC calculation requires rather complicated mathematics, so a simplermethod called 'Geometric Mean Chord' GMC or 'Standard Mean Chord' SMC may


36/55

be used as shown on the drawings below. MAC is only slightly bigger than GMCexcept for sharply tapered wings.

Taper ratio = tip chord/root chord.

The mean aerodynamic chord (MAC) distance from the centre line is

calculated as follows:

Geometrical mean chord, GMC = (Croot+ Ctip) / 2.

DETERMINATION OF AC FOR WING:


37/55


38/55

Power required

PR=K1V3+K2/V in kw

Where

K1=1/2 scd0

K2= bw2/ s

b = wing span in m

W= over all weihjt in kg

= density kg/m

3

S= wing area m2

Cdo= zero drag co efficient

Thrust required

TR=PR/V IN KN


39/55

Power available

Pavl = T V in KW

Thrust available

Tavl= pavl /v in KN


40/55

VELOCITYPOWER

REQUIRED

THRUST

REQUIRED

THRUST

AVAILABLE

POWER

AVAILABLE

192 1544523.879 8044.395201 135672 26049024

212 1614265.48 7614.45981 135672 28762464

232 1736332.399 7484.191376 135672 31475904

252 1909954 7579.182538 135672 34189344

272 2135561.83 7851.330259 135672 36902784

292 2414377.921 8268.417538 135672 39616224

312 2748161.422 8808.209685 135672 42329664

332 3139046.82 9454.9603 135672 45043104

352 3589437.31 10197.26508 135672 47756544

372 4101932.561 11026.70043 135672 50469984

392 4679278.595 11936.93519 135672 53183424

412 5324332.266 12923.13657 135672 55896864

432 6040035.601 13981.56389 135672 58610304

452 6829396.96 15109.28531 135672 61323744

472 7695476.982 16303.97666 135672 64037184

492 8641377.955 17563.77633 135672 66750624

512 9670235.686 18887.17908 135672 69464064


41/55

532 10785213.2 20272.95715 135672 72177504

552 11989495.81 21720.10111 135672 74890944

572 13286287.21 23227.77485 135672 77604384

592 14678806.4 24795.28107 135672 80317824

612 16170285.13 26422.03452 135672 83031264

632 17763965.94 28107.54105 135672 85744704

652 19463100.49 29851.38111 135672 88458144

672 21270948.16 31653.19667 135672 91171584

692 23190775.01 33512.68065 135672 93885024

712 25225852.77 35429.5685 135672 96598464

732 27379458.11 37403.63129 135672 99311904

752 29654871.96 39434.67016 135672 102025344

772 32055378.98 41522.51163 135672 104738784

792 34584267.06 43667.00387 135672 107452224

812 37244826.96 45868.0135 135672 110165664

832 40040351.93 48125.42299 135672 112879104

852 42974137.42 50439.12842 135672 115592544

872 46049480.86 52809.03768 135672 118305984

892 49269681.4 55235.06884 135672 121019424

912 52638039.76 57717.14886 135672 123732864

932 56157858.02 60255.21247 135672 126446304

952 59832439.51 62849.20116 135672 129159744

972 63665088.65 65499.0624 135672 131873184

992 67659110.88 68204.74887 135672 134586624


42/55

POWER REQUIRED VS VELOCITY

0

10000000

20000000

30000000

40000000

50000000

60000000

70000000

80000000

0 200 400 600 800 1000 1200


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THRUST REQUIRED VS VELOCITY

0

10000000

20000000

30000000

40000000

50000000

60000000

70000000

80000000

0 200 400 600 800 1000 1200


44/55

POWER AVAILABLE VS VELOCITY

0

20000000

40000000

60000000

80000000

100000000

120000000

140000000

160000000

0 200 400 600 800 1000 1200


45/55

THRUST AVAILABLE VS VELOCITY

0

50

100

150

200

250

300

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15


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velocityLoad

factor

radius of

turnTurn rate

215 1.000152 270089.4553 91295827

215 1.000609 135003.7091 45634020215 1.001371 89956.75848 30407240

215 1.002439 67419.64701 22789214

215 1.003816 53886.34498 18214686

215 1.005503 44854.99392 15161905

215 1.007502 38396.14186 12978688

215 1.009817 33545.1388 11338947

215 1.012452 29765.98686 10061518

215 1.015411 26737.14824 9037704

215 1.018697 24253.9516 8198334

215 1.022317 22179.98681 7497292

215 1.026277 20420.81231 6902653

215 1.030582 18908.93889 6391611

215 1.035239 17594.91835 5947444

215 1.040257 16441.62361 5557607

215 1.045643 15420.69045 5212510

215 1.051408 14510.03633 4904690

215 1.057559 13692.23934 4628258215 1.064109 12953.36062 4378502

215 1.071068 12282.11045 4151605

215 1.07845 11669.25212 3944447

215 1.086266 11107.15976 3754448

215 1.094532 10589.47101 3579459

215 1.103264 10110.84637 3417674

215 1.112477 9666.762794 3267564

215 1.122189 9253.368747 3127829215 1.13242 8867.362103 2997351

215 1.143191 8505.897085 2875168

215 1.154523 8166.503749 2760446

215 1.16644 7847.031021 2652458

215 1.178969 7545.598431 2550568

215 1.192136 7260.551793 2454216

215 1.205972 6990.427689 2362908

215 1.220509 6733.93365 2276208

215 1.235781 6489.919039 2193726


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velocityLoad

factor

radius of

turnTurn rate

215 1.251825 6257.355523 2115115

215 1.268683 6035.324109 2040064215 1.286398 5822.998482 1968293

215 1.305018 5619.635133 1899552

215 1.324593 5424.562206 1833613

215 1.34518 5237.169869 1770271

215 1.36684 5056.907577 1709339

215 1.389639 4883.270554 1650646

215 1.413648 4715.801183 1594038


48/55

BANK ANGLE VS RADIUS OF TURN

0

5000

10000

15000

20000

25000

30000

35000

40000

0 20 40 60 80 100 120

radius of turn


49/55

BANK ANGLE VS RATE OF TURN

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0 20 40 60 80 100 120

rate of turn


50/55

OVERALL WEIGHT VS WING LOADING

0

100

200

300

400

500

600

700

0 10,000 20,000 30,000 40,000 50,000 60,000 70,000

Wing loading


51/55

OVERALL WEIGHT VS THRUST WEIGHT RATIO

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 10,000 20,000 30,000 40,000 50,000 60,000 70,000

Thrust weight ratio


52/55

OVERALL WEIGHT VS ASPECT RATIO

.

0

1

2

3

4

5

6

7

0 10,000 20,000 30,000 40,000 50,000 60,000 70,000

aspect ratio


53/55

OVERALL WEIGHT VS TAPER RATIO

0

2

4

6

8

10

12

14

16

0 10,000 20,000 30,000 40,000 50,000 60,000 70,000

taper ratio


54/55

THREEVIEWS OF

AIRCRAFT


55/55

Documents

Project RAPP