Aircraft Design Day2

DAY 2

FUSELAGE

FUSELAGE • An aircraft’s main body section that holds crew and passengers or cargo • It is derived from the French word “Fusele – Spindle shaped”

FUSELAGE

FUSELAGE ASSEMBLAGE

TYPES OF FUSELAGE STRUCTURE • Box truss type – The structural elements resemble those of a bridge, with emphasis on using linked triangular elements. The aerodynamic shape is completed by additional elements called formers and stringers and is then covered with fabric and painted

• Monocoque – the exterior surface of the fuselage is also the primary structure

• Semi-monocoque – A series of frames in the shape of the fuselage cross sections are held in position on a rigid fixture, or jig. These frames are then joined with lightweight longitudinal elements called stringers. These are in turn covered with a skin of sheet aluminum, attached by riveting or by bonding with special adhesives

SEMI-MONOCOQUE FUSELAGE Semi-monocoque fuselage structure consists of • Longerons / stringers (Longitudinal members)  Longerons carries the bending load as axial load  Stringers also carry axial load  Stringers stabilize the skin

• Framing (Transverse members)  Provide the shape to the fuselage  Reduce the stringer length thus avoiding overall instability

• Skin  Carries the shear load from the cabin pressure, external transverse and torsional loads

• Bulkheads  Bulkheads are provided at concentrated loading regions such as wing attachments, tail attachments and landing gear locations

SEMI-MONOCOQUE FUSELAGE

COMPARISON OF FUSELAGE & WING STRUCTURE WINGS

FUSELAGE

Spar caps carry axial loads induced by bending

Longerons and stringers carry axial load induced by bending

Shear loads are resisted by spar web

Transverse shear loads are carried by skin

Rib design is influenced by local air loads

Fuselage frames are influenced by concentrated loads

Wing skin thickness is more compared to fuselage skin

Because of the curvature, the fuselage skins under compressive and shear load are more stable Less skin thickness is used

External pressure loads are more on the wing skin

External pressure loads are very less on the fuselage skin

FUSELAGE CONFIGURATION AERODYNAMIC SMOOTHNESS

PASSENGER REQUIREMENT

LANDING GEAR REQUIREMENT

TYPICAL CROSS SECTION

TYPICAL FRAMES • Former frame – Pitch = 20”

• Bulkhead frame

STRINGER FRAME CONNECTION • Stringer and frame are connected through clips • Stringer clips – Transfer the skin panel normal pressure loads to frame – Helps break up of effective column length – Provides some degree of compressive strength at the inner cap – Acts as frame web panel stiffener

FLOOR BEAMS

COCKPIT

EMPENNAGE • Empennage is the tail portion of an aircraft • The empennage gives stability to the aircraft and controls the pitch and yaw. • In simple terms the empennage may be compared to the feathers of an arrow, colloquially; "Tail Feathers" • Structurally, the empennage consists of the entire tail assembly, including the fin, tail plane and the part of the fuselage to which these are attached.

Vertical stabilizer

Rudder Elevator

Fuselage Horizontal stabilizer

LANDING GEAR

• Two landing gears are available – Nose landing gear – Main landing gear

MAIN LANDING GEAR

AIRCRAFT LOADS

LOADS • Ground loads – Landing – Transportation – Taxiing

• Air loads – Manoeuvre – Gust

BASIC FLIGHT LOADS •

Positive high angle of attack – Normal force (N) produces compressive stress on the wing upper portion – Moments from the chord wise force (C) produces compressive stress on the leading edge – Critical for compressive stresses in upper forward region and tensile stresses in lower aft of the wing

•

Positive low angle of attack – Chord wise force (C) is the largest force acting aft on the wing – Wing bending moment produce maximum compressive stresses in upper front spar flange and maximum tensile stresses in the lower front spar flange

•

Negative high angle of attack – Loads are smaller compared to positive high angle of attack – Wing bending moment produce maximum compressive stresses in lower forward region and maximum tensile stresses in the upper aft region

•

Negative low angle of attack – Chord wise force (C) is the largest force acting aft on the wing – Maximum compressive bending stresses in lower aft region and tensile stresses in the upper forward region of the wing

AIRCRAFT LOADS • Loads and forces applied on the aircraft structural components to establish the strength level – Air pressure – Inertial loads – Landing loads

DESIGN LOADS • LIMIT LOAD: – Maximum load anticipated in aircraft service life time – Load aircraft structure should withstand without causing any permanent deformation

