Lift Notes Technical

Lift: – is the part of total reaction which is perpendicular to the relative airflow/flight.

Drag: – is the part of total reaction acting in the same direction of the airflow.

Chord: – is a Straight line joining the leading edge to the trailing edge.

Mean Camber line: – The line which is equidistant from the top and the bottom surface of the aerofoil.

Camber: – The difference between the Chord and the Mean camber line.

Thickness: – It is the distance between the top and the bottom surface the aerofoil.

AOA: – The acute angle between the relative airflow and the chord.

Thickness Ratio: Max. Thickness/Chord.

Finess ratio: – Chord/max. Thickness or Length/Breadth.

Two Aircraft’s having the same finess ratio will have the same coefficient of drag. But, the larger body will have more drag.

The larger body will have more value of S, in the Drag formulae Drag = ½ RHO V2 S. where S is surface area causing Drag.

Angle of incidence: – Angle between the chord and the longitudinal axis.

Aspect ration: –

Span/Chord (for a rectangular wing)

Span/Mean Chord [for a Tapered wing]

(Span)2 /Wing Area [for a Delta Wing Aircraft]

Wing Area/(Chord)2

Wash out: – is an decrease in angle of incidence towards wing tip.

Wash in: – is an increase in angle of incidence towards wing tip.

Boundary Layer: – It is the layer of air on the top surface of aerofoil which slows retardation in velocity (from free stream to zero). The flow of boundary layer is laminar. But, due to friction with in the boundary layer the flow becomes turbulent.

Transition Point: – It is the point on the top surface of aerofoil where the relative airflow changes from laminar to turbulent.

Separation Point: – It is the point on the top surface of aerofoil where boundary layer gets separated from aerofoil.

With increase in the AOA the transition point and the separation point move forward.

Center of Pressure: – It is an imaginary point on the top surface of an aerofoil through which all the upward forces through an aerofoil can be assumed to act.

Stagnation Point: – It is the point on the leading edge of the wing where the airflow comes to a momentary halt. At this point the pressure is equal to dynamic pressure + static pressure.

Relative Airflow: – The airflow acting in the opposite direction of the Aircraft moment. It depends upon the speed of the Aircraft.

Symmetrical Aerofoil: – Aerofoil which has no camber or an aerofoil in which chord line and the mean camber line are same.

In an asymmetrical aerofoil C.P. moves forward with the increase in the Angle of Attack. At the stalling angle it is at the maximum Forward position, while still producing lift. Anymore increase in the Angle of Attack will cause the Centre of Pressure, Transition point separation point to abruptly move backwards and the lift produced suddenly decreases.

It an asymmetrical aerofoil the C.P. generally moves between 20% and 30% of the chord line aft of the leading edge in normal working range of AOA.

Symmetrical Aerofoil produces zero lift at 00 AOA and an Asymmetrical aerofoil produces zero lift at negative angle of attack (Generally At –40).

Max. lift is produced at stalling angle (just before stalling)

Normally a stalling angle is 160‑.

Lift = ½ ℓ V2 S CL

Where

CL = Coefficient of Lift, which depends upon

Shape of the Aerofoil

Angle of Attack

ℓ = Air Density

V = Velocity or Air speed

S = Surface Area

Bernoulli’s Principle: – In a streamline non viscous flow the sum total of pressure, Kinetic Energy and Potential Energy remains constant.

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Characteristic of high lift Aerofoil: –

High Thickness ratio. The maximum Thickness is about 25% to 30% of the chord line behind the Leading edge.

Large Camber

Well rounded Leading Edge.

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Characteristics of High Speed Aerofoil: –

Very low Thickness Ratio, maximum Thickness is about 50% of the chord line behind the leading edge.

Very little or No Camber.

Sharp Leading Edge.

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Pressure Distribution around an Aerofoil:-

At o° Angle of Attack or at Negative Angle of Attack-

At the top surface of the Aerofoil there is a decrease in the pressure (less than the surrounding air pressure). At the bottom surface of the Aerofoil also the pressure is less than the surrounding air and the NET LIFT is O

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At small Positive Angle of Attack-

At the top surface of the aerofoil the pressure decreases (less than the surrounding air pressure) at the bottom surface of the aerofoil also the pressure decrease (i.e. It decreases but it is still less than the free stream atmospheric  pressure or surrounding pressure). But this decrease in pressure is less than the decrease in pressure at the top surface. Therefore there is a net pressure differential which produces lift.

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At large Angle of Attack-

The pressure at the top surface of the aerofoil decreases. (Less than the surrounding atmospheric pressure) but at the bottom surface of the aerofoil the pressure INCREASES (i.e. It increases more than the atmospheric Pressure).

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Aerodynamic Centre: It is the fixed point on the wing profile about which the lift movement co-efficient remains constant and does not change with the changes in the Angle of Attack.

At Subsonic speeds it is located at 25% of the chord line behind the leading edge and at the Supersonic speeds it is at 50% of the Chord line behind the leading edge.

Aerodynamic Centre depends upon the thickness of the wing.

It changes only with the changes in Aerofoil/ Wing thickness.

If Wing Thickness increases the Aerodynamic Centre moves forward (Moves less than 25% aft of the Leading Edge).

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Drag: is the Resistance offered by the airflow to the forward motion of the Aircraft. Or we can say that it is the part of the total reaction acting in the same direction as the airflow

Drag
Parasite drag or Zero lift drag Wing drag or lift dependent drag
Profile drag Interference Drag Induced drag Increment of 0 Lift drag
Form drag Skin friction

Parasite Drag: – is because of Non Lift producing surfaces of the Aircraft.

Profile Drag: – is because of the shape of different parts of the Aircraft. It can be reduced by streamlining the body parts.

1)     Interference Drag: – is caused because of the Eddies. Where ever two different surfaces of the Aircraft are joined together they create a Drag called Eddies. This Drag results in interference Drag. It can be reduced by providing Metal Fairings.

2)    Form drag: – is basically because of the shape of different Aircraft parts.

3)    Skin friction: – is because of the friction being offered by the different parts of the Aircraft.

Parasite drag depends upon the air speed. It is Directly Proportional to the Airspeed or we can say It increases with the Airspeed.

4)    Induced drag: – It is because of lift producing surfaces of Aircraft.

Pressure distribution around a wing: – There is Low pressure on the top surface of the wing and High Pressure at the bottom of the wing. At the wing tips the High Pressure air from below the wing comes up and meets the air on the top of the wing  giving rise to Wing Tip Vortices.

  1. a) At the trailing edge of the wing the airflow at the top has a Down wash and the airflow at the bottom has an Up wash. Where these two air masses flowing in different directions meet each other at the trailing edge they are formed in a shape of whirling air called TRAILING EDGE VORTICES.

After forming Trailing Edge Vortices Move out wards and mix with the Wing Tip Vortices. This gives rise to a drag called INDUCED DRAG.

Airflow at the top of the wing is inclined towards the longitudinal axis and at the bottom of the wing the airflow is inclined outwards from the longitudinal axis.

Increment of Zero Lift Drag: – At Zero degrees Angle of Attack the lift produced is O but a small amount of drag is still produced, this drag is known as increment of Zero Lift Drag.

Note: Aspect Ratio increases, lift increases, Induced Drag reduces, Wing Tip Vortices reduces and Stalling Angle reduces.

Factor effecting Induced Drag:-

  1. a) CL, Coefficient of Lift: – CID * CL2
  2. a) AOA : Induced Drag increases with the increase in Angle of Attack
  3. i) Because with increase in Angle of Attack, Coefficient of Lift increases
  4. ii) With increase in Angle of Attack there is greater low pressure at the top of the wing and greater high pressure at the bottom of the wing, therefore Wing Tip Vortices are produced more, thereby increasing Induced Drag.
  5. b) Air speed: – With the increase in the airspeed induced drag reduces because at higher air speed we fly at lower Angle of Attack to produce the same lift.

If the speed is high, by the time Wing Tip Vortices are formed the Aircraft would have already moved forward.

  1. c) Aspect ratio : –Induced drag is inversely Proportional to Aspect Ratio As the Wing Span increases it cause the Aspect Ratio to increase, which decreases the Induced Drag.
  2. d) Weight of the AIRCRAFT: – With the increase in Weight the Induced Drag increases. Because an increase in the weight requires the lift requirement to be increased and the Aircraft is required to fly at a higher Angle of Attack, to produce more Lift.
  3. e) Finess Ratio: – With the increase in Finess Ratio Induced drag decreases.
  4. f) Total Drag = Induced Drag + Parasite Drag

VMD: – Velocity of minimum Drag

Ø      With the decrease in the airspeed total Drag reduces till a particular airspeed (VMD) after that it increases.

Ø      At VMD Induced Drag = Parasite Drag or we can say at VMD Induced Drag = Parasite Drag

Ø      In the area of Reverse Command to maintain a lower airspeed you need a greater power, but vice versa is not true. Example – The Airspeed can be reduced by increasing the power settings in the Area of Reverse Command. But for a greater power you do not require a lower airspeed. *******************************************

Ø      O° to ½0 Angle of Attack Drag is minimum, Drag is maximum only after stalling Angle. Lift is 0° at negative Angle of Attack and maximum at Stalling Angle.

