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Transcript of CARGO ADP-2
CHAPTER 1
AIM OF
EXPERIMENT
1
AIM OF THE PROJECT
The aim of this design project is to design a cargo aircraft by comparing
the data and specifications of present aircrafts in this category and to calculate
the performance characteristics. Also necessary graphs need to be plotted and
diagrams have to be included wherever needed.
The following design requirements and research studies are set for the
project:
Design an aircraft that will transport cargo of weight around 600000 kg
over a design range of 4000 km at a cruise speed of about 850 km/h.
To operate from regional and international airports.
To use advanced and state of the art technologies in order to reduce the
operating costs.
To offer a unique and competitive service to existing scheduled
operations.
To assess the development potential in the primary role of the aircraft.
To produce a commercial analysis of the aircraft project.
2
CHAPTER 2
ABSTRACT
3
ABSTRACT
The purpose of the project is to design a cargo aircraft. The aircraft will
possess a high wing, tricycle landing gear and a twin tail arrangement. Such an
aircraft must possess a wide body configuration to provide sufficient capacity
for loading. It must possess turbofan engines to provide the required amount of
speed, range and fuel economy for the operator. The aircraft will possess six
engines.
4
CHAPTER 3
INTRODUCTION
5
DESIGN PROCESS-AN OVERVIEW:
The project design process is the means by which the competing factors and
constraints which affect the design are synthesized with the specialist analytical
inputs to produce the overall configuration. The process may be considered in
three different parts:
Conceptual design studies
Preliminary design studies
Detail design studies
CONCEPTUAL DESIGN STUDIES:
The first activity in the project design process is the ‘conceptual design
study’ in this phase, conventional and novel configurations are considered to
determine layouts which are technically feasible and commercially viable at the
start of the phase all options are considered during the concept design phase the
quantity of data generated on each design will be relatively limited and the man
power expended small. The outcome of the study is the knowledge of the
feasibility of the various concepts and an estimate of the rough size of the most
likely configurations.
Conceptual design begins with either a set of design requirements
established by the prospective customer or a company-generated guess as what
the future customers need. Design requirements include aircraft range, payload,
landing and takeoff distance and speed requirements.
The actual design usually begins with conceptual sketch. A good
conceptual sketch will include the approximate wing and tail geometries, the
fuselage shape and the internal locations of major components like engine,
cockpit, landing gear and fuel tanks.
6
PRELIMINARY DESIGN STUDIES:
At the end of the conceptual design phase all the design layouts will have
been analyzed. Those which are regarded as unfeasible or too commercially
risky will be eliminated. The remainder will be compared after careful
consideration of a suitable selection criterion. It is important not to carry too
many options forward to the next stage as this will dissipate the available effort
and slow down the detailed definition of the preferred design. However, care
must be taken to avoid discarding design layouts too quickly as some may lead
to evolutionary configurations which could give the aircraft a competitive
advantage over aircraft from other companies.
DETAILED DESIGN STUDIES
The detailed design phase is started towards the end of the parametric
analysis. In this part of the design process the layout is refined to a greater level
of detail. With the external shape fixed, the structural framework will be
defined. In this phase, there will be an increasing reluctance to make radical
geometric changes the overall layout of the aircraft. Throughout this phase, the
aircraft weight and performance estimates will be continuously updated as more
details of the aircraft layout becomes available.
THE DESIGN
`Design of any system is of successful application of fundamentals of
physics. Thus the airplane design incorporates the fundamentals of
aerodynamics, structures, performance and stability and control and basic
physics. These are based on certain degree of judgments and experience.
Design is a process of usage of creativity with the knowledge of science
where we try to get most the best things available and to overcome the pitfalls
7
the previous deign has. It is an iterative process to idealism toward with
everyone marching still.
Here the preliminary design has been done of cargo aircraft. The basic
requirements are the high endurance, low weight, high accuracy and long range.
Here the most possible considerations have been taken. The flight parameters
and limitations are studied.
This design project also looks at the aspects like improving the
aerodynamic characteristics as well as the payload. The design project has been
classified into different stages in our design will be as follows.
Collection of comparative data
Selection of aircraft parameters
Preliminary weight estimations
Selection of Power plant
Airfoil selection, flaps, t/c, etc.
Wing layout
Layout of landing gear, loads and tyre selection
Critical performance parameters
3 view diagram
8
CHAPTER 4
V-n DIAGRAM
Introduction:
9
Airplanes may be subjected to a variety of loading conditions in flight.
The structural design of the aircraft involves the estimation of the various loads
on the aircraft structure and designing the airframe to carry all these loads,
providing enough safety factors, considering the fact that the aircraft under
design is a commercial transport airplane. As it is obviously impossible to
investigate every loading condition that the aircraft may encounter, it becomes
necessary to select a few conditions such that each one of these conditions will
be critical for some structural member of the airplane.
Velocity –Load Factor (V-n) diagram:
The control of weight in aircraft design is of extreme importance.
Increases in weight require stronger structures to support them, which in turn
lead to further increases in weight and so on. Excesses of structural weight
mean lesser amounts of payload, thereby affecting the economic viability of the
aircraft. The aircraft designer is therefore constantly seeking to pare his
aircraft’s weight to the minimum compatible with safety. However, to ensure
general minimum standards of strength and safety, airworthiness regulations
(Av.P.970 and BCAR) lay down several factors which the primary structure of
the aircraft must satisfy. These are the
Limit load, which is the maximum load that the aircraft is expected to
experience in normal operation.
Proof load, which is the product of the limit load and the proof factor
(1.0- 1.25), and
Ultimate load, which is the product of the limit load and the ultimate
factor (usually 1.5). The aircraft’s structure must withstand the proof load
without detrimental distortion and should not fail until the ultimate load has
been achieved.
10
The basic strength and fight performance limits for a particular aircraft
are selected by the airworthiness authorities and are contained in the flight
envelope or V-n diagram.
Fig.1
MANEUVERING LOADS
11
The greatest air loads on an aircraft usually comes from the generation of
lift during high-g maneuvers. Even the fuselage is almost always structurally
sized by the lift of the wings rather than by the pressures produced directly on
the fuselage. Aircraft load factor (n) expresses the maneuvering of an aircraft as
a standard acceleration due to gravity
Symmetric Maneuver Load:
These will occur when the aircraft’s pilot (or the autopilot) operates the
longitudinal control surface (e.g. the elevator or canard0 to cause aircraft to
pitch nose-up or nose-down. This action may result in two distinct forms of
acceleration:
Translational, which may be either longitudinal or normal to the flight
path
Rotational
Normal Load Factor:
The loads due to symmetric maneuvers are most commonly analyzed
through use of the definition of a normal load factor (n), whereby:
n= Lift (L)/Weight (W)
The load factor is more properly defined as the component of
aerodynamic force perpendicular to the longitudinal axis divided by the aircraft
weight. But, assuming that in terms of gravitational accelerations (“g”).
.
IMPORTANT VELOCITIES:
12
The main velocities that r plotted in the V-n diagram are:
1 – g Stall Velocity
Design Maneuvering Velocity
Design Cruise Velocity
Design Dive Velocity
1-g Stall Velocity, Vs:
where,
ρ = Density at the cruise altitude (0.41 kgm-3)
(WFDWG / S) = Wing Loading = 691.4x 9.81 Nmm-2(from
ADP-I)
where,
CLmax = 2.35
CD at CLmax = 0.15
CN = 2.354
DESIGN LIMIT LOAD FACTORS n lim pos and n lim neg:
13
1-g Stall Velocity, Vs = 118.19m/s
The positive design limit load factor, nlimpos, must be selected by the
designer, but must meet the following condition:
For cargo aircraft, nlimpos= 3.8
n lim neg = -1.52
3.8 ≥ 2.1 + (24000 / (622125+10000)
3.8 ≥ 2.13
Hence our design satisfied the above condition.
The negative design limit load factor nlimneg, must be selected by the
designer, but must meet the following condition:
nlimneg ≥ 0.4nlimpos for normal and for utility category airplanes
Therefore,
1.52 ≥ 0.4 × 3.8
1.52 ≥ 1.52
Hence our design satisfied the above condition.
-1.3 < n < 3.2
DESIGN MANEUVERING SPEED VA :
The design maneuvering speed VA, must be selected by the designer, but
must satisfy the following relationship:
VA ≥ 118.19√3.8
VA ≥ 230.39 ms-1
And, we take our Maneuvering speed as 230.39 ms-1.