• ULTIMATE LOAD = 1.5 x LIMIT LOAD • LOADS SPECIFIED BY THE LOAD GROUP IS ALWAYS THE LIMIT LOAD

SAFETY FACTOR • SAFETY FACTOR = ULTIMATE LOAD /LIMIT LOAD • SAFETY FACTOR IS PROVIDED FOR THE FOLLOWING REASONS – – – – –

UNCERTAINTIES IN LOADS INACCURACIES IN STRUCTURAL ANALYSIS VARIATIONS IN STRENGTH PROPERTIES OF MATERIALS DETORIATION DURING SERVICE LIFE VARIATION IN FABRICATION IN NOMINALLY IDENTICAL COMPONENTS

AIRSPEED • Airspeed (AS) – Speed of the aircraft relative to air

• Indicated Airspeed (IAS) – Speed of an aircraft as shown on its pitot static airspeed indicator

• True Airspeed (TAS) – Physical speed of the aircraft relative to the air surrounding the aircraft

Vt = V g − V w

• Equivalent Airspeed (EAS) – Speed at sea level that would produce the same incompressible dynamic pressure as the true airspeed at the altitude at which the vehicle is flying

Vi 2 1 2 1 1 6 4 = 1 + M + M + M + 2 4 40 4 Ve

LIMITING AIRSPEED • Limiting Airspeed – Speed of the aircraft up to which the structure may not undergo any permanent damage

• Crossing the limiting airspeed may cause – – – – – –

Critical gust Destructive flutter Aileron reversal Wing or surface divergence Stability and control problems Damaging buffets

FORCES ON AN AIRCRAFT

LIFT • Lift is a mechanical force generated by solid objects as they move through a fluid • Lift is the sum of all the fluid dynamic forces on a body perpendicular to the direction of the external flow approaching that body • Lift generated on an airfoil depends on – – – –

Angle of attack Speed of the air flow Total area Density of air

1 2 L = ρV SC L 2

AIRFOIL • Airfoil is the shape of a wing or blade (of a propeller, rotor or turbine) or sail as seen in cross-section • Airfoils are – Symmetric (Mean camber line coincides with chord line)

– Cambered

AIRFOIL TERMINOLOGY Leading edge

Trailing edge

• Leading edge is the front edge of the airfoil • Trailing edge is the back edge of the airfoil • The mean camber line is a line drawn half way between the upper and lower surfaces. • The chord line is a straight line connecting the leading and trailing edges of the airfoil, at the ends of the mean camber line. • The chord is the distance between the leading edge and trailing edge • The maximum thickness and the location of maximum thickness are expressed as a percentage of the chord • Camber is the asymmetry between the top and bottom curves of an aerofoil

PRESSURE VARIATION

  F = ∑ pn A = ∫ pn dA Γ

LIFT GENERATION • Lift is a force generated by turning a moving fluid

Force = mass x acceleration (F=ma) m(V − V ) F = Force = mass x Change in velocity (t − t ) 1

1

0

0

NOTE: Changing either the direction or speed of the flow generates a force

LIFT COEFFICIENT

THRUST • Thrust is a force created by a power source which gives an airplane forward motion. • It can either "pull" or "push" an airplane forward. • Thrust is that force which overcomes drag. • Thrust is generated by – Propellers – Jet engines

DRAG • Drag is the aerodynamic force that opposes an aircraft's motion through the air • Drag is generated by every part of the airplane • Drag is generated by the difference in velocity between the solid object and the fluid • For drag to be generated, the solid body must be in contact with the fluid. If there is no fluid, there is no drag • Drag acts in a direction that is opposite to the motion of the aircraft

CLASSIFICATION OF DRAG • Drag is generally divided into three categories: Parasitic drag Lift-induced drag Wave drag • Parasitic (Parasite) Drag is comprises of Form drag (Pressure drag) Skin-friction drag Interference drag

TYPES OF DRAG • Skin friction drag: Drag due to the skin friction between the molecules of air and the solid surface of the aircraft • Form drag: Form drag is due to aerodynamic resistance to the motion of the object through the fluid. This drag depends on the shape of the aircraft • Interference drag: Whenever two surfaces meet at a sharp angle on an airplane, the airflow has a tendency to form a vortex. Accelerating the air into this vortex causes drag on the plane • Lift-induced drag: is a drag force which occurs whenever a lifting body generates lift • Wave drag: Wave drag is associated with the formation of the shock waves • Ram drag: Ram drag is associated with slowing down the free stream air as air is brought inside the aircraft

SKIN FRICTION DRAG • Skin friction occurs because air has viscosity. • The entire skin friction drag is created within the boundary layer • Because air is viscous there is a shearing force occurring within the boundary layer