Ø      Ratio of Lift and Drag is L/D Ratio.

Ø      L/D ratio depends upon Angle of Attack, with increase in AOA, L/D ratio increases. L/D Ratio is usually maximum at 30to 40 AOA.

L/D ratio indicates the Aerodynamic Efficiency of the wing.

In normal transport category Aircrafts L/D ratio is 15:1 or 20:1

Requirements of ideal Aerofoil: –

1)     High Coefficient of Lift Max: – CL max controls the landing speed higher the CL lower is the landing speed.

2)     Low Coefficient of Drag Min: – If CD is low wing will experience low resistance due to Drag, enabling it to reach higher airspeeds.

3)     High L/D Ratio: – High L/D ratio means good efficiency good weight carrying capacity, less fuel consumed for distance covered giving high Endurance.

4)     Small Movement of Centre of Pressure: – Small movement of C.P. means the greatest pressure on the wing remaining in one fix portion, which can be made more tensile and stronger than the remaining Aerofoil parts and thus reduces the weight of the Aircraft Structure.

5)     Stable Movement of Centre of Pressure: – Stable movement, means that with the increase in AOA nose down movement about the leading edge should increase.

6)     High maximum value of CL3/2/CD, the higher the value of this ratio lesser is the power required.

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When note CL³ ÷ 2CD governs the Power required

As we know that Power = Force × Distance or Drag × TAS

In a straight and level flight Lift = Weight

Lift × Drag = Weight × Drag

Drag = Drag ÷ Lift × Weight

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When LIFT = Lift = ½ ℓ V2 S CL

Where

CL = Coefficient of Lift, which depends upon

  1. Shape of the Aerofoil
  2. Angle of Attack

ℓ = Air Density

V = Velocity or Air speed

S = Surface Area

Therefore V = √ Weight ÷ CL ½ ℓ S                     Lift = Weight

Therefore Power = CD ÷ CL × Weight × √ Weight ÷ CL ½ ℓ S = CL³/2/CD

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 Stalling or Burbling: – AIRCRAFT will stall whenever the Stalling Angle is exceeded but the stalling speed will be different in different conditions.

For every Aircraft stalling angle is always fixed, unless changed by some Lift Augmenting Devices.

Ø      Stall takes place whenever the boundary layer breaks away from the surface of the aerofoil, the lift suddenly reduces, drag increases and Aircraft cannot sustain flight in the air.

Ø      After stalling lift doesn’t become zero, because-

  1. a) There is still high pressure in the Bottom of the wing (30% of the Lift comes from Bottom surface of the wing)
  2. b) Certain amount of Lift is still produced from the Leading Edge of the wing.

Ø      After the stalling Angle is exceeded the Centre of Pressure, Transition Point and Separation point all abruptly move backward.

Ø      Stalling Angle is directly proportional to the Square of the Lift

Factors Effecting Stalling Speed: –

With increase in the weight, Stalling speed increases.

As a rule 10% increase in the Weight increases Stalling Speed by 5%.

Therefore Stalling Speed is directly proportional to the Aircraft All Up Weight.

V2 = V1 √Weight2 ÷ Weight1

V2 = New Stalling Speed

V1 = Old Stalling Speed

W2 = New Weight

W2 = Old Weight

With increase in weight, Stalling Angle is reached at higher air speed.

Power: – Power on stalling lower than the power off stalling speed, because-

  1. a) In a propeller driven Aircraft the slipstream from the propeller (Prop wash) pushes the transition point backward, it reactivates the boundary.
  2. b) The Thrust is divided into two components. One part acts in the direction of Lift and thereby reducing the stalling speed because Lift requirement is reduced.

Load factor: – Lift ÷ Weight or Live Load ÷ Dead Load

Live load: – The Weight of the Aircraft in the air

Dead Load:-The Weight of the Aircraft on the Ground.

Load Factor is the ratio between the total Air Load being imposed on the wing to the Weight of the Aircraft.

With the increase in the load factor stalling speed increases.

**V2 = V1 √Load Factor

With increase in Load Factor the Stalling Angle is reached at a higher air speed or in turns the stalling speed becomes higher.

 (a)      Lift Augmenting Devices: Lift Augmenting Devices increase the value of CL max, therefore Lift requirement is reduced and thereby Stalling Speed reduces.

(b)     Position of Centre of Gravity: – With the Forward C.G. position the stalling speed increases, because the effective weight of the Aircraft increases and with the Rearward C.G. the Stalling Speed reduces.

(c)      Air Density : – has No effect

(d)     Altitude.: –  has No effect

(e)      Wind Velocity: – has No effect

(f)       Turbulence: – In turbulent air Aircraft will stall at a higher Air speed, because of sudden increase in Angle of Attack caused by up gust. Aircraft can stall at any air speed and any attitude but only whenever the Stalling Angle is exceeded.

All the above factors change only the stalling speed except the Lift Augmenting Devices which will alter the Stalling Speed and Stalling Angle as well.

Critical Angle of Attack is same as Stalling Angle of Attack.

Stalling speed is directly proportional to under root of Lift.

Stalling Speed is inversely proportional to under root of All up Weight.

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LEVEL FLT

Forces acting on the Aircraft

1)     Lift

2)     Weight

3)     Drag

4)     Thrust

In Straight and Level Flight Lift = Weight and Thrust = Drag

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All the four forces cancel out each other and Aircraft is in the state of Dynamic Equilibrium and Aircraft continues to move because of inertia.

  • Two Aircrafts are in Straight and level Flight one is flying faster and the other is flying slower both Aircraft have same weight. Will the Lift produced be different for each Aircraft? Both Aircrafts will be producing same lift, but the slower Aircraft is flying at higher AOA.
  • Lift passes through C.P.
  • Weight Passes through C.G.
  • By design in all commercial Aircrafts C.G. is always kept ahead of C.P.
  • Power has no effect on lift

Lift

Lift and Weight form a couple which gives the nose down pitching movement.

Thrust line passes below the drag line, Thrust and drag form a couple which gives nose up pitching movement.

Ø      Lift Weight couple is much more greater than Thrust Drag couple (Because L/D ratio is 15:1 or 20:1), Therefore there is a resultant nose down pitching movement and the balancing force is provided by the tail plane.

Ø      On adding power Aircraft has tendency to pitch up, because there is sudden increase in the Thrust Drag couple.

Ø      The advantage of having Lift Weight couple more then Thrust Drag couple is that in case of engine failure the nose will pitch down and pilot will not have to struggle to maintain airspeed.

FACTORS EFFECTING LEVEL FLIGHT: –

1)      Air speed: – Doesn’t affect level flight directly but in a level flight there is inverse relationship between AOA and air speed. Although all Aircraft’s do not have AOA indicator but knowing the air speed pilot can have a fair idea of the AOA.

2)     Weight: – With increase in wt. lift requirement increases. To increase Lift AOA is increased and therefore as weight increases cruising speed decreases.

Increasing Weight effects-

  1. Range will be less
  2. ANPG (Air Nautical Miles per Gallon) will reduce.

iii.             Fuel consumption will increase

  • A greater payload can be carried to a lesser distance

3)     Temperature: – With increase in temperature power output of the engine reduces and power required by the airframe. In warmer air TAS increases and to maintain the increased TAS more Power is required.

  1. Cursing Speed reduces

During summers Aircraft flies slower.

  1. Range will reduce with increase in temperature specific gravity of the fuel reduces, Therefore Weight reduces and for the same volume of the fuel the weight of fuel carried reduces.

iii.             ANG reduce (Air Nautical Miles per Gallons reduces)

  1. Fuel consumption increases

4)     Attitude: – With increase in altitude the power out reduces and power requirement of the airframe increases.

5)     Centre of Gravity: – With the forward C.G. effective weight of the Aircraft increases, Therefore it causes the following-

  1. Cursing speed reduces
  2. Range reduces

iii.             ANPG reduces (Air Nautical Mile per Gallons reduces)

  1. Fuel consumption increases
  • On a long Cross Country Flight as the fuel is being consumed, the weight of the Aircraft reduces, causing Cruising Speed to increase, Air Nautical Mile per Gallons increase and Fuel consumption to reduce because of reducing weight.
  • To get maximum range Aircraft should be flown at VMD.
  • But this speed is too low and controllability problem arises, therefore Aircraft is flown at 1.10% of VMD or 1.1 VMD and this speed is called RECOMMENDED RANGE SPEED and this is equal to the speed for the best L/D ratio.

Gliding: – is descending without power

Forces acting on the Aircraft during glide: –

  1. Lift
  2. Drag
  3. Weight.

During Gliding Lift and Drag are balanced by

Lift (L) = W Cos q

Drag (D) = W Sin q

q = Gliding Angle

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Gliding Angle: – is the Angle which the flight path of the Aircraft makes with the horizontal plane

Apparent Gliding Angle: – is the Angle which the flight path of the Aircraft makes with a fixed point on the ground.

The Apparent Gliding Angle depends upon the Wind Velocity.

Head Wind increases Apparent Gliding Angle.

Tail Wind decreases Apparent Gliding Angle.

  • Gliding angle is not effected by the Wing Velocity.

