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Principles of Flight > Subsonic Aerodynamics > Flashcards

Subsonic Aerodynamics Flashcards

1
Q

What are the conventions for questions?

A

Effect of a variable under review is the only variation that needs to be addressed.

Different bits of the course will be examined in one question.

Convert knots into m/s and know conversions by heart.
1 knot = 0.514m/s = 1/2 for calcs

Inflow to propellors is the TAS

Not considering fly by wire

Always subsonic, always incompressible.

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2
Q

List SI units.

A 
Mass -kg
Acceleration - m/s2
Weight - N or kgms-2
Velocity - m/s
Density - kg/m3
Temp - F,C,K
Pressure Pa or N/m or kgm-1s-2
Force - Newton
Wing loading = Pressure I think. Newtons per square metre
Power = J/s or Watt - is work per time
Work - Nm
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3
Q

State and interpret Newtons laws

A

1st - Same speed if no resultant force
2nd - F=ma, so acceleration is proportional to resultant force
3rd - Equal but opposite reaction force

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4
Q

Explain air density

A

Mass per unit volume

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5
Q

List atmospheric properties that affect density

A

Temperature and pressure, gas constant

P=RHORT

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6
Q

Define static Pressure

A

Pressure at any point in fluid not associated with its motion

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7
Q

Define dynamic pressure, and its formula

A

Pressure attributed to a fluid’s movement

q = 1/2 * RHO * V^2

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8
Q

Apply formula for given altitude and speed

A

Get altitude, use density chart
10000 0.903 kg/m3
22000 0.609 kg/m3
40000 0.302 kg/m3

Use 1/2 * RHO * V^2

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9
Q

State Bernoulli’s equation

A

Conservation of energy in closed system.

At throat, Velocity increases so pressure decreases, dynamic pressure increases

Remember converging and diverging nozzle

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10
Q

Define total pressure. Apply it to a venturi

A

P+q=P0

AKA stagnation pressure.

In venturi, along streamlines total pressure is constant. Static and dynamic vary.

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11
Q

Describe how IAS is gotten from Pitot-static tube

A

Stagnation (total) pressure and static pressure determined. A diaphram deflects showing the difference. Deflection proportional to dynamic pressure.

Gauge is designed to correct for 1/2RHOV^2

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12
Q

Describe relationship between P,T,RHO for air

Explain continuity

A

P/(RHO*T) = constant (may not be exactly right bit will work for question.

This is conserved.

Mass flow is also conserved/

m dot = AVRHO

A*V is conserved downstream

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13
Q

Define IAS,CAS,EAS,TAS

A

TAS is actual speed through air

IAS is what the pitot shows. This has errors
CAS -calibrated airspeed. This fixes instrument error and position error
EAS - fixed compressibility error

If we can get the TAS, this fixes the density error

ICE T is pretty cool drink
IAS,CAS,EAS,TAS instrument, position, compressibility,density

TAS is used in full Lift equation where density is only taken into account by changing the RHO.

L =0.5RHOTAS^2 *S *Cl

IAS^2 ppl 0.5 * RHO *TAS^2

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14
Q

Describe steady and unsteady airflow

A

Steady - straight streamline. smooth, predictable. Takes path of least resistance

Unsteady - broken streamline

Steady - P,T and RHO are independent of time

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15
Q

Explain concept of streamline

A

The path a bit of fluid takes. Has constant total pressure

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16
Q

Describe airflow through a stream tube

A

A streamtube defines a volume of air. No flow in and out of tube

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17
Q

Describe force resulting from pressure distribution round an aerofoil

A

High pressure below, low pressure on top, so resultant force upwards

18
Q

Resolve resultant force into lift and drag

A

Lift = Perpendicular to free stream flow

Drag - Parallel to free-stream-flow

19
Q

Describe all parameters of a wing

A

Leading edge radius
Chord - is just length
Chord line - straight line from leading to trailing edge
Camber line - line always equidistant from top and bottom surface. Goes through centre of all inscribed circles on wing
Camber - max different of chord line and camber line
AOA - angle between RAF and chord line
Symmetrical - chord line = camber line
Thickness/chord ratio
Nose radius

20
Q

Explain difference of 2D and 3D wing

A

Spanwise flow.

