Post on 04-Apr-2018
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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.
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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
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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
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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
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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
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Cruise Speed VS Maximum Takeoff Weight
Cruise Speed: 560mph
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
Cruise Speed: 560mph
Altitude: 56000 fts
0
10000
20000
30000
40000
50000
60000
70000
0 200 400 600 800 1000
Altitude
Altitude
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Cruise Speed VS Range
Cruise Speed: 560mph
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
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Cruise Speed VS Wing Span
Cruise Speed: 560mph
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
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Cruise Speed VS Wing Area
Cruise Speed: 560mph
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
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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
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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
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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
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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
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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
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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)
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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
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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
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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+)
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= 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
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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
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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
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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
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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
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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.
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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
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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.
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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.
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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
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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.
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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
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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:
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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
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Power available
Pavl = T V in KW
Thrust available
Tavl= pavl /v in KN
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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
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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
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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
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POWER AVAILABLE VS VELOCITY
0
20000000
40000000
60000000
80000000
100000000
120000000
140000000
160000000
0 200 400 600 800 1000 1200
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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
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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
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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
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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
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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
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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
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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
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THREEVIEWS OF
AIRCRAFT
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