Gliding Endurance: is the time taken by the Aircraft to reach the ground during Gliding

Gliding range: ­- is the dist covered by the Aircraft during Gliding.

Factors Effecting Gliding: –

  1. Lift/Drag Ratio: – Higher the L/D ratio, lower (Shallow) is the Gliding Angle.
  2. Air Speed: – For a particular weight of the Aircraft there is only one air speed which will give best L/D ratio and best Gliding Angle.
  • If we fly at a speed higher than the best L/D ratio speed the L/D ratio will decrease. Parasite drag increases, Gliding Angle will be steep.
  • If we fly at a speed lower than the best L/D ratio speed the L/D ratio will ratio will again decrease. Induced drag increases, Gliding Angle will again be steep.
  1. Weight: – With increase in weight, the speed for the best L/D ratio increases and therefore the speed for the best Gliding Angle.
  • A 10% increase in weight, increases the best Gliding Speed by 5%.
  1. a) Gliding Angle: – is not effected by weight, provided the best L/D ratio speed is maintained.
  2. b) Gliding range: – is Not effected
  3. c) Gliding Endurance: – With the increase in weight, Gliding Endurance decreases and Rate of Descent increases.
  4. Temperature: – has No effect.
  5. Altitude: –
  6. Gliding Angle: – has No effect
  7. Gliding Range: – has No effect
  8. Gliding Endurance: – With increase in altitude Gliding Endurance increases, because at higher altitude TAS increases.
  • For the same weight, the IAS for approach and landing remains the same.

A/c with same weight approaching to land at Mumbai or Simla will follow the same IAS, but TAS at Simla is more, due to high altitude airfield, causing Gliding Endurance to increase.

  1. Wind Velocity: –
  2. a) Gliding Angle- has No effect
  3. b) Apparently Gliding Angle-

Head wind increases Apparent Gliding Angle

Tail wind decreases Apparent Gliding Angle.

  1. c) Gliding Endurance- has No effect.
  2. d) Gliding Range-

Head Wind reduces Gliding Range.

Tail Wind increases Gliding Range.

  1. Flaps: – reduce Lift/Drag ratio, the Gliding Angle increases. (without increase in airspeed)
  • Power on Descent: – In a power on descent, Thrust is less than the Drag (That is why Aircraft is Gliding). On descending with Power Gliding Angle reduces and Gliding Range increases.

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FLIGHT CONTROLS

All the axis of the Aircraft passes through the center of gravity (C.G.) and is at 90° to each other.

Flying controls: –

  1. Ailerons
  2. Elevators
  3. Rudders
  4. Differential spoilers

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(1)     Aileron– controls the movement of the Aircraft about Longitudinal Axis, or along the Lateral Axis, they control the Rolling movement.

(2)    Elevators– controls the movement of the Aircraft about Lateral Axis, or along the Longitudinal Axis, they control the Pitching movement.

(3)    Rudders– controls the movement of the Aircraft around/about Vertical Axis, they control the Yawing movement

PRINCIPLE OF CONTROL SURFACE: – is that the increase in Camber increase the Lift.

Aileron Drag: – is the drag experienced by the down going aileron as it moves into the area of high pressure.

The up coming aileron, which moves into the area of low pressure experiences less drag.

Adverse Yaw: – is the yaw in the opposite direction to the roll and it occurs because of the Aileron Drag. It is more prominent at low air speed, since the aileron drag is maximum at the low air (at low air speed control effectiveness is less and to bring in the same Roll movement a greater movement of ailerons is required).

Methods, to overcome Adverse Yaw: –

  1. Rudders: – Primary function of the rudders is to overcome adverse yaw.
  2. Differential Aileron: – The aileron moves in such a way that the down going aileron moves through a lesser angle and the up going aileron moves through a greater angle. This reduces the drag on the downing aileron, which is exposed to area of High Pressure.
  3. Frise Ailerons: – The upcoming aileron moves in such a way that the part of it moves in to the under surface of the wing, but the down going aileron moves in a normal way and creates a smooth surface both above and below the wing. Therefore the drag on the upcoming aileron is increased.
  4. Combination of both Frise and Differential Ailerons are also used.
  5. Control Coupling: – In this Rudders and Ailerons are coupled together, whenever a small Roll is required the same side Rudder is automatically applied to assist in the Roll moment

Aileron Reversal: – is a High Speed Phenomenon that takes place in Sweep Back Wing Aircraft. In this the Aircraft Rolls in the opposite direction to the intended Roll.

E.g. – AIRCRAFT rolling to the left, left aileron comes up right aileron goes down. As the lift aileron comes up Center of Lift moves forward, wing twists downward. Therefore increasing the angle of incidence of the left wing, AOA increases lift produced by the left wing increases and Aircraft rolls to turn right instead of left.

It happens in the sweep back wing because the Tip Chord is weaker than the Root Chord, wing is weaker towards the tip where the ailerons are located, and therefore the wing twists easily.

Ways To Overcome Aileron Reversal

  1. Spoilers: – It is a device which reduces lift (primary function of Spoilers is to dump the Lift). At high airspeeds the aileron are locked and the roll control is provided by the differential spoilers. e.g. Aircraft wants to roll to the left, left side spoiler will come up, right side spoiler will remain retracted, this will reduce the lift on the left wing and Aircraft will roll to the left.

Spoilers can also be used as Air Brakes/Speed Brakes/Lift Dumpers. On using spoilers as speed brakes both side spoilers will come up and they are not available for roll control.

  1. Some Aircraft have two set of aileron :

(a)   Inboard Ailerons: are installed towards the fuselage.

(b)  Outboard Ailerons: are installed towards the wing tips.

At high air speeds only Inboard Ailerons operate, because at high speeds the Air loads are very high and the Inboard Ailerons are of a higher strength to withhold these loads, since they are located closer to the fuselage where the wing is connected to the Fuselage and twisting will not occur.

  1. At slow airspeeds both INBOARD and OUTBOARD Ailerons move together.

FLAPRONS- is when the Inboard Ailerons extend lower as Flaps, where they again work as the Aileron. FLAPRONS prevents the airflow being affected to the aileron as the Flaps go down

The maximum movement of the flying controls is limited by.

ADJUSTABLE STOPS: – They are provided at the Control Surface as well as in the Cockpit. (Control Column)

Turn Buckles: – Are provided to adjust the tension of the Control Cables. The Turn Buckles must be wire locked after they have been adjusted.

Any work which is done on flying controls and engine requires Duplicate Inspection.

Temperature Compensator: – Allows for the changes in the temperature to control the cable tension.

Elevons: – Elevators + Ailerons, present in Delta Wing Aircrafts.

Ruddervators or Elerudder: – Rudders + Elevators present in V-tale Aircraft or Butterfly Tail Aircraft.

Factors Effecting Control Effectiveness:

(1)    Size and Shape of the control

(2)    Deflection Angle

(3)    Airspeed ((EAS)2 )

(4)    Movement Arm (Distance From Centre of Gravity)

Effect of Speed on Controls: – The Deflection Angle requires give an attitude change is proportional to (EAS)2. To the Pilot it means that when the speed is reduced by half, the control deflection must be increased 4 times to achieve the same change of movement.

Effect of forward speed on Ailerons: – The Rolling velocity due to aileron power is directly proportional to Aircraft speed. For a given deflection of aileron the steady rate of roll will also depend on Dx, (where Dx is the change in the AOA of the wing produced by the rolling velocity in the forward flight. It tends to dampen the rolling effect.) dampening in roll effect. As forward speed increases this Dx, dampening effect increases. To counteract this error the rolling movement increases as Aircraft speed increases.

(1)   The Rate of Roll increases in the same ratio as the forward speed increases for a given Aileron defection.

(2)   Effect of Altitude : – With the increase in altitude forward speed increases, since we know as altitude increases TAS increases. Therefore the steady rate of roll at a given Equivalent Air Speed (CAS corrected for Compressibility error) increases. The Dampening in yaw effect also decreases with increase in altitude.

Loose control cables will cause the controls to float during flight i.e. the controls are kept slightly higher than the neutral position, because low pressure is on the top of the wing. This floating is more prominent in ailerons as compared to rudders and elevators. To prevent this up float in the flight an Aileron Droop is given to the ailerons.

BACKLASH- occurs due to Loose control cables. It occurs when the  controls move later than the control column.

If the control cables are too tight it will cause the controls to SNATCH.

(3)   SNATCH: – Snatching occurs due to too tight Control Cables. It occurs at or near stall speed or occurs at very high speeds like Critical Mach No. it is caused by continuous and rapid shifting of C.P. of the ailerons due to the disruption of the airflow over the surface resulting in Snatching or Jerking of the Control Column.

Rudder Reversal: – It takes place at high Mach No./High Speeds (At speeds above Critical Mach No.) due to the formation of Shock Waves and Aircraft yaws in the opposite direction to the intended Yaw.

E.g. Aircraft wants to yaw to the left, left rudder pedal is pressed, rudder goes to the left, velocity of the air on the right side increases, it crosses the Critical Mach No. Shock Waves are formed and CL on the right side reduces, tail goes to the left and AIRCRAFT YAWS TO THE RIGHT.