-1.52 < n < 3.8
14
VA = 230.39 ms-1
DESIGN CRUISE SPEED VC:
The design cruise speed VC, must be selected by the designer, but must
satisfy the following relationship:
kc= 33 for transport category
VC ≥ 33x√ 691.4
VC ≥ 865.2 ms-1
DESIGN DIVING SPEED VD:
The design diving speed must satisfy the following relationship:
VD ≥ 1.05 VC
VD ≥ 1301.57 ms-1
SAMPLE V-n MANEUVER DIAGRAM
15
VC = 867.7 ms-1
VD = 1301.57 ms-1
Fig.2
V-n MANEUVER DIAGRAM FOR OUR AIRCRAFT
Fig.3
16
CHAPTER 5
GUST AND
MANEUVERABILITY
ENVELOPES
GUST AND MANEUVERABILITY ENVELOPES
17
Gust envelope of an aircraft refers to the capabilities of a design in
terms of airspeed and load factor or altitude. The term is somewhat loosely
applied, and can also refer to other measurements such as maneuverability.
When a plane is pushed, for instance by diving it at high speeds, it is said to be
flown "outside the envelope", something considered rather dangerous.
CALCULATION
The gust V-n diagram is given by the following formulae
where,
Ude = equivalent gust velocity (in m/sec)
Ve = equivalent airspeed (in m/s)
Gust Alleviation Factor, Kg= 0.88µg / (5.3+ µg) for cargo aircrafts
Mass ratio, µg= 2(w
S )ρgC Lα c
µg = 7.57
Lift curve slope, CLα = 3.96
Mean Chord, C = 11.39 m
Thus, Kg= 0.518
ui=kû; where i=b,c,d ; ûb=20.11m/s; ûc= 15.24m/s ; ûd= 7.62m/s
At high angle of attack, point B, ub = 10.39 ms-1
At level flight, point C, uc = 7.89ms-1
At dive condition, point D, ud = 3.94 ms-1
And,
VB = VS√nc
18
VB = 227.34 ms-1
The incremental Gust Load Factor is given as,
∆ n= ρuV
2(WS ) * CLα where u = Kgui , i=b,c,d.
Ude for VB gust lines = 75.34
Ude for VC gust lines = 58.34
Ude for VD gust lines = 29.17
Point B = 2.7+2.7
B= 5.4
Point C = 2.5+8.01
C= 10.51
Point D΄ = 2.6+6.02
D= 8.62
Point E = -1.52-6.02
E= -7.54
Point F = -1.52-8.01
F= -9.53
Point G = -1.52-2.7
G= --4.22
Wing loading also affects gust response, the degree to which the aircraft
is affected by turbulence and variations in air density. A small wing has less
area on which a gust can act, both of which serve to smooth the ride.
19
SAMPLE GUST ENVELOPE
Fig.4
GUST ENVELOPE FOR OUR AIRCRAFT
Fig.5
20
CHAPTER 6
CRITICAL
LOADING
PERFORMANCE
21
CRITICAL LOADING PERFORMANCE
The greatest air loads on an aircraft usually comes from the generation
of lift during high-g maneuvers. Even the fuselage is almost always structurally
sized by the lift of the wings rather than by the pressures produced directly on
the fuselage. Aircraft load factor (n) expresses the maneuvering of an aircraft as
a standard acceleration due to gravity
At lower speeds the highest load factor of an aircraft may experience is
limited by the maximum lift available. At higher speeds the maximum load
factor is limited to some arbitrary value based upon the expected use of the
aircraft. The maximum lift load factor equals 1.0 at levels flight stall speed. The
is the slowest speed at which the maximum load can be reached without
stalling.
The aircraft maximum speed, or dive speed at right of the V-n diagram
represents the maximum dynamic pressure and maximum load factor is clearly
22
important for structural sizing. At this condition, the aircraft is at fairly low
angle of attack because of the high dynamic pressure, so the load is
approximately vertical in the body axis. For a subsonic aircraft, maximum speed
is typically 50% higher than the level-flight cruise speed.
The load factor for different maneuvers found out and load factor during
critical performance like pull up, pull down etc., are to be found.
For level turn,
23
Minimum radius of turn,
Load factor at min. radius of turn,
Maximum Turn Rate,
Load factor at max. rate of turn,
For Pull up maneuver,
24
.Load factor can be found from,
For Pull down maneuver,
Load factor can be found from,
During Climb,
During Dive,
Vdive= 1.5 Vcruise
25
Substituting the other values, the load factor for different critical
performance are,
K= 0.04
MANEUVERS LOAD FACTOR (n)
Max. Turn Rate 0.53
Level Turn 1.40
Pull Up 1.54
Pull Down 1.67
Climb 4.33
Dive 2.46
26
CHAPTER 6.1
FINAL V – n
DIAGRAM
27
FINAL V – n DIAGRAM
CALCULATION:
From ADP – I, = 2.7; = 0.419
; (W/S)= 691.4kgm-2
AT CRUISE ALTITUDE = 10000 m
(-)ve n = 0.000124V2 (+)ve n = 0.000805V2
(-n) V (+n)
No unit in ms-1 No unit
0 0 0
0.12 100 0.81
0.17 118.19 1.14
0.50 200 3.28
0.67 230.39 4.35
2.02 400 13.12
8.08 800 52.51
12.64 1000 82.05
28
Fig.6
29
CHAPTER 7
STRUCTURAL
DESIGN STUDY –
THEORY
APPROACH
30
STRUCTURAL DESIGN STUDY – THEORY APPROACH
Aircraft loads are those forces and loadings applied to the airplanes
structural components to establish the strength level of the complete airplane.
These loadings may be caused by air pressure, inertia forces, or ground
reactions during landing. In more specialized cases, design loadings may be
imposed during other operations such as catapulted take-offs, arrested landings,
or landings in water.
The determination of design loads involves a study of the air pressures
and inertia forces during certain prescribed maneuvers, either in the air or on the
ground. Since the primary objective is an airplane with a satisfactory strength
level, the means by which this result is obtained is sometimes unimportant.
Some of the prescribed maneuvers are therefore arbitrary and empirical which is
indicated by a careful examination of some of the criteria.
Important consideration in determining the extent of the load analysis is
the amount of structural weight involved. A fairly detailed analysis may be
necessary when computing operating loads on such items as movable surfaces,
doors, landing gears, etc. proper operation of the system requires an accurate
prediction of the loads.
Aircraft loads is the science of determining the loads that an aircraft
structure must be designed to withstand. A large part of the forces that make up
design loads are the forces resulting from the flow of air about the airplane’s
surfaces- the same forces that enable flight and control of the aircraft.
Load factors
In normal straight and level flight the wing lift supports the weight of the
airplane. During maneuvers or flight through turbulent (gusty) air, however,
additional loads are imposed which will increase or decrease the net loads on
the airplane structure. The amount of additional loads depends on the severity of
31
the maneuvers or the turbulence, and its magnitude is measured in terms of load
factor.
The maximum maneuvering load factor to which an airplane is designed
depends on its intended usage. Fighters, which are expected to execute violent
maneuvers, are designed to withstand loads commensurate with the
accelerations a pilot can physically withstand. Long range, heavily loaded
bombers, on the other hand, are designed to low load factors and must be
handled accordingly.
For a typical two spar layout, the ribs are usually formed in three parts
from sheet metal by the use of presses and dies. Flanges are incorporated around
the edges so that they can be riveted to the skin and the spar webs Cut-outs are
necessary around the edges to allow for the stringers to pass through Lightening
holes are usually cut into the rib bodies to reduce the rib weight and also allow
for passage of control runs fuel electrics etc.
32
STRUCTURAL DESIGN CRITERIA
The structural criteria define the types of maneuvers, speed, useful loads,
and gross weights which are to be considered for structural design analysis.
These are items which are under the control of the airplane operator. In
addition, the structural criteria must consider such items as inadvertent
maneuvers, effects of turbulent air, and severity of ground contact during
landing. The basic structural design criteria, from which the loadings are
determined, are based largely on the type of the airplane and its intended use.
33
CHAPTER 8
LOAD
ESTIMATION OF
WINGS
34
AIR LOADS ON WING
With the V-n diagram complete, the actual loads and load distribution on
the wing can be determined. Before the actual structural members can be sized
and analyzed, the loads they will sustain must be determined. Aircraft loads
estimation, a separate discipline of aerospace engineering, combines
aerodynamics, structures and weights.
