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NAVAL AVIATORS (BASIC AD)

1. At 40,000 ft pressure is 19% of sea level value.

Pressure ratio = = P/PoTemperature ratio = 0 = T/To

Density ratio = =

Is directly proportional to 1/T & PDensity ratio = Pressure ratio

Temp ratio

Kinematic Viscosity = =

2. SPEED MEASUREMENT

At high AOA, static pressure distribution varies a lot & it is difficult tominimize static source error.

At 100 Kts 0.05 psi position error gives an error of 10 Kts.

3. If friction & compressibility effects are not considered, the velocity would decrease to

zero at the aft stagnation point and the full stagnation pressure would be recovered.

4. Maximum velocity & minimum pressure points on the aerofoil do not necessarily

occur at the point of maximum thickness.

5. A/c stalls at same AOA regardless of wt, dyn pressure, bank angle etc. Stall speed

effected by wt etc.

6. Aerofoil properties diff from wing / a/c because of effect of planform. Areofoil two

dimensional.

7. Symm aerofoil zero lift at zero AOA . +ive camber aerofoils zero lift at ve AOA.

Cl max high Bss low.

8. If a/c is in steady flt at (L/D)max, the total drag is at a minimum. Any AOA lower or

higher than this reduces L/D & drag increases.

At L/D max R A G E

P J B J.If L/D max = 15, then a/c covers 15 miles horizontally for each mile of ht lost.

Increase in wt causes increase in glide speed.

9. At low Cls, skin friction drag predominates. At high Cls form or pressure, drag

predominates & the drag coeff varies rapidly with lift coeff.

10. Laminar boundary layer causes reduced skin friction drag vis a vis TBL.

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11. Stall speed - a 2% change in weight causes a 1% change in stall speed.

Eg. At W = 10000Kg VBSS = 100 Kts

So at 11000Kg Vstall = 100 * 11000/10000

= 105 Kts.

12. No appreciable change in load factor below 30 degree angle of bank. Rapid increaseabove 45 degree bank angles.

13. Increase in altitude produces decrease in density & increases the TAS at stall. Alsodue to varn in compressibility , & viscosity effects will cause indicated stall speed to

increase. (effect significant above 20000ft).

14. TE flaps increase Cl at each AOA. Increase Cl max and red stall AOA.

15. Greater drag due flaps advantageous to have steeper appr/ reqt of high power fromengine so better engine response.

16. Plain, split & slotted flap increase camber whereas fowler flap also increase area.

17. Positive camber causes nose down twisting moment great when large camber is

used well aft on the chord. Higher the Cl max produced, greater the twisting

moments. In fowler flap maximum.

18. Flaps less effective on wings with thin sections , sweep back also reduces

effectiveness of flaps.

19. Slot no change in camber. Cl max inc achieved at higher AOA ie stall is delayed.

Due to this higher AOA , so landing gear is complicated.

20. Slot negligible change in pitching moment & no sig change in drag at low AOA.

21. Tailless a/c can only utilize high lift devices, which have little effect on pitchingmoments. Slots & slats used to increase Cl max in high sp a/c when compressibility

effects are considerable.

22. LE devices more effective than TE devices on swept wings as flow pattern is

controlled. Slats most effective on wings of low thickness & sharp leading edges.

23. Boundary layer control increases AOA for maximum lift. So best to combine withflaps which red AOA. Both negate each other.

24. Flap retraction change in trim, reduced drag, greater acceleration and increase inattitude to maintain Cl.

25. First 50% of flap deflection causes more than half the total change in Cl max & last50% causes more than half of total change in Cd.

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26. For jets since little effect on induced flow. Hence, stall AOA power on & off is same.However power on stall speed is less. In props, max lift AOA incr & stall speeds

reduced.

27. Symm aerofoil at zero lift no pitching moments about AC.

Location of AC not effected by camber, thickness & AOA.

Generally AC located between 23 & 27% of chord.

+ive camber has ve Cm AC ie CP behind AC & vice versa.AC should be behind CG for long stability.

Subsonic AC at 25%. In supersonic flow at 50% chord point. So a/c in transonic flt

experiences large changes in long stability (because large changes in AC).

28. Flow in turbulent layer allows particles to travel from one layer to another producing

an energy exchange. Heat transfer easy. So frost, water soil film removed easily aft oftransition point.

