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Incompressible Flow > Chapter 4 > Flashcards

Chapter 4 Flashcards

1
Q

Lift and moments on an airfoil are due mainly to ______, which is dictated by ____ flow.

A

Pressure distribution
Inviscid flow

Ref: Anderson, Pg 324

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2
Q

Define the mean chamber line.

A

The locus of points halfway between the upper and lower surfaces as measured perpendicular to each point.

Ref: Anderson, Pg 326

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3
Q

Define the chord.

A

The straight line that connects the leading and trailing edges.

Ref: Anderson, Pg 326

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4
Q

Define “chamber.”

A

The maximum distance between the mean camber line and the chord line, measured perpendicular to the chord line.

Ref: Anderson, Pg 326

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5
Q

Define “thickness” with respect to an airfoil.

A

The distance between the upper and lower surfaces measured perpendicular to the chord line.

Ref: Anderson, Pg 326

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6
Q

What is the first digit of the NACA four series airfoils?

A

Maximum chamber in hundredths of the chord.

Ref: Anderson Pg 327

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7
Q

What is the second digit of the NACA four series airfoils?

A

Location of the maximum chamber in tenths of the chord from the leading edge.

Ref: Anderson Pg 327

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8
Q

What do the last two digits of a NACA four series airfoil represent?

A

Max thickness in hundreds of the chord.

Ref: Anderson Pg 327

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9
Q

A symmetric airfoil has ___ chamber.

A

No

Ref: Anderson Pg 327

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10
Q

The aerodynamic forces and moments on a body are due to ____ and ______.

A

Pressure distribution, and shear stress distribution, integrated over the body surface.

Ref: Anderson Pg 19 & 23

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11
Q

With respect to airfoils, explain the rationale for the sign convention for moments.

A

Moments that tend to increase the angle of attack (pitch up) are positive.

Ref: Anderson Pg 23

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12
Q

What is the formula for dynamic pressure and what are it’s units?

A

q_i := (1/2)rhoV_i^2
Units: Pa

Ref: Anderson Pg 24

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13
Q

On the graph of cl vs alpha, up to what point is the graph linear?

A

Just before cl_max and then stall.

Ref: Anderson Pg 329

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14
Q

With respect to aircarft performance and handling, the higher cl_max, the lower the _____.

A

Stalling Speed

Ref: Anderson Pg 329

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15
Q

For a symmetric airfoil, what is the angle of attack at zero lift? For a non-symmetric airfoil?

A

Zero.
A negative number.

Ref: Anderson Pg 329

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16
Q

True or False
Inviscid flow theory allows the calculation of Lift slope and the angle of attack at zero lift. It does not predict cl_max.

A

True

Ref: Anderson Pg 329

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17
Q

Is the lift slope dependant upon Reynolds number? Is cl_max?

A

No
Yes

Ref: Anderson Pg 329

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18
Q

Is the moment coefficient for an airfoil sensitive to Re?

A

No except at very large angle of attack.

Ref: Anderson Pg 330

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19
Q

c_d is the _____.

A

Profile drag coefficient.

Ref: Anderson Pg 332

20
Q

The profile drag is the sum of ____.

A
  1. Skin friction drag.
  2. Pressure drag due to flow separation (form drag)

Ref: Anderson Pg 330 and 331

21
Q

Is the profile drag coefficient sensitive to Re?

A

Yes, it is a viscous effect.

22
Q

Define the aerodynamic center.

A

The point on the airfoil about which the moment is independent of the angle of attack.

Ref: Anderson Pg 331

23
Q

In general moments on an airfoil are a function of _____, unless the moment is taken about the ____.

A

Angle of attack
Aerodynamic center.

Ref: Anderson Pg 331

24
Q

The higher the value of L/D_Max the more ____ the airfoil.

A

Efficient

Ref: Anderson Pg 333

25
Q

What direction of flow constitutes a negative value of Circulation? A positive value?

A 
Negative = CCW
Positive = CW

(Same as aerodynamic moments)

Ref: Anderson Pg 333

26
Q

The circulation around a point vortex is equal to the _____________.

A

Strength of the vortex.

Ref: Anderson Pg 335

27
Q

The local jump in tangential velocity across the vortex sheet is equal to ____.

A

The local sheet strength. gamma = u1 - u2.

Ref: Anderson Pg 336

28
Q

Why is replacing the airfoil boundary with a vortex sheet appropriate?

A

Because there is a thin boundary layer on the surface due to the action of friction between the surface and the airflow. This boundary layer is highly viscous and large velocity gradients produce substantial vorticity.

Ref: Anderson Pg 337

29
Q

State the first Kutta Condition.

A

Steady flow over an airfoil at a given angle of attack, nature adopts a particular value of Gamma such that the flow leaves smoothly at the trailing edge.

Ref: Anderson Pg 340

30
Q

State the second Kutta Condition.

A

If the trailing edge angle is finite, then the trailing edge is a stagnation point.

Ref: Anderson Pg341

31
Q

State the third Kutta Condition.

A

If the trailing edge is cusped, then the velocities leaving the top and bottom surfaces at the trailing edge are finite and equal in magnitude and direction.

Ref: Anderson Pg 341

32
Q

According to the Kutta Condition, what is the vortex strength at the trailing edge?

A

Zero.

Ref: Anderson Pg 342

33
Q

Why is lift easily investigated by inviscid theories?

A

Pressure is the dominant force in the generation of lift, and shear stress has a negligible effect.

Ref: Anderson Pg 342

34
Q

How does nature enforce the Kutta condition?

A

Via friction in the boundary layer.

Note: Interestingly enough, if friction did not exist the Kutta condition could not be enforced and lift would never be generated.

Ref: Anderson Pg 342

35
Q

State Kelvin’s circulation theorem.

A

The time rate of change of circulation around a closed curve consisting of the same fluid elements is zero.

Ref: Anderson Pg 343

36
Q

True or False

The circulation around the airfoil is equal and opposite to the circulation around the starting vortex.

A

True

Ref: Anderson Pg 345

37
Q

What is the lift slope for a thin symmetric airfoil?

A

2*pi

Ref: Anderson Pg 352

38
Q

For a symmetric airfoil, where is the location of the center of pressure, aerodynamic center?

A

Both are at c/4.

Ref: Anderson Pg 352

39
Q

True or False

The quarter chord is the center of pressure for a cambered airfoil.

A

False.

Ref: Anderson Pg 360

40
Q

For a cambered airfoil, where is the theoretical location of the aerodynamic center?

A

c/4

Ref: Anderson Pg 360

41
Q

The center of pressure for a cambered airfoil varies with _____.

A

The angle of attack.

Ref: Anderson Pg 361

42
Q

True or False

The lift on an airfoil is primarily due to the integrated pressure distribution exerted on its surface.

A

True

Ref: Anderson Pg 379

43
Q

True or False

The shear stress distribution acting on the airfoil when integrated in the lift direction is usually negligible.

A

True

Ref: Anderson Pg 379

44
Q

What phenomenon is completely responsible for aerodynamic drag?

A

Viscosity.

Ref: Anderson Pg 380

45
Q

Drag acts through what two mechanisms?

A
  1. Skin friction drag due to shear stress.
  2. Pressure drag due to flow separation (aka form drag).

Ref: Anderson Pg 380

46
Q

Form drag is the same as _____. It results from ____.

A

Pressure Drag
Flow separation

Ref: Anderson Pg 380

47
Q

What is the formula for the boundary layer thickness over a flat plate in laminar flow?

A

delta = 5x/sqrt(Rex)

Ref: Anderson Pg 381

Incompressible Flow (6 decks)

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