Initially we have to calculate the lift produced by the wings. Once the lift
on the wings is known, the span-wise and chord-wise load distributions can be
determined.
According to classical wing theory, the span wise lift or load distribution
is proportional to the circulation at each station. A vortex lifting –line
calculation will yield the span-wise lift distribution. For an elliptical plan form
wing, the lift and load distributions is of elliptical shape.
For a non-elliptical wing, a good semi-empirical method for span-wise
load estimation is known as Schrenk’s Approximation. This method assumes
that the load distribution on an untwisted wing has a shape that is the average of
the actual plan-form shape and an elliptical shape of same and area. The total
area under the lift load curve must sum to the required total lift.
Fig.7
35
To find the lift distribution in aircraft wing, the following procedure is followed:
1) Plan-form shape wing is plotted.
2) Elliptic distribution is drawn using the formula
CALCULATIONSWe know,
Where; a = Spanofthewing
2
a = 88.4/2 = 44.2 mπ * 44.2 * b / 4= 899.75/2b = 12.96 m
To construct the ellipse,
Using the above equation, for various values of x, the values of y are found and the ellipse is drawn
SPANWISE (m)
CHORDWISE (m)
0 12.963 12.936 12.849 12.6812 12.4715 12.1918 11.8321 11.4024 10.8827 10.2630 9.5133 8.6236 7.5139 6.0942 4.03
44.2 0
36
Fig.8
The load intensity at each grid point on the wing plan-form is calculated as
follows.
Load intensity at root = W2
× y0
Areaunderthecurve
Where y0 is the lift distribution at the root
Load intensity at root = (618488.15 X 12.96)/380.16 =21084.82N/m
Area of Schrenk’s curve= (2/3) x (12.96x44) = 380.16 m2
Load at any location ‘n’ = Load Intensity at root × yn
y0
Where yn is the lift distribution at the corresponding grid point
Lift on each element is calculated using the following formula and a graph is
plotted between lift on element and wing span.
37
Lift on element = Load intensity at grid point * Distance between two
grid points
Structural load of the wing, W Wing = ∫0
b2
K CX2 dx
CX = A + Bx where CX is the chord at each station
At x = 0, CX = CR = 16.28 m
At x = b2 , CX = CTIP = 4.07 m
Using the above conditions, we get,
A = 16.28; B = - 0.27
CX = 16.28 – 0.27x
To find the value of K, first the total structural weight of the wing is taken as the
WING LOAD.
Wwing=C1 C2 C3 WdgC
4 nC5 Sw
C6 AC
7 (t/c)C8 (C9+)C
10 (cos)C11 Sf
C12 qC
13 WfwC
14
A being the aspect ratio of the wing
n being the load factor
q being the dynamic pressure
Sw being the planform area of the main wing
Sf being the planform area of flapped portion of the main wing
t/c being the max.thickness-to-chord ratio of the wing
Wdg being the design gross weight
Wfw being the weight of fuel stored in the wing
being the sweep angle of the max.thickness
being the taper ratio
Type C1 C2 C3 C4 C5 C6 C7 C8 C9 C10 C11 C12 C13 C14
Fighter 0.0103 Kdw Kvs 0.500 0.500 0.622 0.785 -0.4 1.0 0.050 -1.0 0.04 0 0
38
Transport 0.0051 1 1 0.557 0.557 0.649 0.500 -0.4 1.0 0.100 -1.0 0.10 0 0
Gen.
Aviation0.0090 1 1 0.490 0.490 0.758 0.600 -0.3 0 0.004 -0.9 0 0.006 0.0035
W Wing = ∫0
b2
K CX2 dx= 37942.46 N
On solving the above equation, we get K = 7.26
Using the above value of K, the wing structural loads at other locations
are calculated and tabulated.
The resultant can be found as the difference between the structural load
intensity and lift load intensity
Resultant Load Intensity = Structural Load Intensity – Lift Load
Intensity
Span(m)Chord(m
)
LIFT LOAD
INTENSITY(N/m)
LIFT ON
ELEMEN
T (N)
STRUCTURAL
LOAD
INTENSITY
(N/m)
RESULTANT
LOAD
INTENSITY(N/m
)
0 12.96 21084.82 632.54 37935.18 16850.36
3 12.93 21036.10 1893.25 37442.56 16406.46
6 12.84 20889.59 3342.33 36768.01 15878.42
9 12.68 20629.28 4332.14 36081.37 15452.09
12 12.47 20287.63 5680.53 34758.83 14471.2
15 12.19 19832.09 7139.55 32809.92 12977.83
18 11.83 19246.40 8275.95 31679.00 12342.6
21 11.40 18546.83 9644.35 29314.46 10767.63
24 10.88 17700.83 10974.51 27299.89 9599.06
27 10.26 16692.15 12519.11 24551.65 7859.5
30 9.51 15471.96 13770.04 21392.86 5920.9
33 8.62 14024.01 15566.65 17794.94 3770.93
39
36 7.51 12218.13 17349.74 13729.27 1511.14
39 6.09 9907.91 20140.29 9167.3 -740.61
42 4.03 6556.46 26422.53 4080.48 -2475.98
Fig.9
40
Fig.10
41
Fig.11
From the above graphs, it can be inferred that all the three parameters
decrease along the span of the wing.
SHEAR FORCE AND BENDING MOMENT DIAGRAM
To determine the shear force and bending moment diagram for the wing
we assume that the wing is a cantilever beam with the root end fixed while the
tail end is free.
For a cantilever beam the shear force is a given by,
Tabulation for the values of shear force and bending moment at various
positions along the span is as follows.
SPAN (m)
RESULTANT LOAD INTENSITY (Nm-1)
SHEAR FORCE(N)
BENDING MOMENT(Nm)
0 16850.36 744785.91 16459768.61
3 16406.46 675946.15 13924490.69
42
Shear Force = Rx
Bending Moment = Rx2/2
6 15878.42 606555.64 11585212.72
9 15452.09 543913.56 9572878.65
12 14471.2 465972.64 7502159.50
15 12977.83 378952.63 5532708.39
18 12342.6 323376.12 4236227.17
21 10767.63 249809.01 2897784.51
24 9599.06 193901.01 1958400.20
27 7859.5 135183.4 1162577.24
30 5920.9 84076.78 596945.13
33 3770.93 42234.41 236512.69
36 1511.14 12391.34 50804.49
39 -740.61 -3851.17 -10013.04
42 -2475.98 -9298.32 -10228.15
Fig.12
43
Fig.13
LOAD ESTIMATION OF WINGS
WING STRUCTURAL LAYOUT
Specific Roles of Wing (Mainwing) Structure:
The specified structural roles of the wing (or main plane) are:
To transmit: wing lift to the root via the main span wise beam Inertia loads from the power plants, undercarriage, etc., to the main
beam. Aerodynamic loads generated on the aerofoil, control surfaces & flaps
to the main beam.
To react against: Landing loads at attachment points Loads from pylons/stores Wing drag and thrust loads
To provide:
44
Fuel tank age space Torsional rigidity to satisfy stiffness and aero-elastic requirements.
To fulfill these specific roles, a wing layout will conventionally
compromise:
Span wise members (known as spars or booms) Chord wise members(ribs) A covering skin Stringers
Basic Functions of Wing Structural Members
The structural functions of each of these types of members may be
considered independently as:
SPARS
Form the main span wise beam Transmit bending and torsional loads Produce a closed-cell structure to provide resistance to torsion, shear and
tension loads.
45
In particular:
Webs – resist shear and torsional loads and help to stabilize the skin. Flanges - resist the compressive loads caused by wing bending.
SKIN
To form impermeable aerodynamics surface Transmit aerodynamic forces to ribs & stringers Resist shear torsion loads (with spar webs). React axial bending loads (with stringers).
STRINGERS
Increase skin panel buckling strength by dividing into smaller length sections.
React axial bending loads
RIBS
Maintain the aerodynamic shape Act along with the skin to resist the distributed aerodynamic pressure
loads Distribute concentrated loads into the structure & redistribute stress
around any discontinuities Increase the column buckling strength of the stringers through end
restraint Increase the skin panel buckling strength.
46
STRUCTURAL ANALYSIS OF WING
Wing is the major lift producing surface. Therefore, the analysis has to be
very accurate. The structural analysis of the wing by defining the primary load
carrying member Spars is done below.