Both are ideal fluid and we will consider viscosity

21
Q

Explain streamline pattern around aerofoil

A

Top - close together, fast, less static pressure

Bottom - further apart/same as free stream. Higher static pressure

22
Q

Describe stagnation point

A

V = 0. Static pressure - Total pressure. So high pressure

23
Q

Explain the effect of AOA changes on stagnation point

A

Will change its position accordingly…

24
Q

Explain local pressure changes with camber and AOA

A

On cambered - increasing AOA moves centre of pressure forwards. So minimum pressure moves forwards (top of wing)

Symmetrical - increasing AOA increases total reaction but CP does not move

Pattern of pressure distribution determines the position of the centre of pressure. It is affected by the angle of attack and the camber but not the flow speed.

25
Q

Describe effect of AOA increase and camber increase on aerofoil

A

Up to stall point, bothe increase everything

Top - Faster flow, lower pressure
Bottom - Slower, higher pressure

More lift, more drag. Angle of resultant lift increases, so higher proportion of drag

26
Q

Explain the centre of pressure

A

Point through which the resultant force acts.
Affected by camber and AOA, but NOT by airspeed
Assumed to be at 25% of chord line

If question says airspeed increases in straight and level flight, this means that the Lift must remain constant, so the angle of attack must decrease, with all attributable affects.

Magnitude of total lift force is constant.

27
Q

Explain aerodynamic centre

A

Where the pitching moment coeff does not change for changes in AOA. Pitching moment about this point does not change as AOA increases.

This is taken to be 25% too.

28
Q

List two physicsal phenomena that cause drag.

A

Skin friction drag - as fluid has viscosity, it sticks to the surface and boundary layer

Form drag - produced by the effect of viscosity on the pressure distribution around the object

29
Q

Why do drag and wake cause loss of momentum?

A

Some energy is lost to the object, and Bernoulli is therefore not upheld. (Ptot != const).

As the conversion from dynamic to static pressure behind object is not upheld.

This is a wake of turbulent flow.

30
Q

Explain influence of AOA on lift. Symmetrical and Cambered

A

Increases linearly until stalling or critical point, then decreases.

Symmetrical - zero lift at 0 AOA.

Cambered - has lift at AOA. Has zero lift at the zero-lift AOA

31
Q

Why are coefficients used in general?

A

They make the equations fit reality. Cl describes the natural ability to produce lift.

32
Q

Describe the lift formula and perform simple calcs

A

Lift = 1/2 * RHO * S * V^2

33
Q

Compare high and low speed aerofoils

A

Low speed - high camber/ high thickness ratio for max lift

High speed - low camber/ low thickness ratio for max lift

34
Q

Describe drag formula

A

Same as lift except it has a coefficient of drag

35
Q

Discuss effect of shape of body on drag

A

More frontal area - more drag

36
Q

Define the following terms

A 
span
tip and root cord
taper ratio
wing area
wing planform
mean geometric cord  -average chord
mean aerodynamic chord - imaginary wing that has the same pitching characteristics and has similar longitudinal stability properties
aspect ratio - can work out with area and chord
dihedral angle
sweep angle
wing twist
angle of incidence
37
Q

Explain causes of spanwise flow

A

Because of wingtip vortices the air keeps moving on top/bottom of the wing.

Interaction of these causes small vortices at trailing edge.

More effect near to wing tips

38
Q

Downwash

A

Heavy near tips due to vortices

39
Q

Boundary layer

A

Laminar - No flow perpendicular to surface
sheets of increasing speed
2mm thick
less skin friction drag
less energy at surface so can be overcome by a negative pressure gradient
mean speed is lower as not as far as turbulent

Turbulent BL
Flow in all directions (mixing)
Greater skin friction drag
more kinetic energy at base so resists separation
Is deeper 2-3cm
Mean speed is faster in BL as further to go

40
Q

What are 3 points of flow on win. What happens when change AOA

A

Stagnation point moves down
Transition point moves forward
Separation point moves forward

Due to adverse pressure gradient. Bad question no. 15544

41
Q

Explain 2D drag

A

Profile drag = skin friction drag + form drag.

I think that the entirity of the profile drag is calculated via

D = 0.5 RHO V^2 * S * Cd

S here is the frontal area!!

Principles of Flight (1 decks)

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