FORM DRAG • Form drag depends primarily upon the size and shape of the object • Form drag is due to the pressure difference between front and behind the object • Flow separation increases form drag Ex: When we walk through water we will feel as though we are being held back. This is due to the combination of the build up of pressure in front of us, and the decrease in pressure behind us. This can be seen visually by the rising wave in front of us and the depression in the water behind us

FORM DRAG • Form drag depends primarily upon the size and shape of the object • Form drag is due to the pressure difference between front and behind the object

INTERFERENCE DRAG • Interference drag is the effect of an aerodynamic component on another: wing-body, wing-nacelle, vertical-horizontal tail, junctions

INDUCED DRAG • Induced drag occurs because the flow near the wing tips is distorted spanwise as a result of the pressure difference from the top to the bottom of the wing • Swirling vortices are formed at the wing tips, which produce a down wash of air behind the wing which is very strong near the wing tips and decreases toward the wing root • The local angle of attack of the wing is increased by the induced flow of the down wash, giving an additional, downstream-facing, component to the aerodynamic force acting over the entire wing

LIFT INDUCED DRAG F

L

D Induced drag 2

2W Di = πρeV 2 b 2

REDUCTION OF INDUCED DRAG • Winglets offer the best reduction in induced drag • Winglets also produce additional thrust

WAVE DRAG • Wave drag is caused by the formation of shock waves around the aircraft • Shock waves radiate away a considerable amount of energy, energy that is experienced by the aircraft as drag • Although shock waves are typically associated with supersonic flow, they can form at much lower speeds at areas on the aircraft where, according Bernoulli’s principle, local airflow accelerates to supersonic speeds over curved areas • The effect is typically seen at speeds of about Mach 0.8, but it is possible to notice the problem at any speed over that of the critical Mach of that aircraft's wing • The magnitude of the rise in drag is impressive, typically peaking at about four times the normal subsonic drag

RAM DRAG • Ram drag is associated with slowing down the free stream air as air is brought inside the aircraft V0 FD r = w0 g w0 – Intake air mass flow g – Acceleration due to gravity v0 – Flight velocity

TOTAL DRAG Total drag

1 2 DT = ρV SC d 2

C d = C di + C dp

DRAG COEFFICIENT Bodies

Drag Coefficient

Airfoil Section, minimum [1]

0.006

Airfoil Section, at stall [1]

0.025

2-Element Airfoil

0.025

4-Element Airfoil

0.05

Subsonic Aircraft Wing, minimum [2]

0.05

Subsonic Aircraft Wing, at stall [2]

0.16

Subsonic Aircraft Wing, minimum [3]

0.005

Subsonic Aircraft Wing, at stall [3]

0.09

Subsonic Transport Aircraft

0.012

Supersonic Fighter, M=2.5

0.016

Airship

0.020-0.025

WEIGHT ESTIMATION

AIRCRAFT WEIGHT DISTRIBUTION • Weight breakup of the vehicle is – Structural weight (40%) – Payload (20%) – Fuel weight (40%)

WEIGHT

WEIGHT • Design takeoff gross weight is the total weight of the aircraft as it begins the mission for which it was designed • Aircraft weight consists of – – – – –

Pay load Fuel Power plant Fixed equipment Structure

Wo = W payload + W fuel + W power plant + W structure + W fixed equipment

PAYLOAD • Payload consists of the weights of passenger, crew and luggage Wpayload = Wpassenger+Wcrew+Wbags+Wcargo

EMPTY WEIGHT • Empty weight of the aircraft

Wempty = W structure + W power plant + W fixed equipment • Empty weight fraction for – Jet aircraft 0.45 to 0.55 – Turbo prop 0.55 to 0.61

STRUCTURAL WEIGHT • Structural weight of the aircraft WStructure = Wwing+Wfuselage+Wtail+Wlanding gear

WING WEIGHT Wing weight = 8% of airplane weight W wing= 0.0051(Wdg N Z )

0.557

S

0.649 w

A

0.5

( c) t

−0.4

0.1 −1.0 0.1 ( 1 + λ ) ( CosΛ ) S csw root

Ref: P.Raymer

Wing weight fraction ratio Wwg Wzf

b  0.55 −0.3  = 0.0017 nult t r ,max   cos Λ c / 2 

W wing= 4.22 S wg + 1.642 * 10

−6

1.05

 1+  

6.25cos Λc / 2 W zf   b  S

N ult b 3 TOW * ZFW (1 + 2λ ) t Cos 2 ΛS wg (1 + λ ) c avg

( )