LAMINAR TURB

(1) Low vel next to surface (1) High vel next to surface

(2) Vel profile not so full (2) Fuller velocity profile(3) Less thick (3) Thicker

(4) Lesser KE (4) Higher KE(5) Less skin friction (1/3rd of turb SF) (5) More skin friction

(6) Frost water & oil film

removed

29. Cds for either turbulent or laminar flow decrease with increase in RN since vel grad

decreases as BL thickens & redn in vel grad => less skin friction drag.

30. RNo < million laminar boundary layer.

Rno between 1 & 5 million partly laminar & partly turb.Rno > 10 million boundary layer predominantly turbulent.

31. Greater changes in Cl max occur where laminar layer predominates. Higher the

velocity earlier the transition point.

32. Vortex generators applied to surfaces of a high sp a/c may delay compr buffet. They

create a strong vortex which introduces high velocity, high energy air next to thesurface to reduce or delay the shock induced separation.

33. Variation in ad characteristics with Reynolds no. is termed scale factor.

34. Increase in Rno increases Cl max & red Cd of the section .

35. Rno =

36. Roughness on ball etc. causes earlier transition and del sepn hence lesser form drag.Beneficial if redn in form drag more than increase in SF drag.

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37. Aerofoil section 2 dimension. Planform 3 dimensions.

38. Due prodn of lift downward vel imparted to flow. Downward vel created at AC is one

half (w) the final downward velocity imparted to the airstream (2w). Downwardangle (epsilon) downstream.

Sections of the wing operate in an average relative wind which is inclined downward

one half the final downwash angle.

Therefore, Wing (AOA) = AOA (secn) + AOA (induced).

39. Load factor 2 => Cl is doubled and Ind Drag is 4 times as great.

LF 5 Ind drag 25 times.Ind drag is directly proportional L*L

Di2 / Di1 = (L2/L1)* (L2/L1)

10% higher gross wt increase ind drag 21%.

40. Di2 / Di1 =

Ind drag increases with alt.Eg. At 40000ft ( = 0.25) a/c would have 4 times as much ind drag than at sea level

(at constt TAS). However, at same EAS ind drag will not vary.

41. Induced drag decreases with increase air speed.

Di2/ Di1 = (V1 / V2) * (V1 / V2) (L, S, AR, are constt).

At high speeds induced drag less.

Jet -> low level high speed ind drag 1% of total drag. For same a/c just above stallsp ind drag approximately 75%.

42. Wing with low AR is less sensitive to changes in AOA & requires high AOA for max

lift.

43. CDwing = CDi + CDsecn.

44. Flare reqts set the best glide speeds of low aspect ratio airplanes above the speed for

(L/D) max. Addl sp provides favorable margin of flare capability for flameout

landing from steep glide path (low A, low (L/D) max, low glide ratio).

45. Due to prop wash a prop powered a/c may have its most undesirable stall characs

during the power on stall rather than power off stall.

46. Since the upper wing surface has the most critical press grads, a low wing posn on a

circular fuselage would create a greater interference drag than a high wing position.

47. Total Drag = CDp + Cdi

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Both CDp & Cdi vary with Cl. Varn in CDp can be considered with Cdi by a constt

factor called airplane efficiency factor, e,CD = Cdpmin + Cdi / e

Cdpmin Parasite drag coeff (drag at zero lift)

Cdi / e - Ind drag coeff (due to lift)Para drag 50% wing, fuselage & nacelles

40% & tail 10%

e varies from 0.6 to 0.9

Fairly accurate drag calculation upto 70% of Clmax due sharp increase in parasitedrag above that AOA. Valid for subsonic performance. Max varn for low asp ratio &

sweptback wings. Also for high sp compr effects destroy this.

D = Dp + Di Dp = Cdpmin qsDi all drag due to lift

Dp parasite drag completely indp of lift (also called barn door drag)

Dp = fq f = equivalent parasite area , q = dyn pressure

f = CdpminS (gives an impression of barn door size).

Parasite drag can be appreciated as the result of the dyn pressure, q, acting on theequivalent parasite area, f. (this area defined as hypothetical surface with CD = 1.0

which produces same parasite drag as the airplane.) Analogy barn door in airstream which is equivalent to the airplane.

This are 4sqft for fighter to 40 sqft for tpt a/c.

If f is incr Dp increases in direct proportion.