The configuration used in our aircraft is the Box Beam (distributed
flange) concept-built-up or integral construction.
This method is more suitable for aircraft wings with medium to high load
intensities and differs from the mass boom concept in that the upper and lower
skins also contribute to the span wise bending resistance
Another difference is that the concept incorporates span wise stringers
(usually “z” section) to support the highly –stressed skin panel area. The
resultant use of a large number of end-load carrying members improves the
overall structural damage tolerance.
The concept described above is commonly known as built-up
construction method. The concept is simple in that the skin-stringer panels are
47
manufactured singly from large billets of metal. Advantages of the integral
construction method over the traditional built-up method include:
Simpler construction & assembly Reduced sealing/jointing problems Reduced overall assembly time/costs Improved possibility to use optimized panel tapering.
This configuration is used in order to offer provision for our High Load
Intensity of 22000 Nm-1.
SPAR DEFINITION:
The maximum bending moment from previous section was found to be as
2897784.51 Nm. Therefore we define 3 Spars with front spar at 15% of
chord, middle spar at 45% of chord and rear spar at 70% of chord. The
position of the three spars from the leading edge of the root chord is given
below as follows:
Front spar - 15% of chord = 2.442 m
Middle spar - 45% of chord = 7.326 m
Rear spar - 70% of chord = 11.396 m
Bending moment M = Max BM * FOS * n
48
= 2897784.51 × 1.5 × 3.8
= 16517371.71Nm
The Structural load bearing members in the wing are the Spars and
Stringers. The bending moment carried by the Spars is 70% and that of
Stringers is 30% of the total Bending Moment.
Bending Moment taken by Spars is = 0.7 x 16517371.71 = 11562160.19 Nm
The cross section of the spar chosen here is an I-section.
For each spar we are determining the following parameters:
1) Centroid
2) Moment of Inertia
3) Bending Moment
4) Bending Stress
FRONT SPAR
Height of the spar = 38 cm
Breadth of the spar = 16 cm
Thickness of the spar = 4.5 cm
49
Cross Section of Front Spar
To find out the centroid, the following calculations are made.
Element
Area(A)
(cm2)
x(
cm)
y(cm)
Ax(cm3)
Ay(cm3)
Ax2
(cm4)Ay2
(cm4)Icx
(cm4)Icy
(cm4)
1 72 8 2.25 576 162 4608 364.5 121.5 1536
2 130.5 8 19 10442479.
58352 47110.5 9145.8 220.22
3 72 835.7
5576 2574 4608 92020.5 121.5 1536
Total 274.5 2196 5215. 1756 139495. 9388.8 3292.2
50
5 8 5 7 2
Front Spar Calculations
Centroid = X = ΣAxΣA = 8 cm ; Y=
ΣAyΣA = 19 cm
I xx= Σ Icx + ΣAy2 – ΣAY2
I xx = (9388.87) + (139495.5) – (274.5)(19)2
I xx = 49789.88 cm4
I yy = Σ Icy+ ΣAx2 – ΣA X2
I yy = (3292.22) + (17568) – (274.5)(8)2
I yy = 3292.22 cm4
The FRONT SPAR carries 35 % of the BM carried by the Spars. Thus,
Front spar BM = 0.35 x 1156216019 N-cm
= 404675606.7 N cm
Front Spar Bending Stress
Bending Stress, σ z = (M x / I xx) y
POINTSCOORDINATES (y)
(cm)BENDING STRESS (N/cm2)
A 19 154425.68B 14.5 117851.18C 14.5 117851.18D -14.5 -117851.18E -14.5 -117851.18F -19 -154425.68
The bending stress at various points whose co-ordinates are determined
with centroid as the origin are calculated from above formula and tabulated.
51
Bending Stress diagram for I-Section
MIDDLE SPAR:
Height of the spar = 41.6 cm
Breadth of the spar = 18 cm
Thickness of the spar = 5 cm
Cross Section of Middle Spar
To find out the center of gravity, the following calculations are made:
Area(A x y Ax Ay Ax2 Ay2 Icx Icy
52
Element
) (cm2)
(cm)
(cm)
(cm3) (cm3) (cm4) (cm4) (cm4) (cm4)
1 90 9 2.5 810 225 7290 562.5 187.5 2430
2 158 920.8
14223286.
41279
868357.1
213147.
7329.17
3 90 939.1
810 3519 7290137592.
9187.5 2430
Total 338 30427030.
42737
8206512.
513522.
75189.1
7
Middle Spar Calculations
Centroid = X = ΣAxΣA = 9 cm ; Y=
ΣAyΣA = 20.8 cm
I xx= Σ Icx + ΣAy2 – ΣAY2
I xx = (13522.7) + (206512.5) – (338)(20.8)2
I xx = 60467.7 cm4
I yy = ΣIcy+ ΣAx2 – ΣA X2
I yy = (5189.17) + (27378) – (338)(9)2
I yy = 5189.17 cm4
The bending moment carried by the middle spar is 40% of the total bending
moment carried by the spars.
Middle Spar BM = 462486407.6 N-cm
Bending Stress σ z = (M x / I xx) y
The bending stress at various points whose co-ordinates are determined
with centroid as the origin are calculated from above formula and tabulated:
POINTS COORDINATES (y) (cm) BENDING STRESS (N/cm2)A 20.8 159088.52B 15.8 120846.09C 15.8 120846.09D -15.8 -120846.09E -15.8 -120846.09
53
F -20.8 -159088.52
REAR SPAR
Height of the spar = 17.72 cm
Breadth of the spar = 7.6 cm
Thickness of the spar = 2.5 cm
Cross Section of Rear Spar
To find out the centroid, the following calculations are made:
Element
Area(A) (cm2)
x(
cm)
y(cm)
Ax(cm3)
Ay(cm3)
Ax2
(cm4)Ay2
(cm4)Icx
(cm4)Icy
(cm4)
1 19 3.8 1.25 72.2 23.75274.3
629.6875 9.896 91.45
2 31.8 3.8 8.86 120.8 281.74 459.1 2496.28 428.76 16.56
54
4 8 9 7
3 19 3.816.4
772.2 312.93
274.36
5153.957
9.896 91.45
Total 69.8265.2
4618.42
81007.
97679.93
2448.55
2199.4
6
Rear Spar Calculations
Centroid = X = ΣAxΣA = 3.8 cm ; Y=
ΣAyΣA = 8.86 cm
I xx= Σ I cx + ΣAy2 – ΣAY2
I xx = (448.552) + (7679.932) – (69.8)(8.86)2
I xx = 2649.184 cm4
I yy = Σ Icy+ ΣAx2 – ΣA X2
I yy = (199.46) + (1007.9) – (69.8)(3.8)2
I yy = 199.46 cm4
Rear Spar carries 25 % of the spar Bending Moments.
Rear Spar Bending Moment = 289054004.8 N-cm
Bending Stress σ z = (Mx/Ixx)y
The bending stresses at various points are obtained as:
Rear Spar Bending Stress
POINTS COORDINATES (y) (cm) BENDING STRESS (N/cm2)A 8.86 966719.74B 6.36 693943.29C 6.36 693943.29D -6.36 -693943.29E -6.36 -693943.29F -8.86 -966719.74
55
CHAPTER 9
LOAD ESTIMATION OF FUSELAGE
56
LOAD ESTIMATION OF FUSELAGE
FUSELAGE STRUCTURAL LAYOUT
The fuselage is the main structure, or body, of the aircraft.
It provides space for personnel, cargo, controls, and most of the
accessories. The power plant, wings, stabilizers, and landing
gear are attached to it.
There are two general types of fuselage construction—
welded steel truss and monocoque designs. The welded steel
truss was used in smaller Navy aircraft, and it is still being used
in some helicopters.
The monocoque design relies largely on the strength of
the skin, or covering, to carry various loads. The monocoque
design may be divided into three classes - monocoque,
semimonocoque and reinforced shell.
57
The true monocoque construction uses formers, frame
assemblies, and bulkheads to give shape to the fuselage.
However, the skin carries the primary stresses. Since no
bracing members are present, the skin must be strong
enough to keep the fuselage rigid. The biggest problem in
monocoque construction is maintaining enough strength
while keeping the weight within limits.
Semimonocoque design overcomes the strength-to-weight
problem of monocoque construction. In addition to having
formers, frame assemblies, and bulkheads, the
semimonocoque construction has the skin reinforced by
longitudinal members.