Effectiveness of the elevators in the addition to the above factors also depends upon the position of center of gravity.

With the forward located C.G. the greater movement of the elevator is required to bring in the same changes in the attitude.

Elevons: – They are present in Delta Wing Aircrafts where the functions of Elevators and Ailerons are combined.

If the control column is pulled back both the Elevons will come up causing both ailerons to act as an Elevator and go up, giving pitch up moment, nose pitches up Now if the control column is moved to the one side the angular position of both elevators changes (one elevon moving up through a lesser angle then the other). This causes the rolling movement in the required direction.

Ruddervator: – Are provided in a “V” tail or butterfly tail Aircrafts. In this the function of Rudder and Elevator are combined. The lift force being produced is 90° to each of the Ruddervator surface and therefore it is divided into two components. 1. Vertical components are used to provide pitching movement. 2. Horizontal component is used to provide the yawing movement.

  • All moving (slab or flying tail): – In this full or accurate control retained at all Mach No. of high speeds. Forward movement of control  columns increases the Angle of Incidence of the tail plane and produces an upward force necessary to lower the nose. In some Aircrafts elevator is retained and it is linked to tail plane in such a way that the movement of the tail plane causes elevator to move in the normal direction to help the tail plane.
  • STAB TAIL PLANE- is when there is no separate Elevator in the system and the whole tail plane assembly moves on moving controls columns. It is used at high airspeeds because conventional tail plane looses its effectiveness at high airspeeds or at high Mach No. due to the formation of Shock Waves.

Tail or Variable Incidence Tail, it is used in some Aircrafts in addition to the trimming tabs. The V tail plane is more effective than the tabs at high Mach No. By varying the incidence of the tail plane the “OUT OF TRIM” forces can be balanced.

CONTROL BALANCING

Stick Loads: – They are the aerodynamic loads exerted on the controls or the load exerted by the controls. They are felt because of the deflection of the control surface into the airflow. They increase with the increase in the airspeed.

Hinge:-  is a point around which  the control surfaces move up and down

Moment: – It is hinge moment = L  X D

Hinge moment is the force required by the pilot to move the controls.

This force depends upon

(1)    D = Distance between Hinge point and the point from where the Lift is acting.

(2)    L = Lift being produced by the Control surfaces.

(3)    The Surface Area of the Control Surfaces.

Ways of control Balancing: –

  • Control balancing is the means of reducing the Hinge moment and therefore reducing the physical force reputed by the pilot to move the controls.
  1. Mass Balance: – A mass is attached ahead of the Hinge. This will shift the Center of Gravity closer to the hinge, but C.G. is still kept behind the hinge.
  • Advantages of Mass Balancing
  1. It reduces the force required by the Pilot to move the Controls.
  2. It reduces the Control Flutter.
  • Control Flutter: – Is the phenomenon of high speed where the Control Surfaces moves up and down about its mean position.
  1. Aerodynamic Balancing: – In this we create Aerodynamic forces in such a manner that they help the pilot to move the control.
  2. Inset or set back hinge: – In this Hinge point is moved backwards i.e. closer to the C.G. but it is still kept ahead of the C.G. Thus reducing the Hinge moment. The amount of inset is limited to 20% to 25% of the Chord Length. (This ensures that the C.P. of the control does not move ahead of Hinge at large angles of deflection.
  3. Internal Balance: – The area between Tail plane and Elevator is sealed and the tail plane and elevator are joined by a diaphragm. The high pressure on the Leading edge of the Elevator moves the elevator up by moving the diaphragm up and therefore the trailing edge goes down.
  4. Horn balancing: –
  5. Horn : – It is the part of control surface which projects ahead of the Hinge line.

E.g.- If the elevator goes down the horn comes up airflow from the top tends to keep the Horn up and therefore keeps the Elevator down.

Horns are of two types 1. Shielded and  2. Unshielded

Disadvantages of Horn Balancing:-

  1. At high speeds these type of Balancing can lead to the Control Flutter.
  2. The entire Balance portion is at one end of the Hinge and the forces on it can exert the twist on the Hinge.

System of Tabs: –

Tab– is an Auxiliary Aerofoil attached at the Trailing edge of the Control Surface.

  1. Fixed Tab- It can only be adjusted on ground. They cannot be adjusted in the air. Tabs move in the opposite direction of the control surface.
  2. Balanced tab: – They have a mechanical linkage to the control column and they automatically move in the opposite direction to the control surface. The movement of the tab is related to the movement of the control column and not to the load on the control surface. They are not of much use because at high speeds the control surface moves less and therefore tab will move less although the load is more. Therefore, balancing force provided is less.
  3. Spring tabs: In this a spring is attached in a mechanically linkage which prevents its movement at low air speeds. This spring will allow the tab to move only at high airspeeds. Spring tabs may be pre loaded to prevent them from coming into operation until the stick (Rudder) force exceeds the predetermined value. This is done to keep it out of action at low air speed thus avoiding excessive lighting and lack of Controls feel.

Disadvantage: – Sudden change in stick force occurs when the tab comes into action at high speed.

  • During preflight check- on moving the control column only the Control Surface moves, Spring Tab will not move.
  • During preflight check- on moving the control column only the balanced tab will move but spring tab will not move
  1. Anti Balance Tab- They move in the same direction as the control surfaces, thus they spoil the balance. They are provided in Aircrafts which are over balanced. They provide better feel of controls to the Pilot.
  2. Servo Tabs- In servo tabs are connected to the control column but the control surface is not connected to the control column. Depending on the movement of the tab and the airspeed control surface moves. They are Less Effective at low Air speeds.
  • Full and free movement of the controls is done as a part of external check not vital actions. Because by moving the control column only the tab will move, but the control surface does not move.
  1. Trim Tabs: – They have a separate control in the cockpit and they can be deflected to any desired position. They can be Mechanically or Electrically operated.
  2. Tail heavy, which way will the trim tab move.

(a) up          (b) down      (c) No movement

Answer ( C )

  1. Tail heavy, which way the pilot will adjust the trim tab.

(a)      up     (b) Down     (c) No movement.

Answer ( A )

Hydraulic controls: – are of two types

  1. Power Assisted/Hydraulic Assisted Controls
  2. Power Operated/Hydraulic Operated Controls.
  3. Power Assisted Hydraulic: – In this the Hydraulic pressure helps the Pilot to move the controls, Pilot has the feel of the controls.
  4. Power operated: – In this pilot only sends the commands to the hydraulic system to move the controls. The controls are moved by hydraulic actuators/Pubs. In this type of system the Pilot does not have any feel of the controls. An artificial feel is provided to  the pilot.

The feel provided in the cockpit is directly proportional to the amount of deflection of the Control Surfaces.

Hydraulic Controls are of IRREVERSIBLE TYPE i.e. these controls can be moved from the cockpit but the outside gust loads cannot move these controls.

At very high Mach No (Critical Mach No.) due to the formation of the Shock Waves these controls suffer and get jammed due to JACK STALL.

JACK STALL- is when the Adverse pressure are formed near the Hydraulic Actuators due to the formation of the Shock Waves and jam the Control Movements.

Fly By Wire: – In this system Pilot’s command to the hydraulic system are send in the form of electrical pulse and not mechanically through cables and pulleys as was the case in the older Aircrafts.

  1. Rudder Boost System: – This system helps the pilot to maintain the directional control in case of the engine failure by sensing the difference in torque between two engines which arises when one of the engines fails. It applies the correct Rudder by correct amount and maintains direction in case of an engine failure.

Lift Augmenting Devices: –

They are the devices which increases the Lift. There are 3 types of Lift Augmenting Devices:

  1. Stats and Slots
  2. Flaps
  3. Boundary layer control

Slats and Slots: – Slats are the auxiliary aerofoil attached to the leg edge of the wing . On extending the slat a gap appears between the wing and this gap is called as a Slot.

  • When the airflow goes though the Slot it accelerates in velocity, because of the venturi effect and therefore it penetrates further against the Adverse Pressure Gradient.
  • The Transition point moves backwards and therefore it reactivates the Boundary layer.
  • The low pressure area is uniformly distributed over the wing and therefore a greater part of the wing is used to produce lift. The action of the slat is to flatten the marked peak of the low pressure area and to change it with the gradual Pressure Gradient.

The Slat delays the separation of the Boundary layer to about 28° AOA. Whereas a Normal Wing will stall at 150 – 160AOA.

Slats and Slots, increase the CL max or we increase the AOA to increase CL max.

With Slat and Slots, CL max is increased up to 70% and they increase the Stalling Angle by approximately 10°.

Types of slats :

  1. Fixed slats : – They give more lift at lower airspeeds but more drag at higher airspeed.
  2. Automatically Slats: – They automatically get extended at higher AOA because (a) Slats themselves are aerofoil sections and when placed in the airflow they produce lift and get extended (b) with the increase in the AOA, the low pressure area moves forward on to the slat and therefore it gets extended.

Therefore the disadvantage of Fixed Slat being extended at Normal AOA and thereby associated Drag is overcome.

  1. Pilot control: – Pilot has the control over the slats whenever he wishes he can extend the slat.
  2. Built in Slots: – In this Slots are built into the wing tips behind the leading edge. At higher AOA air from below the wing flows through the slot on to the upper surface and thereby re-activates the Boundary layer.