  

−0.3

FUSELAGE WEIGHT • For L/d>5; the gross area of fuselage cabin is  2  S g = π dL 1 −   ( L/d)

2/3

 1  1+ 2  ( L / d ) 

• Weight of fuselage is

W f = 0.021S

1.2 g

lt VD, E d

• Where W – Weight (lbs) V – Design dive speed (knots) f

D,E

LANDING GEAR WEIGHT • For main landing gear and nose landing gear

Wmg = 40 + 0.16W

3/ 4 to

Wng = 20 + 0.10W

3/ 4 to

−5

+ 0.019Wto + 1.5 ×10 W −6

+ 2 ×10 W

3/ 2 to

• For A/c with take off weight > 10000 lbs, landing gear weight is 3.5% to 4.5%

3/ 2 to

TAIL PLANE WEIGHT • For main landing gear and nose landing gear

Wmg = 40 + 0.16W

3/ 4 to

Wng = 20 + 0.10W

3/ 4 to

−5

+ 0.019Wto + 1.5 ×10 W −6

+ 2 ×10 W

3/ 2 to

• For A/c with take off weight > 10000 lbs, landing gear weight is 3.5% to 4.5%

3/ 2 to

FUEL WEIGHT • Fuel weight fraction (Wf/W0) can be estimated based on the mission profile of the aircraft • Mission profile consists of – – – – – –

Takeoff Climb Cruise Descend Loiter Landing

 Wx = 1.061 − W0  W0

Wf

  

TYPICAL MISSION PROFILE 3 2 4

5 1 0

Takeoff (0-1) Climb (1-2) Cruise (2-3) Descend (3-4, 5-6) Loiter (4-5) Landing (6-7)

6

7

MISSION SEGMENT WEIGHT • Mission segment weight fraction (Wi/Wi-1 ) is the weight at the end of the segment divided by weight at the beginning • Mission segment weight fractions – Takeoff 0.970 – Climb 0.985 – Landing 0.995

CRUISE MISSION WEIGHT • Cruise segment weight fraction is calculated using Breguet range equation V L  W i −1   R= ln C D  Wi 

W i −1  − RC   = exp Wi  V ( L / D) 

LOITER FUEL WEIGHT • Cruise fuel weight fraction is calculated from endurance equation

1 L  W i −1   E= ln C D  Wi  W i −1  − EC   = exp Wi  ( L / D)  C – specific fuel consumption

SPECIFIC FUEL CONSUMPTION • Specific fuel consumption is the rate of fuel consumption divided by the resulting thrust • SFC is measured as fuel mass flow per hour per unit thrust force • Unit for SFC is – lb of fuel/hr/lb of thrust (FPS system) – mg/Ns (Metric system)

• SFC for propeller engines are measured as Cbhp pounds of fuel per hour to produce one hp at the propeller shaft C = C bhp

V 550η p

SPECIFIC FUEL CONSUMPTION Typical Jet SFC lb of fuel/hr/lb of thrust (mg/Ns) Type

Cruise

Loiter

Pure Turbojet

0.9 (25.5)

0.8 (22.7)

Low-bypass turbofan

0.8 (22.7)

0.7 (19.8)

High-bypass turbofan

0.5 (14.1)

0.4 (11.3)

Propeller SFC lb/hr/bhp (mg/W.s) Piston prop (fixed pitch)

0.4 (0.068)

0.5 (0.085)

Piston prop (variable pitch)

0.4 (0.068)

0.5 (0.085)

Turboprop

0.5 (0.085)

0.6 (0.101)

EXAMPLE • Takeoff weight W1/W0 =0.97 • Climb • Cruise

W2/W1 =0.985

W3  − RC   = exp W2  V ( L / D) 

W3  − 2800000 * 0.000139  = exp  = 0.855 W2 204 * 13.9  

• Loiter

W4  − EC  = exp  ( ) W3 L / D  

W4  − 10800 * 0.0001111  = exp  = 0.917 W3 16  

• Land

W5/W4=0.995

• W5/W0=0.995*0.917*0.855*0.985*0.97=0.745 • Wf/W0=1.06*(1-0.745)=0.270

AIRCRAFT COORDINATE SYSTEM

STABILITY AND CONTROL

AIRCRAFT MOTIONS (Y)

(X) (Z)

AIRCRAFT MOTIONS

ROLLING

ROLLING

ROLLING

PITCHING

PITCHING

YAWING

YAWING

SPOILERS

FLAPS AND SLATS