58
The reinforced shell has the skin reinforced by a complete
framework of structural members.
Different portions of the same fuselage may belong to any one
of the three classes. Most are considered to be of
semimonocoque-type construction.
The semimonocoque fuselage is constructed primarily of
aluminum alloy, although steel and titanium are found in high-
temperature areas. Primary bending loads are taken by the
longerons, which usually extend across several points of
support. The longerons are supplemented by other longitudinal
members known as stringers. Stringers are more numerous and
lightweight than longerons.
The vertical structural members are referred to as
bulkheads, frames, and formers. The heavier vertical members
are located at intervals to allow for concentrated loads. These
members are also found at points where fittings are used to
attach other units, such as the wings and stabilizers.
The stringers are smaller and lighter than longerons and
serve as fill-ins. They have some rigidity but are chiefly used for
giving shape and for attachment of skin. The strong, heavy
longerons hold the bulkheads and formers. The bulkheads and
formers hold the stringers. All of these join together to form a
rigid fuselage framework. Stringers and longerons prevent
tension and compression stresses from bending the fuselage.
The skin is attached to the longerons, bulkheads, and
other structural members and carries part of the load. The
fuselage skin thickness varies with the load carried and the
stresses sustained at particular location.
59
There are a number of advantages in using the
semimonocoque fuselage.
The bulkhead, frames, stringers, and longerons aid in the
design and construction of a streamlined fuselage. They
add to the strength and rigidity of the structure.
The main advantage of the semimonocoque construction
is that it depends on many structural members for
strength and rigidity. Because of its stressed skin
construction, a semimonocoque fuselage can withstand
damage and still be strong enough to hold together.
FUSELAGE STRUCTURAL ANALYSIS
60
Structural analysis of fuselage like that of wing is of prime importance
while designing an aircraft. As the fuselage is the one which houses the pilot,
the power plant and also part of the payload its structural integrity is a matter of
concern. While analyzing the fuselage structure the section must be idealized.
Idealization involves the conversion of a stringer and its accompanying skin
thickness into a concentrated mass known as a boom. The shear flow analysis
of the fuselage simulating flight conditions is shown below.
(a) Actual fuselage section; (b) idealized fuselage section
The stringer used is of Z type. The following are its dimensions
Cross sectional area of each stringer is 100 mm2
Cross section of Z-section
61
The above stringer section is uniformly used throughout the fuselage as
shown above in order to provide the fuselage the required load carrying
capacity. The diagram showed adjacent is of the idealized fuselage structure.
The idealization process is carried out in the following way.
STRESS ANALYSIS:
IDEALIZATION:
The boom 1 is given by
Where
B1 = Area of Boom 1
tD = Thickness of skin panel
b = Circumferential distance between 2 stringers
By Symmetry,
B1 = B9, B2 = B8, B10 = B16, B3 = B7 , B11 = B15, B4 = B6 = Bl2 = B14 ,B5 = B13
B1=100+(0.65*1.37*106/6)[2+(4110/5500)]+(0.65*1.37*106/6)[2+(4110/5500)]
=815582.12 mm2
Similarly for boom 2 ,
B2 =815582.12 mm2
Similarly B3 = 815582.12 mm2, B4 =815582.12 mm2. We note that stringers 5 and 13 lie on the neutral axis of the section and are therefore unstressed; the calculation of boom areas B5 and B13 does not then arise.
Thus, we have B1:B16=815582.12 mm2
62
We know that,
Ixx = By2
Ixx1=24.67 m4; Ixx2=13.77 m4; Ixx3=6.12 m4; Ixx4=1.11 m4
Maximum bending moment = 2897784.51 Nm
Hence the Bending moment acting on the fuselage M = Max.B.M * n* FOS
=2897784.51 * 3.8*1.5
=16517371.71 Nm
Ixx = 24.67 m4
The value of stress acting is given by the expression :
= (16517371.71 *y)/24.67
Stress in Stringers
STRINGER/BOOM Y (m) STRESS x106 (Nm−2)
1 5.5 3.68
2, 16 4.11 2.75
3, 15 2.74 1.83
4, 14 1.37 0.9
5, 13 0 0
6, 12 -1.37 -0.9
7, 11 -2.74 -1.83
8, 10 -4.11 -2.75
9 -5.5 -3.68
63
CHAPTER 10
BALANCING AND
MANEUVERING LOADS
ON TAIL PLANE, RUDDER
AND AILERON LOADS
64
Maneuvering loads.
Each horizontal surface and its supporting structure, and the main wing of a
canard or tandem wing configuration, if that surface has pitch control, must be
designed for the maneuvering loads imposed by the following conditions:
A sudden movement of the pitching control, at the speed VA, to the
maximum aft movement, and the maximum forward movement, as
limited by the control stops, or pilot effort, whichever is critical.
A sudden aft movement of the pitching control at speeds above VA,
followed by a forward movement of the pitching control resulting in the
following combinations of normal and angular acceleration
At speeds up to VA, the vertical surfaces must be designed to withstand the
following conditions. In computing the loads, the yawing velocity may be
assumed to be zero:
With the airplane in unaccelerated flight at zero yaw, it is assumed that
the rudder control is suddenly displaced to the maximum deflection, as
limited by the control stops or by limit pilot forces.
With the rudder deflected ,it is assumed that the airplane yaws to the
overswing sideslip angle. In lieu of a rational analysis, an overswing
angle equal to 1.5 times the static sideslip angle may be assumed.
(A yaw angle of 15 degrees with the rudder control maintained in the
neutral position (except as limited by pilot strength).
The airplane must be yawed to the largest attainable steady state sideslip
angle, with the rudder at maximum deflection caused by any one of the
following:
65
i. Control surface stops;
ii. Maximum available booster effort;
iii. Maximum pilot rudder force as shown below:
The rudder must be suddenly displaced from the maximum deflection to the
neutral position.
The yaw angles may be reduced if the yaw angle chosen for a particular
speed cannot be exceeded in--
i. Steady slip conditions;
ii. Uncoordinated rolls from steep banks; or
iii. Sudden failure of the critical engine with delayed corrective action.
The ailerons must be designed for the loads to which they are subjected--
In the neutral position during symmetrical flight conditions; and
By the following deflections (except as limited by pilot effort), during
unsymmetrical flight conditions:
i. Sudden maximum displacement of the aileron control at VA.
Suitable allowance may be made for control system deflections.
ii. Sufficient deflection at VC, where VC is more than VA, to produce
a rate of roll not less than obtained
iii. Sufficient deflection at VD to produce a rate of roll not less than
one-third of that obtained
66
(a)Symmetric maneuvering conditions:
For the analysis of the maneuvering flight conditions specified in
paragraphs (b) and (c) of this section, the following provisions apply:
Where sudden displacement of a control is specified, the assumed rate of
control surface displacement may not be less than the rate that could be
applied by the pilot through the control system.
In determining elevator angles and chordwise load distribution in the
maneuvering conditions, the effect of corresponding pitching velocities
must be taken into account. The in-trim and out-of-trim flight conditions
must be considered.
(b) Maneuvering balanced conditions:
Assuming the airplane to be in equilibrium with zero pitching acceleration, the
maneuvering conditions A through I on the maneuvering envelope must be
investigated.
(c) Pitch maneuver conditions:
The conditions specified in paragraphs (c) must be investigated. The movement
of the pitch control surfaces may be adjusted to take into account limitations
imposed by the maximum pilot effort, control system stops and any indirect
effect imposed by limitations in the output side of the control system (for
example, stalling torque or maximum rate obtainable by a power control
system.
67
Maximum pitch control displacement at VA.:
The airplane is assumed to be flying in steady level flight and the cockpit pitch
control is suddenly moved to obtain extreme nose up pitching acceleration. In
defining the tail load, the response of the airplane must be taken into account.
Airplane loads that occur subsequent to the time when normal acceleration at
the c.g. exceeds the positive limit maneuvering load or the resulting tailplane
normal load reaches its maximum, whichever occurs first, need not be
considered.
(2) Specified control displacement:
A checked maneuver, based on a rational pitching control motion vs. time
profile, must be established in which the design limit load factor will not be
exceeded. Unless lesser values cannot be exceeded, the airplane response must
result in pitching accelerations not less than the following:
A positive pitching acceleration (nose up) is assumed to be reached
concurrently with the airplane load factor of 1.0. The positive
acceleration must be equal to at least
where-
n is the positive load factor at the speed under consideration; and V is the
airplane equivalent speed in knots.