Stall Characteristics with Slats: – AOA of approximately 25° the low pressure envelope over the top of the wing is considerably enlarge. When the AOA is further increased the powerful low pressure envelope collapses resulting in sudden loss of lift and sudden changes in the attitude. This situation can cause a disaster especially when one wing stalls before the other creating a strong rolling moment to set up.

Function of Slats: –

  1. Primary Function: – Flaps give us steeper Climbing Angle.
  2. Slats decrease the Stalling Speed.
  3. Slats give us lower Takeoff Speed
  4. Slats give us lower Takeoff run.
  5. Slats enable Aircraft to climb at lower Air speeds, reducing the distance required for a Climb, which is very beneficial for the Aircrafts operating at Aerodromes with surrounding high terrains.
  6. Slats give us lower Approach and Landing Speed.
  7. Slats give us lower Float period
  8. Landing run reduces because of lower Approach speed and lesser Float period

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Flaps

Flaps are auxiliary Aerofoil attached on the trailing edge of the wing.

Principle of flaps: – Flaps increases the effective camber of the wing.  Therefore increasing the downwash of the airflow over the wing, thereby increasing the Coefficient of Lift.

With flaps the Stalling Angle reduces and CL increases.

Pitching Movements with Flaps: –

(a)        All trailing edge flaps produce an increase nose down pitching moment due to the change in the pressure distribution around the flaps.

(b)        Flaps when lowered increase the down wash on the tail, (airflow meeting the Tail plane), this depends upon size and position of Tail plane and provide a nose up pitching moment.

  • Both (a) and (b) oppose each other which ever aspect is dominant in particular Aircraft along with other considerations like High Wing Aircraft or Low Wing Aircraft etc. will govern whether nose up or nose down trimming is required.

Types of flaps:

  1. Plain Flaps: – Normal flaps
  2. Slotted Flaps: – They form a slot between wing and the Flap. Slotted flaps can be (a) Single Slotted (b) Double Slotted (c) Triple Slotted
  3. Split Flaps: – The bottom part of the wing splits into two.
  4. Fowler Flaps: – The flap first goes back and then goes down.
  • They increase the surface area as well as the camber.
  1. Zap flap: – Similar to Fowler flaps but mechanically different.
  2. Kruegger Flaps/Leading Edge Flaps/Nose Droop: – These Flaps are located at the Leading Edge.
  3. They give more lift at higher AOA
  4. They increase the Stalling Angle.
  • The increase in CL max varies between 50% for a Single Slotted Flap to 90% for the Fowler Flap.
  • Flaps give more lift without increase in the AOA. (Unlike Slats and Slots)
  • Flower flaps give max increase in CL with min increase in Drag.
  • Split flaps causes minimum Pitching moment.
  • Flaps are retracted in the stages only after reaching a SAFE ALTITUDE because there is sudden decrease in lift on retracting flaps, and the Aircraft tends to sink.
    • After Takeoff, before retracting the flaps speed must be increased to FLAP RETRACTION SPEED- VFRI. If the speed is not increased the Aircraft will develop a sink and this sink can be best control by lowering the nose of the Aircraft.
    • There is a maximum speed above which the flaps cannot be used which is known as VFE, because with flaps CL increases (speed is already high and CP will move back, CG is already ahead and this will cause the Structural damage.

    VFE is different for different flaps setting, higher the flaps setting lower is the VFE.

    Flap increase the lift they also increase drag the increase in drag is much more than the increase in the lift. Therefore L/D ratio decreases and thereby Gliding Angle steeps.

    Function of Flaps: –

    1. Primary Function: – Flaps give us steeper Gliding Angle, without increasing the airspeed.
    2. Flaps decrease the Stalling Speed.
    3. Flaps give us lower Takeoff Speed
    4. Flaps give us lower Takeoff run.
    5. Flaps give us lower Approach and Landing Speed.
    6. Flaps give us lower Float period
    7. Landing run reduces because of lower Approach speed and lesser Float period.
    8. Better Control Effectiveness at lower airspeeds. Rudder and Elevator effectiveness increases, because with flaps drag increases and to counteract this Drag more power is required. This increases the Slip Stream effect, on to the Rudder and Elevators. which is more pronounced in Propeller driven Aircrafts
    9. Better forward visibility during Approach.

    Effect of Sweep Back Wing on Flaps- The Flaps reduce the effectiveness of the Trailing edge control surfaces and high lift devices, because there is a decrease in frontal area of flaps with the Sweep Back.

    Boundary Layer Control

    Suction Devices- are installed on the Trailing edge on the top surface of the Wing.

    Blowing Jest- are high speed jets of air, which blows air backward on the Trailing edge of the wing and therefore it attracts the breaking Boundary layer.

    Vortex Generators- are small Longitudinally strips of metal, which are approximately 3” to 4” inches long and 1” to 2” inches high, attached on the top surface of the wing on the leading edge.

  • They induct air of high kinetic energy into the Boundary layer and therefore re activates the Boundary layer. Vortex Generators decrease the Stalling Speed and decrease the Stalling Angle.Normally they are attached on he wings, but sometimes they are also attached on both sides of the FIN, for better Rudder Controllability.

    Blowing Jets and Suction Devices are used in Fighter planes, because they require a lot of engine power.

    All Flaps except the Kruegger Flaps/ Leading Edge Flaps decrease the Stalling Angle, because they displace the effective Chord by changing the shape of the Aerofoil.

    The Stalling Angle is always measured with reference to the Original Chord.

    CLIMBING: –

    Forces acting on the Aircraft during climb is-

    Lift, Weight, Thrust, Drag.

    Where Weight is more than Lift

    And Thrust is more than Drag

    1. Is the angle of climb

    With the increase in Angle of Climb, Q, W sin q increases Cos q decreases. In a 900 climb and W Cos q is ‘O’ and W Sin q is maximum i.e. Lift is O and Thrust = Weight + Drag.

    Rate of Climb- is the height gained in a given time.

    Angle of Climb: – is worked out by using the height gained and the distance traveled in the time. It gives the flight path gradient.

    Climb Gradient:- Slope along which the Aircraft climbs, It depends upon the ROC and IAS.

    Flight Path Gradient: Is worked out by Airspeed and ROC.

    Climb Gradient is not affected by Wind Velocity where as Flight Path Gradient depends upon the Wind Velocity.

    Head wind increases Flight Path Gradient.

    Tail wind decreases Flight Path Gradient.

    ROC does not depend upon Wind Velocity, But H/W and T/W will affect the Ground speed of the Aircraft. This H/W or T/W will give a gradient which is different from the climb gradient called as Flight Path Gradient.

    Factors effecting winds: –

    1. Airspeed: – With the increase in airspeed more power is required to maintain level flight therefore Extra Power available is less, on which ROC depends.
    2. Minimum: airspeed at which level flight can be maintained.
    3. Maximum: airspeed at which level flight can be maintained

    Best rate of climb is available at a point where the difference between power available and power required is maximum.

    1. VMD
    2. Best rate of climb speed

    Best rate of climb sped is slightly greater than VMD

    In a jet Aircraft thrust available curve is a straight line there are two points where the difference between power available and power required is maximum and the jet Aircraft climbs at higher airspeed, for better engine cooling.

    We climb at CRUISE CLIMB SPEED: – speed which is slightly higher than the best Rate of Climb speed for better cooling of the engine with increase in height, we reduce the Cruise climb speed because temperature is less and True Air Speed is more.

    1. Weight: – With increase in weight. Power required increase, therefore Extra Power available decreases, and ROC reduces.
    2. Temperature: – With increase in temperature power output of the engine decreases and power required increases, therefore ROC reduces.
    • In summers ROC is less
    1. Altitude: – With the increase in altitude, power output of the engine decreases, therefore power required increases causing ROC to reduce.

    Service Ceiling: – It is the altitude at which ROC drops to 100ft/minute.

    Absolute Ceiling: – It is the altitude at which ROC drops to zero feet/minute.

    With increase in altitude the range of maximum and minimum airspeed at which level fight can be maintained reduces, the Power Required increases. The Extra Power available reduces.

    Absolute ceiling is the altitude, at which only one airspeed can be maintained for level fit, & no further climb is possible.      

    1. Wind Velocity: – ROC is not affected by Wind Velocity.

    Angle of Climb increases with Head wind.

    Angle of Climb decreases with Tailwind.

    1. Flaps: – With flaps power requirement increases because Drag increases and therefore ROC and Angle of Climb reduces.
    2. Density Altitude: – is the Pressure Altitude corrected for non standard Temperature and Pressure condition.
    • D.A. = P.A. + 120 × (Actual temperature at Pressure Altitude – ISA temperature at that Altitude)
    • Density altitude depends upon temperature. If temperature is higher than ISA then Density Altitude is more than Pressure Altitude.  .
    • With increase in Density Altitude ROC reduces because high density altitude means either more temperature or higher altitude.
    1. Drift Down Altitude: – It is the altitude up to which a Multi Engine Aircraft can descent and maintain altitude in case of a critical engine failure. This is also called as Single Engine Ceiling. Drift down altitude is always approached from the top. It depends upon
    • (1) Temperature (OAT) (2) Aircraft All Up Weight at the time of engine failure.
    • Drift Down Distance: – The distance travelled during descent to Drift Down Altitude factors.