Also,

Dp2 / Dp1 =

Parasite drag reduces with alt at constt TAS. However, if at constt EAS Dp remainssame.

Dp2 / Dp1 = (V2 /V1) * (V2 / V1)

Parasite drag of most importance at high speed. At low speed 25% of total drag is

parasite & at high speed is nearly 100% of total drag. This emphasizes great

aerodynamic cleanliness.Wing contribution nearly half the total parasite drag.

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NAVAL AVIATORS HIGH SPEED

1. Speed of sound is the rate at which small pressure disturbances will be propagated

through the air and this propagation speed is solely a function of air temperature.Temp red speed of sound red.

2. 3 types of waves-(i) Oblique Shock Waves (compression).

(ii) Normal Shock Waves (compression).

(iii) Expansion Wave (no shock).

3. OBLIQUE SHOCKAir slowed. Mno red but flow still supersonic. Flow direction changed. Staticpressure behind wave increases. Density of air behind wave increases. Some

energy lost due conversion to heat energy.

Shock wave angle red as MNo increases and increases with increase in wedge

angle. If this increase is too much, shock wave detaches at leading edge (this will

lead to subsonic flow behind shock wave).

4. Three dimen flow weaker shock, less change in pressure & density at same Mno.

5. NORMAL SHOCK WAVEWave detaches from leading edge. Shock wave perpendicular to upstream flow.Behind wave flow subsonic.

Whenever a supersonic air stream is slowed to subsonic without a change in

direction, a normal shock wave will form. Transfer of flow from subsonic to

supersonic is smooth & no shock wave forms.Local Mno behind wave is approximately equal to reciprocal of Mno ahead of the

wave. 1.25 M -> 1/1.25 = 0.8 MAirflow direction unchanged.

Static pressure behind wave increases greatly, density increases greatly.

Energy of air stream reduces greatly.

6. EXPANSION WAVE

Flow turns around a corner.(i) Air Stream Accelerated.

(ii) Flow Direction Changed.

(iii) Static Pressure Decreased.

(iv) Density decreased.(v) No shock & no loss of energy.

In three dimensions, principle diff is the tendency for the static pressure tocontinue to increase past the wave.

7. Wave drag varies as the square of the thickness ratio if thickness is reduced 50%,the wave drag is reduced 75%.

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8. The Mno which produces a sharp change in the Cd is termed the force divergence

mach no & usually exceeds Mcrit by 5 to 10%. Condition also referred to as dragdivergence or drag rise.

9. A decrease in downwash on horizontal tail will create a diving moment & a.c willtuck under.On swept wing, Cp shift contributes to trim change root shock more

Cp aft & adds to diving moment. Shock formn at wing tips moves Cp fwd so

climbing moment & tail downwash which contribute to pitch up.

10. Compressibility effects at high alt & low dyn pressure of little consequence.

However, if at low alt & high dyn pressure will cause greater trim changes, heavier

buffet etc.

11. To minimize supersonic wave drag - low t/c ratio, sharp LE, high fineness ratio.

12. Rearward movement of AC implies long stab increases so powerful control surfaces

required to achieve controllability.

13. Shock wave on a/c creates problems outside the imm vicinity of the a/c surfaces.

While shock waves a dist away can be quite weak, the pressure waves can be ofsufficient magnitude to create an available disturbance (SONIC BOOMS).

14. Wave drag coeffs vary as square of t/c ratio.

15. For a wing in supersonic flow, no upwash exists ahead of the wing, 3D effects areconfined to the tip cones & no local induced velocities occur along the span between

the tip cones.

16. The supersonic drag due to lift is a function of the section and AOA while the

subsonic induced drag is a function of Cl & aspect ratio.

17. SUPERSONIC CONFIGURATION

(i) Wing low aspect ratio, highly tapered, sweep back, low thickness ratio&

sharp leading edges.

(ii) Fuselage & nacelles of high fineness ratio.(iii) Tail surfaces similar to wing. Controls fully powered & irreversible with

all moving surfaces.

(iv) Gross cross section may be area ruled to reduce interference drag.

18. At stagnation points max change in velocity & thereby temperature.

19. In supersonic flight ram temperature rise which is independent of altitude & is afunction of TAS.

Act temp = temp rise + amb air temp.

Low alt flt at high Mnos the highest temps.

20. High temp prod redn in str of aluminum alloy so titanium alloy/ stainless steel are

required to be used.