A negative pitching acceleration (nose down) is assumed to be reached
concurrently with the positive maneuvering load factor. This negative
pitching acceleration must be equal to at least
where-
68
n is the positive load factor at the speed under consideration; and V is the
airplane equivalent speed in knots.
Balancing loads
A horizontal surface balancing load is a load necessary to maintain
equilibrium in any specified flight condition with no pitching
acceleration.
Horizontal balancing surfaces must be designed for the balancing loads
occurring at any point on the limit maneuvering envelope and in the flap
conditions
It is not require to balance the rudder because it will not deflect due to
gravity.
Aileron will defect in vice versa direction so it is doesn’t require
balancing load.
69
CHAPTER 11
DESIGN OF
COMPONENTS OF
THE WING
70
DESIGN OF COMPONENTS OF THE WING
FUEL TANKS
Aircraft typically use three types of fuel tanks: integral, rigid removable, and
bladder.
Integral tanks are areas inside the aircraft structure that have been sealed
to allow fuel storage. An example of this type is the "wet wing"
commonly used in larger aircraft. Since these tanks are part of the aircraft
structure, they cannot be removed for service or inspection. Inspection
panels must be provided to allow internal inspection, repair, and overall
servicing of the tank. Most large transport aircraft use this system, storing
fuel in the wings, belly, and sometimes tail of the airplane.
Rigid removable tanks are installed in a compartment designed to
accommodate the tank. They are typically of metal construction, and may
be removed for inspection, replacement, or repair. The aircraft does not
rely on the tank for structural integrity. These tanks are commonly found
in smaller general aviation aircraft, such as the Cessna 172.
Bladder tanks are reinforced rubberized bags installed in a section of
aircraft structure designed to accommodate the weight of the fuel. The
bladder is rolled up and installed into the compartment through the fuel
filler neck or access panel, and is secured by means of metal buttons or
snaps inside the compartment. Many high-performance light aircraft and
some smaller turboprops use bladder tanks. One major down-side to this
type of tank is the tendency for materials to work harden through
extensive use making them brittle causing cracks. One major plus side is
the ability to utilise as much of the aircraft as possible to store fuel.
Combat aircraft generally use self-sealing fuel tanks.
71
The study of the fuel tanks and its corresponding feed systems is of
importance to us because, as the fuel is stored in the wings and also
externally. Thus structural analysis results are incorporated here to ensure
that the fuel tanks are positioned suitably and are adequate in capacity so that
our aircraft has an endurance of 8 hours. The C.G analysis carried out in
ADP-I let us know that our aircraft’s C.G remains within acceptable limits
even on consumption of fuel consumption.
Integral tanks are areas inside the aircraft structure that have been sealed
to allow fuel storage. An example of this type is the "wet wing" commonly
used in larger aircraft. Since these tanks are part of the aircraft structure, they
cannot be removed for service or inspection. Inspection panels must be
provided to allow internal inspection, repair, and overall servicing of the
tank. Most large transport aircraft use this system, storing fuel in the wings,
belly, and sometimes tail of the airplane.
Internally, the fuel is carried in the wings. The wing tanks are capable of
withstanding up to 90000 kg of fuel. Externally however we had two
choices, conformal tanks or drop tanks. Drop tanks are heavy and are an
aerodynamic liability (drag).
72
RIB LOCATION AND DIRECTION
The span-wise location of ribs is of some consequence. Ideally, the rib
spacing should be determined to ensure adequate overall buckling support to the
distributed flanges. This requirement may be considered to give a maximum
pitch of the ribs. In practice other considerations are likely to determine the
actual rib locations such as:
Hinge positions for control surfaces and attachment/operating points for
flaps, slats, and spoilers.
Attachment locations of power plants, stores and landing gear structure.
A need to prevent or postpone skin local shear or compression buckling,
as opposed to overall buckling. This is especially true in a mass boom
form of construction.
Ends of integral fuel tanks where a closing rib is required. When the wing
is unswept, it is usual for the ribs to be arranged in the flight direction and
thereby define the aerofoil section. While the unswept wing does give
torsional stiffness, the ribs are heavier, connections are more complex and
in general the disadvantages overweigh the gains.
FIXED SECONDARY STRUCTURE
A fixed leading edge is often stiffened by a large number of closely
pitched ribs, span-wise members being absent. Providing care is taken in the
detail design of the skin attachment it is possible to arrange for little span-wise
end load to be diffused into the leading edge and buckling of the relatively light
structure is avoided. This may imply short spam-wise sections. The presence of
thermal de-icing, high-lift devices or other installations in the leading edge also
has a considerable influence upon the detail design. Bird strike considerations
are likely to be important.
73
HORIZONTAL STABILISER
When the horizontal stabilizer is constructed as a single component
across the centerline of the aircraft, the basic structural requirements are very
similar to those of a wing.
VERTICAL STABILISER
The vertical stabilizer presents a set of issues which are different from
those of the main plane or horizontal stabilizer. Relevant matters are:
It is not unusual to build the vertical stabilizer integrally with the rear
fuselage. The spars are extended to form fuselage frames or bulkheads.
A ‘root’ rib is made to coincide with the upper surface of the fuselage
and is used to transmit the fin root skin shears directly into the fuselage
skin. Fin span-wise bending results in fuselage torsion. Often it is logical
to incline the rear spar bulkhead to continue the line of the rear spar
since it is usually the end of the main fuselage structure. On the other
hand, the front spar and any intermediate attachment frames are often
best kept perpendicular to fuselage fore and aft datum. The change in
direction being made at the fin root rib. Otherwise the structural form
can follow that of a wing.
AUXILIARY SURFACES
The structural layout of the auxiliary lifting surfaces is generally similar
to that of the wing but there are differences, in part due to the smaller size and
in part due to the need to provide hinges or supports. The latter implies that
each auxiliary surface is a well-defined.
74
HINGED CONTROL SURFACES
Conventional training edge control surfaces are almost invariably
supported by a number of discrete hinges, although continuous, piano type,
hinges may be used for secondary tabs. To some degree the number and
location of the discrete hinges depends upon the length of the control. The
major points to be considered are:
a) The bending distortion of the control relative to the fixed surface
must be limited so that the nose of the control does mot fouls the
fixed shroud.
b) The control hinge loads and the resulting shear forces and bending
moments should be equalized as far as is possible.
c) Structural failure of a single hinge should be tolerated unless each
hinge is of fail-safe design and can tolerate cracking one load path.
These points suggest the use of a relatively large number of discrete
hinges but there are difficulties associated with this solution there are the
obvious loads likely to be induced in the control by the distortion under load of
the main surface to which it is attached may be significant. These problems do
not arise if only two hinge points are used as any span-wise distortion or
misalignment can be accommodated by designing one of the hinges so that it
can rotate about a vertical axis. When more than two hinges are used the
‘floating’ hinge concept
PIVOTED CONTROL SURFACES
In certain high-performance aircraft, the whole of a stabilizing or
control surface on one side of the aircraft may be pivot about a point on its root
chord. Clearly in this case, the structural considerations are dominated by the
need to react all the forces and moments at the pivot and operating points. Thus
the structural layout may consist of an integral root rib or pivot or stub spar
arrangement to which is attached a number of shear webs fanning out towards
75
the extremities of the surface, possibly in conjunction with full depth
honeycomb. High skin shear loading is inevitable due to the need to bring the
loads to the two concentrated points. Shear loads due to torsion may be limited
by locating the operating point on the root rib some distance away from the
pivot.
HIGH LIFT SYSTEMS
There is a wide variety of leading and trailing edge high-lift systems.
Some types are simply hinged to the wing, but many require some degree of
chord-wise extension. This can be achieved by utilizing a linkage, a mechanism,
a pivot located outside the aerofoil contour or, perhaps most commonly, by
some form of track. Trailing edge flaps may consist of two or more separate
chord-wise segments, or slats, to give a slotted surface and these often move on
tracts attached to the main wing structure.
The majority of flaps and slats are split into span wise segments of no
greater lengths than can be supported at two or three locations. As with control
surfaces, the locations of the support points are established so as to minimize
local deformations since the various slots are critical in determining the
aerodynamic performance. Sometimes the actuation may be located at a
different pan wise position from the support points. This is often a matter of
convenience, layout clearances, and the like.