    The Factors affecting are:-

    1. Air speed (maintaining best L/D speed will give more distance traveled)
    2. Head wind or tail wind

    iii. True Air Speed, which will depend upon Temperature and Pressure.

    Drift Down Time: – Time required to descend to Drift Down Altitude. This will depend upon Airspeed and Weight.

    • If Drift Down Distance indicates that you cannot clear the terrain (after engine failure) you will have to reduce the Takeoff weight.
    • Because a higher carried weight requires a higher Drift Down Altitude.
    • If ROC decreases Angle of Climb also decreases, but vice versa is not true.
    • Angle of Climb depends upon Power.
    • Rate of Climb depends upon Thrust.
    • Velocity of Minimum Power (VMP) is slightly less than the Velocity of minimum Drag or Thrust (VMD)

Stability

It is of two types

  1. Static stability           2. Dynamic stability

Static Stability: – It is an inherent quality of the aircraft whereby it has a tendency to come back to its equilibrium position on its own without any pilot action once the disturbing force has been removed.

Dynamic Stability: – It is the stability of the Aircraft in oscillations or oscillatory stability.

Static Stability: – Stability in Pitch Axis is called as Longitudinal Stability or stability of the Aircraft about Lateral Axis is known as Longitudinal Stability.

Stability in Roll Axis is called as Lateral Stability or stability of the Aircraft about Longitudinal Axis is called as Lateral Stability.

Stability in yaw is called as Directional Stability or stability of the Aircraft about the Normal Axis is called as Directional Stability.

Stability is called a Positive Stability: – when the body has tendency to come back to its original position, when disturbed.

Stability is called a Negative Stability: – when the body tends to move away from its original position, when disturbed.

Neutral Stability: – When the body has neither tendency to come back to its original position nor the tendency to move away

Longitudinal Stability: – Factors effecting Longitudinal Stability:

Wings or Movement of Centre of Pressure– E.g. If the nose of the Aircraft pitches up the CP moves forward, Lift will increase, Lift Weight couple will increase and the nose will tend to pitch down (Lift Weight couple has a tendency to put the Aircraft in Nose Down Attitude)

Aerodynamic Center: – It is a fixed point on the wing profile where the Lift coefficient moment remains the same and does not change with the changes in the AOA.

     C.G. is always kept ahead of Aerodynamic Center and therefore C.G. is always ahead of C.P.

 Tail Plane: – It gives maximum Longitudinal Stability. It is made up of a Symmetrical Aerofoil section.

The airflow to the tail plane is affected by the Down wash of the wing. The Tail plane is kept at the lower angle of incidence than the main plane. The airflow to the tail plane comes from a direction different from the airflow to the main plane. Generally the airflow to the tail plane comes from the top, since it is kept at a Negative Angle of Attack, it produces the Negative Lift.

Longitudinal Dihedral: – It is a difference between the Angle of Incidence of the Main Plane and the Tail Plane.

  • Lift provided  by a Tail plane depends upon
  • Angle of incidence of the Tail plane (Longitudinal Dihedral)
  • Shape of the Tail plane
  • Size of the Tail plane
  • Center of Gravity or its position: –
  • FWD C.G. increases Stability.
  • Rearward C.G. decreases Stability.

Neutral Point: – It is that position of C.G. at which stability is Neutral. For Stability C.G. has to be always ahead of Neutral Point.

Static Margin: – It is a distance between C.G. and the Neutral Point. If then static margin increase the Stability will increase.

  • If stability increases controllability reduces.
  • Stability and Controllability are inversely proportional to each other.
  • For every Aircraft the Forward and the Rearward limits of C.G. are defined and for the Stability Aircraft must be loaded within these limits.
  • If the forward C.G. limit is exceeded: –
  • Aircraft will be Stable at all airspeed.
  • Stalling speed will increase.
  • Landing is the most critical phase of the flight, and Aircraft with too Forward C.G. experiences difficulty in flaring out at the time of landing and it can get into a phenomenon of wheel Barrowing.

Wheel Barrowing: It is a phenomenon where the Weight of the Aircraft is on the Nose wheel and not on the Main wheels.

  1. Why Aircraft experiences difficulty in flaring when forward limit of C.G. is exceeded: –
  • Because.1. C.G. is Forward 2. Speed is less, therefore Tail plane effectiveness is less.
  • If the Rearward C.G. limit is exceeded: –
  • Stalling speed decreases
  • AIRCRAFT is more stable at higher air speeds and less stable at lower air speeds.
    • AIRCRAFT can enter Spin from the Stall condition and the recovery is difficult.
    • The distance between C.G. and Tail plane is called as Tail Lever Arm.
    • If tail lever arm is more stability is more, because correcting moments are more and controllability will be less.

    Variable Incidence Tail Plane: – In this there is a separate lever in the cockpit using which the pilot can change the Angle of Incidence of the Tail Plane.

    • Advantage of Variable Incidence Tail Plane: –
    • It reduces the Trim Drag.
    • Trim Drag: – is a drag being experienced by trimmers and flying controls as they move in to the airflow.
    • It provides full elevator movement.
    • It increase the Centre of Gravity Range
    • E.g. If the pilot wants to pitch up a nose, select a slightly lower Angle of Incidence, by moving the Angle of Incidence of the Variable Incidence Tail plane down.
    • Effect of high Mach No. on Longitudinal Stability: – When flying at high Mach No. (close to Critical Mach No.) Longitudinal Stability decreases, because of the formation of Shock Waves. The air behind the Shock Waves is turbulent, boundary layer thickens and when this disturbed airflow hits the Tail Plane, its effectiveness reduces thereby reducing Longitudinal Stability. Under these conditions, an all moving tail plane or Slab tail plane (as it is usually called) is more effective than a conventional Tail Plane.
    • The combination of Positive Static and Positive Dynamic Stability is called Damp Oscillation.
    • For an Aircraft to fly it must have Static Stability it may or may not have the Dynamic Stability.

    Lateral Stability: – Factors effecting Lateral  Stability: –

    1. Wing Dihedral : – A wing with a dihedral when sideslips, the lower wing meets the Relative Airflow at a higher AOA as compared to the upcoming wing. Therefore it produces more lift and tends to come back to its Equilibrium position.

    High Wing AIRCRAFT: – High wing Aircrafts are more Laterally Stable  because of the Pendulous Effect. The high wing Aircrafts are so much laterally stable that an Anhedral is sometimes given to decrease Lateral Stability. E.g. IL-76.

    Sweep back: –

    A wing with a Sweep Back when side slips the down going wing operates under conditions where the effective sweep back is less and it produces more lift where as the up going wing out of the wind experiences increase in the Sweep Back and therefore Lift decreases.

Effective span of the down going wing reduces Aspect Ratio reduces, Induced Drag increases and lift will increase.

The lateral stability due to Sweep Back will be Zero.

The contribution of Lateral stability by Sweep Back is proportional to Coefficient of Lift and the Angle of Sweep Back. Therefore If Sweep Back wing is not producing any lift the Lateral Stability due to Sweep Back will be Zero.

High Keel Surface: – Any vertical surface behind the C.G. is called Keel surface and any vertical surface behind and above C.G. is called High Keel Surface.

 When an Aircraft Side Slips the relative airflow comes from a direction opposite to the Side Slip, hits the Keel Surface and Aircraft comes back to its original position.

E.g. of High Keel Surfaces is FIN, DORSAL FIN,

Low Keel Surface: – Any vertical surface behind the C.G. is called Keel surface and any vertical surface behind and below C.G. is called Low Keel Surface.

E.g. of low keel surface is under carriage VENDRAL FIN.

