Basics Flashcards

1
Q

Moment Coefficient equation

A

Cm = M / q.S.L

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
2
Q

What is aerodynamics

A

The behaviour of air flow around an object and the effect it has on it

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
3
Q

List things to think about when designing in 2D

A
  1. Assumes infinite wing span
  2. Inaccurate for finite wings
  3. Approximation is invalid at wing tips
  4. Approximation is okay for wings with large AR
  5. For swept wings, the section and the coefficients need to be transformed accordingly
How well did you know this?
1
Not at all
2
3
4
5
Perfectly
4
Q

Reynolds number =

A

Re = pVl / u

p - density
u - viscocity
l - reference length

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
5
Q

Mach number =

A

M = V/a

a - speed of sound

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
6
Q

If 2 geometries are similar, the air flows are said to be …

A

Dynamically similar

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
7
Q

List properties of Flow similarity

A
Cl1 = Cl2
Re1 = Re2
M1 = M2
  • Streamline patterns are the same
How well did you know this?
1
Not at all
2
3
4
5
Perfectly
8
Q

Why is flow similarity useful

A
  • Can do scaled tests in wind tunnel (cheaper)
How well did you know this?
1
Not at all
2
3
4
5
Perfectly
9
Q

To match Re of real sized objects, smaller models have to either …

A
  1. Have a higher density/pressure
  2. Lower temp/viscosity
  3. Higher velocity
How well did you know this?
1
Not at all
2
3
4
5
Perfectly
10
Q

List compressibility categories and there related Mach numbers

A

Incompressible - M<0.3

Subsonic - 0.3

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
11
Q

For inviscid compressible flow what is Cl a function of

A

Cl = f(Mach)

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
12
Q

For viscid incompressible flow what is Cl a function of

A

Cl = f(Reynolds)

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
13
Q

What is the centre of pressure

A

The location where the resultant of a distributed load effectively acts on the body

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
14
Q

Equation for Centre of Pressure

A

Xcp = - Mle / n

n - normal force
Mle - Moment

For small incidence:
Xcp = Mle / L

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
15
Q

How does centre of pressure move with reducing lift

A

Xcp moves downstream as lift reduces (can be outside aerofoil)

How well did you know this?
1
Not at all
2
3
4
5
Perfectly
16
Q

What is the aerodynamic centre

A

Fixed point about which the moment is independent of incidence (about 25% of chord length behind leading edge)

17
Q

How do thickness and viscous effects change the aerodynamic centre

A

moves aerodynamic centre upstream

18
Q

How does compressibility effects change the aerodynamic centre

A

moves aerodynamic centre downstream

19
Q

Equation for vorticity

A

Vorticity = ∇ x V(velocity vector)

20
Q

What does vorticity equal if flowfield is irrotational

A

vorticity = 0

21
Q

Define circulation

A

the negative line integral of the velocity vector around a curce

22
Q

For an irrotational flowfield, the circulation of a closed curve equals =

A

zero (contained in the flowfield)

23
Q

What are the boundary conditions for inviscid flow

A

for inviscid flow, the flow cannot penetrate the solid surface i.e. velocity vector is tangent to the surface

24
Q

What is Kelvin’s circulation theorem

A

circulation around a closed curve consisting of the same fluid elements remains constant as the fluid elements move through the flow

25
Q

What is Kutta’s condition

A

Due to viscous effects, the flow has to leave the trailing edge smoothly. This results in circulation and therefore lift

26
Q

The lift generation per unit span on an aerofoil is given as =
(Kutta-Joukowski theorem)

A

l = ρVΓ

Γ - circulation

27
Q

What is the origin of the circulation around the aerofoil

A
  1. Before aerofoil moves there is no circulation
  2. As aerofoil movies a vortex rolls up and around the sharp trailing edge and is flushes downstream
  3. This starting vortex has a circulation with the same magnitude but opposite sign as that around the aerofoil
28
Q

What is the thin aerofoil theory

A

A thin aerofoil can be simulated by a vortex sheet places along the camber line

29
Q

Equations for a symmetrical thin aerofoil

A
Cl = 2πα
dCl/dα = 2π
Xcp = Xac = Xc/4
30
Q

Equations for a chambered thin aerofoil

A
Cl = 2π(α-α(l=0))
dCl/dα = 2π
Xac = Xc/4
31
Q

Explain panel methods to solve for the incompressible inviscid flow field

A
  1. Aerofoil surface divided into ‘panels’
  2. Elementary solutions of unknown strength are linearly distributed on these panels
  3. velocity/boundary conditions give n+1 equations and n+1 unknowns
  4. Solutions of linear system gives functions for the potential flow fields
  5. From this, work out velocity vector and then pressure from Bernoulli;s equation
  6. Deduct forces and moments from surface pressure
32
Q

What application can the panel method have

A

used for 3D geometries

33
Q

Explain digits in NACA-4-series

A

NACA2412
2 - 2%chord length is the max camber
4 - 40%chord length is the position of max chamber
12 - 12%chord length is the thickness

34
Q

Explain digits in NACA-6-series

A
NACA65-218
6 - indicates 6 series
5 - 50%c is the minimum pressure
2 - 0.2 is the design lift coefficient
18 - 18%c is the thickness
35
Q

Describe stall

A
  • due to viscous effects a boundary layer forms on the aerofoil surface
  • as the streamwise adverse pressure gradient increases the boundary layer reduces
  • at certain point the boundary layer separates
  • after separation from the upper surface, the lift drops and drag increases
  • this is known as stall
36
Q

As Re increases Cl …

Why

A

Clmax increases

This is because at high Re, the boundary layer is more resistant to separation

37
Q

As Camber increases Cl …

A

Clmax increase

  • shifts curve up and left
  • stall angle reduces
  • increased Cdmax too
38
Q

As thickness increases Cl …

Why

A

Clmax increases up to a certain point, then reduces again.
This is because thin aerofoil have smaller leading edge radius, leading to higher adverse pressure gradient
Drag penalty with increased thickness