ATTACHMENT OF LIFTING SURFACES
The joint of the fuselage with the wing is subjected to heavy load inputs and
there is a potential for considerable relative distortion. This distortion is usually
accepted and the wing center box is built completely into the fuselage, the
resulting constraint stresses being allowed for. It is usual for the wing structure
of large aircraft to include a production joint at the side of the fuselage and this
is virtual essential for swept wings.
76
DESIGN OF
COMPONENTS OF
THE FUSELAGE
77
CHAPTER 12 DESIGN OF LANDING GEAR
We have designed the landing gear characteristics by following a step by
stepmethod.
1)Landing gear System
We have chosen a Retractable system landing gear which will be
retracted in to the fuselage after the take off.
2)Landing Gear Configuration
The landing gear configuration we have adapted is the Tri-cycle type
with a nose wheel in front. From an ease of ground manoeuvring viewpoint as
well as ground looping the nose wheel configuration is preferred.
3)Preliminary landing gear strut disposition
There are two geometric criteria which are required to be considered on
decidingthe disposition of landing gear struts are:
A)Tip-over criteria
B)Ground clearance criteria
A)Tip-over Criteria :
a)Longitudinal Tip-over Criterion :
For tricycle gears the main landing gear must be behind the aft CG
location. The 15 deg angles shown in the Fig. represents the usual relation
between main gear and the aft CG.
78
b)Lateral Tip-over Criterion :
The lateral tip-over is dictated by the angle ψ in the Fig.
B)Ground Clearance Criterion :
a)Longitudinal Ground Clearance Criterion :
b)Lateral Ground Clearance Criterion :
79
4) Number of Wheels :
Nose landing gear-2
Main landing gear-14
Angles of Pitch and Roll during Takeoff and Landing
The available pitch angle (θ) 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 geometric
limitation to the roll angle (φ) could result in undesirable operational limit under
cross-wind landing condition.
For a given aircraft geometry and gear height (hg), the limit for the
takeoff/landing pitch angle follows directly from the roll angle at which the tip
of the wing just touches the ground is calculated using the expression.
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) 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,
and s the distance between the engines.
80
Pitch Angle Required for Liftoff
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
Where αLOF is the highest angle of attack anticipated for normal
operational use, VLOF ist he liftoff speed, g is the gravitational acceleration CL,
LOF is the lift coefficient, and dCL/dα is the lift-curve slope.. For large transports,
the typical value for the rate of rotation (dθ/dt) is taken as four degrees per
second
Pitch and Roll Angles during Landing
With the flaps in the fully-deflected position, the critical angle of attack
of the wing during landing is smaller than in takeoff. Consequently, the pitch
angle during landing is generally less than that during takeoff. In the absence of
detailed information, the pitch angle on touchdown (θTD) may be assumed equal
to θLOF. As for the roll angle upon touchdown, an upper limit of between five
and eight degrees is generally applied to large transport aircraft.
81
CHAPTER 13 TIRE SELECTION
Aircraft tires are designed to withstand extremely heavy loads for short
durations. The number of tires required for aircraft increases with the weight of
the plane (because the weight of the airplane is distributed better). Aircraft tire
tread patterns are designed to facilitate stability in high crosswind conditions, to
channel water away to prevent hydroplaning, and for braking effect. Aircraft
tires are usually inflated with nitrogen or helium in order to minimize expansion
and contraction from extreme changes in ambient temperature and pressure
experienced during flight
Tire Sizing:
Nearly 90% of the load is carried by the main landing gear.
• Only of 10% of aircraft is carried by nose wheel. But it experience dynamic
loads.
• Nose wheel size could be 60-100% of size of main wheel.
• But in the bicycle and quarter cycle configuration the sizesame.110
Tire Description:
The main landing gear of our aircraft is set at a camber to increase load
carrying capacity. As a result, the tires’ footprints can change based on the
aircraft’s weight. Extreme operating parameters like these demand tires with
incredible durability.
Radial tires can offer low weight but tend to be less retrievable than a bias ply tyre and can exhibit weaker sidewalls.
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Tread
A specially compounded rubber formulated to resist wear, cutting, chunking and heat build-up. Most Dunlop aircraft tyre designs feature circumferential grooves moulded into the tread to disperse water from beneath the tread in wet runway conditions.
The tread helps to reduce the risk of aqua planing and improves traction and grip between the tread and runway surface.
The Casing
The basic strength of the tyre is provided by the casing plies. Casing plies are layers of fabric cord coated with hi-modulus rubber on both sides. Casing plies are held in place by being wrapped around the beads to provide the casing ply turn up.
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Beads
The bead wire anchors the tyre to the rim and ensures an airtight seal.
Beads consist of bundles of high-tensile steel wires, each strand of which is coated in rubber compound and wound into coils of the correct diameter for a given tyre size.
Chafers
Chafers are made of tough nylon material and are fitted around the bead clinch area to resist chafing damage to both tyre and rim flange.
Sidewall
The area of the tyre between the sidewalls is covered with a layer of specially formulated rubber treated with anti-oxidants. The sidewall protects the casing plies from the effects of weathering and offers resistance to cuts and flexing.
Inner Liner
Tubeless tyres have a layer of rubber bonded to the inside of the first casing ply from bead to bead to resist the permeation of nitrogen and moisture into the casing.
Undertread
The under tread is a layer of rubber that is designed to improve the adhesion between tread/ ITF and the casing plies. During the retreading process, the layer acts as the interface for the application of fresh tread rubber.
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The tire that has been selected as per the above considerations is
DUNLOP DR32518T
CharacteristicsChined NoPly Rating 18Tubed/Tubeless TLAspect Ratio 0.79Speed MPH 210Max Load lbs 22620
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Inflation & DimensionsInflation Pressure Unloaded psi 172Inflation Pressure Loaded psi 179Inflation Pressure Type StandardInf Dim Width Min 11.85Inf Dim Width Max 12.50Inf Dim Width Shoulder 10.80Inf Dim OD Min 37.20Inf Dim OD Max 38.00Inf Dim OD Shoulder 36.10
Weight And DimensionsMax Load lbs 22620Loaded Radius 16.00Rim Dim Width Between Flanges 8.66Rim Dim Ledge Diameter 18.43Rim Dim Flange Height 1.49Rim Dim Min Ledge Width 2.75
CHAPTER 14 BRAKE SYSTEM
A brake is a device which inhibits motion. Its opposite component is a
clutch. Most commonly brakes use friction to convert kinetic energy into heat,
though other methods of energy conversion may be employed. For example
regenerative braking converts much of the energy to electrical energy, which
may be stored for later use. Other methods convert kinetic energy into potential
energy in such stored forms as pressurized air or pressurized oil. Still other
braking methods even transform kinetic energy into different forms, for
example by transferring the energy to a rotating flywheel.
Almost all wheeled vehicles have a brake of some sort. Even baggage
carts and shopping carts may have them for use on a moving ramp. Most fixed-
wing aircraft are fitted with wheel brakes on the undercarriage. Some aircraft
also feature air brakes designed to reduce their speed in flight. Notable
examples include gliders and some World War II-era aircraft, primarily some
fighter aircraft and many dive bombers of the era. These allow the aircraft to
maintain a safe speed in a steep descent. The Saab B 17 dive bomber used the
deployed undercarriage as an air brake.
DETAILED DESIGN OF A SEGMENTED ROTOR DISK BRAKE
SYSTEM
Due to heavy landing loads and speeds we are using segmented rotor
brakes. They are heavy duty brakes especially adapted for high pressure
pneumatic systems. Braking can be accomplished by means of several sets of
stationary, high friction type brake linings making contact with rotating (rotor)
segments. This brake system was designed in Pro-E modeling.
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A segmented rotor brake is a multiple disk system. The Brake assembly consists
of
1. Carrier
2. Piston
3. Pressure plate
4. Rotor disk
5. Stator disk
6. Backing plate
7. Flange bolt
8. Pusher
9. Pipe
The above parts are explained in brief:
CARRIER
The Carrier is the basic unit of the brake. It is the part which is attached
to the axle. A groove is machined on the carrier to receive the piston. Air is
admitted through the pipe attached to the outside of the carrier. Inside the pipe
it consists of the pusher with a return spring. It also consists of slots to house
the pressure plate and the stator disk.