Effect of High Mach No. an Lateral Stability: – Lateral stability reduces with the increase in airspeed because: –

  • The minor differences in the construction of wings become apparent at high speeds and will change the Lift and Drag being produced by the two wings.
  • At low speeds we use trimmers to maintain Lateral Stability but close to the Critical Mach No. due to the formation of Shock Waves these trimmers become in effective. Therefore Lateral Stability is lost.
  • Directional Stability: – Factors effecting the Directional Stability: –
  • Fin or Vertical Stabilizer: – They increase Directional Stability because of the Wind Cocking Effect or Weather Cocking Effect.
  • Wind Cocking Effect/Weather Cocking Effect- if the nose yaws to the left the fin will go to the right, the relative airflow coming from the right hits the fin and pushes it back to its Equilibrium position.
  • Dorsal Fin: – provides Directional Stability at high angles of Side Slips. Dorsal Fin reduces the effective Aspect Ratio of the Fin, by increasing the Effective Chord of the Wing and increasing the Stalling Angle of the wing.
  • Dynamic Stability : – Dynamic stability can be of various types: –
    • a. Spiral Divergence or Spiral Instability or Grave Yard Spiral
    • b. Directional Divergence
    • c. Dutch Roll
  • Dutch Roll: – It is a condition of Dynamic Instability, in which the Aircraft rolls and yaws simultaneously. Roll is more prominent than Yaw, Roll and Yaw are in opposite directions. It happens in Aircraft’s having large Dihedral Angle and small Fin Area i.e. Aircraft’s having more Lateral Stability and less Directional Stability.
    • It occurs due to
    • Excessive or incorrect use of Rudders.
    • Turbulence.
    • In advertent Lateral Roll tendency of the Aircraft because as altitude increases, True Air Speed increases, Fin Area remains the same but the restoring moment reduces.
    • Dutch roll can be overcome by using the Ailerons against the upcoming wing. i.e. the wing which is trying to come up try to keep it down with the help of the Ailerons.
    • Dutch Roll can also be corrected by using Yaw Dampers.
    • Spiral Instability: – Takes place in the Aircraft having large fin area and less Dihedral i.e. It occurs in Aircrafts having more Directional Stability and less Lateral Stability. It can be corrected by using opposite Rudder. Aircraft is in a nose down Attitude and the turn becomes fighter and tighter with each turn the Spiral Instability reduces with increase in altitude or Aircraft becomes more Spirally Stable as altitude increases.
    • e.g.: If the aircraft is Rolled to the right it will Side Slip to the right, the Dihedral being inadequate will not be enough to cause to Aircraft to retain Lateral position. As the aircraft is Side Slipping the relative airflow on the large Fin area will tend to Yaw the Aircraft to the right and the nose will further go down. If this is not corrected the Aircraft will continue to Side Slip and Yaw in a Tightening Spiral Dive.
      • Directional Divergence: – Results from a Negative Directional Stability. Aircraft develops Side Slip after being disturbed in Roll or Yaw and develops a yawing moment that causes the Aircraft to yaw further in the same direction.
      • Note: – Dynamic stability is the stability in the oscillations or oscillatory stability. Dynamic stability is the movement of the Aircraft with respect to time. If the Aircraft has been disturbed from its equilibrium position and the maximum displacement decreases with time it is said to be in a Dynamic Stability condition.
      • Dynamic Stability:- If the Amplitude of the Oscillations decreases with time it is called Dynamic Stability
      • Dynamic Instability:- If the Amplitude of the Oscillations increases with time it is called Dynamic Instability.
      • Neutral Stability:- If the Amplitude of the Oscillations remain same with the time, it is called Dynamic Neutral Stability.
      • Longitudinal Dynamic Stability: – Oscillations decrease with time in Pitch Axis.
      • Lateral Dynamic Stability: – Oscillations decrease with time in Roll Axis.
      • Directional Dynamic Stability: – Oscillations decrease with time in Yaw Axis.
      • The Stability of an Aircraft with free Floating Control Surface is called Stick Free Stability.
      • The stability of an Aircraft with the controls held in Fixed position is known as Stick Fixed Stability.
      • There are two types of Dynamic Oscillations: –
      • Long Period Oscillations, with Poor Damping Oscillations
      • Short period Oscillations
      • Long period is also know as PHUGOID Mode Oscillations
      • PHUGOID Mode Oscillations is the one in which the airspeed Pitch and Altitude of the Aircraft varies widely but the AOA remains nearly constant. The Motion is so slow that the effect of Inertia forces and Dampening forces are very low.

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Wings Plan form : –

  • Rectangular Wing : – In case of Rectangular Wing CL max is higher at the Wing roots and lower towards the Wing tips, therefore Wing Root Stalls first.
  • Tapered Wing: – CL max is more towards the tips and CL max is less towards roots, therefore Tip stalls first.
  • Note: – Taper Ratio : – Tip Chord/Root Chord
  • Elliptical Wing: – CL max is same along the entire length of the wing. It has the best Stall characteristic because the whole wing stalls at the same time.
  • Aspect Ratio: – Is the ratio between Span and the Mean Chord. It helps in determining the Aerodynamic Characteristic and Structural Weight of the Aircraft.
  • Sweep Angle: – It is the Angle between the line of 25% Chords and at 90° to the root chord
  • Condition: – Taper wings are of two types: –
  • Moderate Taper (e.g. Taper Ratio .5) It has the Lift distribution, which resembles Elliptical Wing. Therefore Stall pattern is similar to that of an Elliptical Wing.
  • Highly Tapered Wing (Taper Ratio .25) it has more Lift Coefficient towards the tip, therefore Tips stall first.
  • Pointed Tip (Taper ratio = O) It has high Lift Coefficient at the tip. Therefore for all practical purposes the pointed tip will stall first.
  • Sweep back: – Effect of sweep Back on Lift. A Sweep Back Wing Aircraft has a lower CL max, as compared to the Normal Wing Aircraft.
  • In a Sweep Back Wing Aircraft as the airflow strikes the Leading Edge it is divided into the two parts: –
  • Span wise component
  • Chord wise component
  • Lift of the wing is produced because of the Chord wise component, since this component is less than the total airflow. Therefore Coefficient of Lift is less than the Straight Wing Aircraft.
  • Sweep Back Wing Aircraft has high value of Critical Mach No. Since the Chord wise component is less than the speed of the total airflow. Therefore Aircraft speed can be increased more before it reaches the Critical Mach No.
  • Therefore Sweep Back improves the High Speed Characteristic of the wing.
  • Effect of Sweep Back on Drag: – Sweep back increases the Induced Drag, because by sweeping the wing CLis reduced and to maintain the same lift AOA has to be increased, as we know as Angle of Attack increases Induced Drag increases.

Because of greater Induced Drag Minimum Drag Speed (VMD) of a Sweep Back Wing is higher than the Straight wing Aircraft and the Approach speed is generally less than the VMD. If a pilot makes a small change in the attitude of the Aircraft (e.g. pitch up). The Lift will increase slightly but there is a large increase in Drag which will result in rapid fall of speed, and to counter this large increase in power is required.

  • On some Aircraft’s to over come this problem high Drag devices e.g. Air Brakes are installed to increase the Profile Drag this results in a Flatter Drag curve and VMD being closer to the Approach speed.
  • At low Air speeds Sweep Back Wing Aircraft is flown at a higher AOA to generate the same Lift (e.g. A Boeing Aircraft decides in a nose high attitude in landing.)
  • Stall characteristic of Sweep Back Wing Aircraft.
  • A sweep Back wing Aircraft by design is the tapered wing Aircraft, therefore it is more prone to wing tip stalling.
  • Sweep Back wing Aircraft is a low Aspect Ratio wing, therefore It has more Wing Tip Vortices which cause the Stalling AOA towards root to increase, therefore it is prone to Wing Tip Stalling.
  • The Span Wise component also causes Wing Tip  Stalling because it thickness the Boundary layer over the outer part of the wing making it more susceptible to separation.
  • Note:– Wing Tip Vortices induce lower AOA at the wing roots this causes Wing Tip Stalling.
  • High Aspect Ratio Wing Aircraft stalls at the lower AOA because it has less Wing Tip Vortices.
  • Forward sweeps: – Sweep back is of two types.
  • Forward sweep
  • Backward sweep
  • Forward Sweep: – gives more advantage in Subsonic/Trans sonic region.

Innovations to prevent Wing Tip Stalling:-

  • Boundary layer fence or Wing Tip Fences- are small Longitudinal stripes attached Chord wise over the wing.
  • Washout towards the Wing Tips: – For the same attitude of the Aircraft the tips will be at lower Angle of Incidence and roots will be at higher Angle of Incidence, therefore tips will reach at the Stalling Angle later or both Tip and Root will stall at the same time.
  • Spoilers on the leading edge towards the roots: – Horizontal strips attached towards the wing roots, they reduce the lift towards the roots.
  • Slats and Slots towards the Wing Tips.
  • Boundary layer control towards the Wing Tips.
  • Winglets prevent the formation of wing tip vortices, therefore they reduce INDUCED DRAG. The Stalling angle towards Root does not increase.

  • Leading Edge Extensions:- They produce Eddies preventing sideway drifting of the Boundary layer.

    • Leading Edge Notches. They also produce eddies, which prevents sideways drifting of the Boundary layer.
    • Swing Wing Aircraft: – From the Cockpit, the Pilot can change the wings Plan form. At low Air speeds wing is shaped as a Straight wing and at high Airspeeds it is shaped as Sweep Back. This type of Aircraft is also called Polymorpheic Wing/Aircraft.
    • Canard Wing: – Canard Wing is an Aircraft which has a Fore plane attached ahead of the Main Wing.
    • Advantages of the Canard Wing: –
    • The For plane Control surface is located ahead and is therefore clear of any Shock Waves that may form in the Main plane.
    • In Aircraft with Tail mounted engines, the C.G. is backward, therefore a Canard Wing is advantageous because it gives a longer moment arm. This enables the Stability and Trim requirements to be met by a Fore plane of small area.
    • At high Airspeeds, Trim Drag will be less.
    • The Trim Drag problem will be reduced, because at high speeds an upload will be required on the Fore plane to trim the Aircraft.
    • In a Canard Wing, if the wing stalls first, the Stability is lost but if the Fore plane stalls first then the control is lost and the maximum value of CL is reduced.
    • The airflow from the Fore plane interferes with the airflow around the main plane (wing) and the vertical Fin. This results in reduction in lift from the Main Wings and loss of Directional Stability (given by the Fin). This stability can be increased by two Vertical Fins instead of one.
    • Stresses on an Aircraft: –
    • Tension- It is due to the Cabin Pressurization of the Aircraft.
    • Torsion: – Twisting Stress.
    • Shearing: – Cutting Stress caused by Centrifugal Force acting on the Propeller.
    • Bending: – It is due to the lift, the wing tends to bend upwards.
    • Compression: –
    • Lift increases Bending Stress increases, for this reason we have a limitation called Maximum Zero Fuel Weight. (MZFW).
    • MZFW: – It is the Weight of the Aircraft above which any load taken must be in the form of Fuel only, since the weight of the fuel in the Wings counteracts the Bending Stress.