Carrier Pusher Piston
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PRESSURE PLATE
The pressure is a flat, non-rotating plate notched on the inside diameter
which fits over the slot in the carrier. It is also an auxiliary stator plate. The
Brake lining is present on one side of the plate. Ceramic brake linings are used
for more efficiency at higher temperatures, noise control, increased life and
better braking performance.
STATOR DISK
Stator disk also a non-rotating plate notched on the inside diameter to fit
on the carrier. In this brake linings are present on both sides of the disk. The
brake lining are in the form of separate multiple circular blocks to aid the
dissipation of heat. For maintaining clearance between stator and rotor an
automatic adjuster with a spring and conical pin is provided.
ROTOR DISK
The assembly consists of two rotor disk (1) in between pressure plate and
stator, (2) in between backing plate and stator. The rotor segments are notched
on the outside circumference, so that it is keyed to the wheel and rotate with it.
Many holes are provided to aid the dissipation of heat.
BACKING PLATE
It is the final unit of the assembly. It is a non-rotating plate with brake
linings present on one side. It receives the ultimate force resulting from brake
application, so it is provided with the stiffeners on the outside. It is bolted to the
front face of the carrier.
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OPERATION
During the application of brakes through the pneumatic system, the force
first acts on the pusher rod forcing the piston outward. In this process the spring
attached to the rod get compressed over the back face of the carrier. The piston
applies the force against the pressure plate which contacts the first rotor disk.
This lateral movement continues till the brake linings are in contact with
the rotor. The pressure plate, stator disk and backing plate are prevented against
rotation. Thus the non-rotating linings are all forced in contact with the rotors,
creating enough friction to stop the wheel, to which the rotors are keyed.
The function of the automatic adjustor is to maintain the correct friction
between the disks. When the stator is engaged with the rotor, the conical pin
touches the groove in the rotor. This process compresses the small spring in the
adjuster and allows the brake lining to touch the rotor disk.
Exploded view of Segmented Rotor Disk Brake System
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CHAPTER 15 FLIGHT CONTROLS
FLAPS
Our transonic cargo aircraft being of Sweep Wing design requires the use
of flaps as control surfaces as they are devoid of a horizontal stabilizer. Flaperon
is the term generally applied to wing trailing edge surfaces that combine the area
normally occupied by flaps and ailerons, and that move as a unit. A Flaperon
system will generally have a mechanical mixer that takes input from the roll
control system and flap control system, and deflects the right and left flaps
according to the combined input.
Flaps are used just like a system that has separate flaps and ailerons. When
you want to bank, you move the stick right or left. When you want to re-camber
the wing to make it more effective at different speeds, you move the flap control.
The mixer takes these inputs and makes the wing surfaces do the rest. When the
stick right and left, the entire Flaperon surface goes up on one side, and down on
the other.
Flaps are in most respects better than separate flaps and ailerons. Because
the entire trailing edge deflects on roll inputs, they provide more effective aileron
area, and consequently better roll control, than separate ailerons. And because
they entire trailing edge deflects on flap inputs, they allow for a more effective
span wise lift distribution than separate flaps. Flap settings used can be as high as
60 degrees. The one area in which flaps are an operational issue is on take-off
and landing.
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SLATS
As our aircraft have to quickly and effectively climb to assigned altitude
without stalling, we have used slats to enable our aircraft perform both climbing
at high angle and low speed cushion approach. Slats are aerodynamic surfaces on
the leading edge of the wings of fixed-wing aircraft which, when deployed, allow
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the wing to operate at a higher angle of attack. A higher coefficient of lift is
produced as a product of angle of attack and speed, so by deploying slats an
aircraft can fly more slowly or take off and land in a shorter distance
SPOILERS
Due to the high landing speeds and heavy weight of our aircraft, we have
fitted spoilers on to our aircraft. Thrust reversers are also available with our
engine. Thus spoilers are used in conjunction with thrust reversers to slow down
the aircraft while landing. A spoiler is a device intended to reduce lift in an
aircraft
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LOCATION OF CONTROL SURFACES AND OTHER COMPONENTS
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CHAPTER 16
DETAILED DESIGN
REPORT
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S. No.
Design Parameter Value Unit
1 Length (l) 84 m
2 Height (h) 18.1 m
3 Wing Span (b) 88.4 m
5 Aspect Ratio (A.R) 8.6 no unit
6 Taper ratio (𝛌) 0.25 no unit
7 Root Chord (Cr) 16.28 m
8 Mean Chord (Cm) 11.46 m
9 Tip chord (Ct) 4.07 m
10 Lift coefficient (CL) 0.419 no unit
11 Wing Area (S) 899.75 m2
12 Wing Sweep () 35 degree
13 Cruising Altitude 10000 m
14 Service Ceiling 11000 m
15 Range (R) 4000 km
16 Number of Engines 6 (no unit)
17 Engine selected ZMKB Progress D-18T
18 Maximum Thrust Capability 230 kN
19 Thrust to Weight ratio 5.7:1 no unit
20 Maximum Take Off Weight (Wo) 622125 kg
21 Empty Weight (We) 28500 kg
22 Fuel Weight (Wf) 285000 kg
23 Wing Loading (W/S) 691.4 kg/m2
24 Wing-Tail Configuration H-Tail No unit
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25 Area of the vertical stabilizer(SVT) 170.43 m2
26 Area of the aft stabilizer (SHT) 244.00 m2
27 Lift at Cruise (L(cruise)) 6097603.38 N
28 Lift at take-off (L(take-off)) 3657199.86 N
29 Lift at landing (L(landing)) 3735743.12 N
30 Drag at cruise (D(cruise)) 50220.33 N
31 Drag at take-off (D(takeoff)) 67817.67 N
32 Drag at landing (D(landing)) 52547.80 N
33 Max. Rate of Climb(R/C max ) 1.68 m/s
34 Gliding Angle () 3.814 degree
35 Take-off Runway Distance (Sto) 1454.35 m
36 Landing Runway distance (Sl) 1041.13 m
37 1-g Stall Velocity (Vs) 118.19 m/s
38 Load Factor (n) -1.52 < n < 3.8 no unit
39 Design Maneuvering Speed (Va) 230.39 m/s
40 Design Cruise Speed (Vc) 865.2 m/s
41 Design Diving Speed (Vd) 908.46 m/s
42 Max. Bending Moment (B.M) 2897784.51 Nm
43 Front spar Bending Moment 404675606.7 Ncm
44 Middle spar Bending Moment 462486407.6 Ncm
45 Rear spar Bending Moment 289054004.8 Ncm
46 Bending Moment on the fuselage 16517371.71 Nm
46 Landing GearNose landing gear-2 no unit
Main landing gear-14 no unit
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CHAPTER 18
CONCLUSION
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CONCLUSION
In conclusion, the series of cargo aircrafts incorporated many unique
design of future that were never seen on an operational aircraft. The design of
these aircrafts points the way for the design of future of very high mach
aeroplanes.
The aeroplane has gone through many design modifications since its early
conceptual designs expected, among these was a growth in weight. The
document to provide information on the trends in various aircraft characterstics
that may influence general long-term airport planning and design.
These are strong indications that future trends could see the coexistence
of very high capacity aircraft modules of similar capacities for the long
range/very long range operations. Cargo payloads, which include mail, express
and freight, are increasing in size and weight as larger aircraft service with the
airlines,
To ensure continued growth in payload and the profitability of cargo
operations, improvements in methods, equipment and terminal facilities will be
required in order to reduce cargo handling costs and aircraft ground time and to
provide improved service for the shippers.
We have enough hard work for this design project. A design never gets
completed in a flutter sense but it is one step further towards ideal system. But
during the design of this aircraft, we learnt a lot about aeronautics and its
implications when applied to an aircraft design.
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CHAPTER 19
BIBLIOGRAPHY
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BIBLIOGRAPHY
1. Introduction to Flight by J.D.Anderson.
2. Aircraft Performance and Design by J.D. Anderson.
3. Design of Aircraf by Thomas.C.Corke.
4. Aircraft Structures by T.H.G.Megson.
5. Aircraft Structures by D.J.Peery
6. Airframe Stuctural Design by Michael Chun-Yung Niu
7. FAA Pilot's Handbook of Aeronautical Knowledge
WEBSITE REFERENCES
1. www.wikipedia.org
2. www.nasa .gov
3. www.worldaircraftdierctory.com
4. www.airliners.net
5. www.globalsecurity.org
6. www.antonov AN-225 Mriya.com
7. www.passion for aviation.com
8. And other websites related to design of aircrafts.
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