    • Sweep Backs Advantage: –

      • It is the value of Critical Mach No.
      • It gives more lateral stability.
      • Disadvantage: –
      • Less Lift at low Airspeed.
      • Tendency of Dutch Roll.
      • Tendency of Wing Tip Stalling.
      • Ailerons disadvantage.
      • Ground Effect: – is an upward pushing effect which an Aircraft feels when flying close to the surface of the Earth.
      • It is due to the air trapped between the wings and the surface of the wing.
      • Ground Effect is more prevalent in the Low Wing Aircraft than the High Wing Aircraft.
      • It is present up to the height of one Wing Span but it is perceptible up to a height of one-half Wing Span.
      • Effects of Ground Effect: –
      • Aircraft tends to get Airborne at a speed less than the normal Takeoff speeds.
      • With in the Ground Effect and as compare to outside the Ground Effect the Aircraft will require a lower AOA.
      • Outside the Ground Effect as compare to with in the Ground Effect Aircraft requires a higher AOA.
      • With in the Ground Effect, Wing Tip Vortices do not from Induced Drag is less.
      • With in the Ground Effect Tail Plane effectiveness decreases, due to decrease in Down Wash of Tail plane.
      • Aircraft feels nose heavy and backward pressure (up elevator pressure or trim up) is required on the stick.
      • Aircraft experiences difficulty in climbing out after Takeoff or Aircraft tends to settle back on the ground after Takeoff.
      • As we move out of the Ground Effect Induced Drag increases and more power is required for Climb, that excess power is not available, therefore Aircraft cannot climb out.
      • Aircraft floats on landing. This floating is common on both during Take off as well as Landings.
      • Ground Effect is of advantage for Short Field Takeoff.
      • Explanation of point (6)
      • Outside the Ground Effect Lift produced is less, therefore to generate the same lift AOA is increased, speed reduces further, to maintain that speed extra power is required but that extra power is not available. Therefore lower the attitude of the Aircraft and have to fly parallel to the ground to increase the airspeed to generate the same lift, giving difficulty in climbing.
      • With in the Ground Effect more Lift is produced.
      • Wake Turbulence: – It is a Turbulence experienced by a Small Aircraft when flying behind the Heavy Aircraft. After Takeoff as the Heavy Aircraft is developing Lift, Air speed is low and it produces Wing Tip Vortices which cause Turbulence.
      • These Wing Tip Vortices increases if the Aircraft is Heavy, Slow and developing Lift.
      • Danger to a Small Aircraft is Induced Roll.
      • Wing Tip Vortices after forming move Outwards.
      • In case of a light Cross wind Upwind Vortex remain on the Runway for longer duration.
      • These Wing Tip Vortices sink downwards after forming at the rate of 500feet/Minute.
      • On the Right Wing Tip the Wing Tip Vortices rotates in the Anti Clockwise direction.
      • On the Left Wing Tip the Wing Tip Vortices rotates in the Clockwise direction.
      • For this reason Small Aircrafts are not allowed to Takeoff after heavy Aircrafts.
      • During Takeoff: – To get Airborne behind the heavy Aircraft, small Aircraft should get Airborne at a point before the Airborne point of Heavy Aircraft and maintain a flight path above the flight path of the Heavy Aircraft.
      • During Landing: – Maintain a Flight path above the Flight path of the Heavy Aircraft and touch down at a point ahead of the touchdown point of the Heavy Aircraft.
      • Centre of Gravity (C.G.): – Center of Gravity of an Aircraft depends upon the loading position of the Aircraft i.e. it depends upon the distribution of the weight.
      • It depends upon the position where the weight is being placed in the Aircraft.
      • C.G. does not depend upon the total weight of the Aircraft.
      • Every Aircraft has the reference datum called as Centre of Gravity Datum and all distances are measured with respect to C.G. Datum.
      • Moment = Weight × Distance
      • Total Moment ÷Total Weight = Distance of C.G. from the Datum.
      • Any Weight which is placed behind the C.G. Datum is given a Positive sign and any Weight which is placed ahead of C.G. Datum is given a Negative sign.

      • To avoid calculation errors in all Modern Aircrafts C.G. Datum is kept at a point ahead of the nose or it is kept in the nose of the Aircraft, therefore all the moments will be Positive.

        • Q. In an Aircraft if all moments are Positive what is the position of C.G. datum.
        • Ans: – Either in the Nose or at the Imaginary point ahead of nose.
        • Station Number: – It is distance in inches behind the C.G. Datum.
        • Every point in an Aircraft is given a Station Number and is written at prominent places inside the Aircraft.
        • Moment Index = Moment ÷ Reduction Factor
        • Reduction Factor: – It is a fixed number by which all the Moments are divided or reduced. In a particular Aircraft a Reduction Factor is fixed.
        • Sum of Moment Indexes ÷Total Weight × Reduction Factor = Distance of C.G. from the Datum
        • If any load is shifted forward, C.G. moves forward and if any load is shifted backwards C.G. moves backward. (Irrespective of the fact that the load could be ahead or behind the C.G.)
        • Weight Shifted ÷ Total Weight = Change in C.G. ÷ Distance Weight Shifted
        • In conventional Aircraft with wing tanks as the fuel is consumed, C.G. moves forward, because tanks are usually located in the Wings, which is behind the C.G.
        • Weight Added/Removed ÷ Total New Weight = Change in C.G. ÷ Distance between Weight and Old C.G.
        • If a new Weight is added ahead of C.G., the C.G. will move forward and vice versa.
        • If a new Weight is removed ahead of C.G., the C.G. will move backward and vice versa.
        • In big Aircrafts C.G. Range is given.
        • In small Aircrafts C.G. Envelope is given.
        • Note: – In case of C.G. Envelope Weight and Moment Index must intersect at a point with in the Envelope, for the Aircraft to be with in the C.G. limits.
        • In very big Aircrafts C.G. Range is given as the length of MAC or Mean Aerodynamic Chord.
        • Propellers: – is a device which converts the Mechanical Energy of the Engine into Forward Thrust.
        • Blade Angle: – is an angle between Chord and the plane of the rotation of Propeller.
        • Helix Angle: – is an Angle between the Relative Airflow and the plane of rotation of the Propeller.
        • Angle of Attack: – is the Angle between Relative Airflow and the Chord of the Propeller.
        • Blade Angle = Angle of Attack + Helix Angle
        • Face of the propeller is the Flat side of the Propeller, which is visible from the Cockpit in a Tractor Propeller Aircraft.
        • Back of the Propeller is the Curved side of the Propeller.
        • Fine Pitch: – Low blade Angle
        • Course Pitch: –   High blade Angle
        • All the parts of the Propeller blades do not rotate at the same Velocity, tips rotate at a higher Velocity and hub rotates at the lower Velocity. Therefore a HELICAL TWIST is given to a propeller, so that all the parts of the Propeller meet the Relative Airflow at the same AOA.
        • Helical Twist is given in such a way that the Propeller is Fine towards the Tip and is coarse towards the Hub.
        • Propelller produces Thrust and Torque: – Thrust acts in the direction of the movement of the Aircraft and Torque acts in a direction opposite to the direction of rotation the Propeller.
        • Helix Angle is also called as the Angle of Advance.
        • It depends upon the Air speed of the Aircraft.
        • Higher the airspeed, higher is the Helix Angle or Angle of Advance.
        • Pitch of the Propeller: – It is the distance moved by the Propeller in one rotation.
        • Theoretical or Geometrical Pitch: – It is the distance traveled by the Propeller in one rotation when Propeller is at O0 AOA.
        • Experimental Pitch: – Distance moved by the Propeller in one rotation when Propeller is producing Zero Thrust.
        • Practical/Effective Pitch: – It is the distance actually moved by the Propeller in one rotation. Practical pitch is always less than the Theoretical pitch.

        • Slip of the Propeller: – It is the difference between the Experimental pitch and the Effective pitch.

          • Blade Angle and pitch are related to each other.
          • If blade Angle is more, Helix Angle will be more, therefore distance moved by the Propeller will be more.  (i.e. pitch will be more)
          • Relative Airflow to the prop depends upon: –
          • Airspeed
          • RPM
          • Effect of RPM (Airspeed constant)
          • If RPM increases AOA will increase and vice versa.
          • Effect of Airspeed (RPM constant)
          • If Airspeed increases AOA decreases and vice versa.
          • AOA governs the Total Reaction or Thrust and Torque.
          • Best Thrust ÷ Torque is available at one particular AOA called the Optimum AOA.