Basics Flashcards
Moment Coefficient equation
Cm = M / q.S.L
What is aerodynamics
The behaviour of air flow around an object and the effect it has on it
List things to think about when designing in 2D
- Assumes infinite wing span
- Inaccurate for finite wings
- Approximation is invalid at wing tips
- Approximation is okay for wings with large AR
- For swept wings, the section and the coefficients need to be transformed accordingly
Reynolds number =
Re = pVl / u
p - density
u - viscocity
l - reference length
Mach number =
M = V/a
a - speed of sound
If 2 geometries are similar, the air flows are said to be …
Dynamically similar
List properties of Flow similarity
Cl1 = Cl2 Re1 = Re2 M1 = M2
- Streamline patterns are the same
Why is flow similarity useful
- Can do scaled tests in wind tunnel (cheaper)
To match Re of real sized objects, smaller models have to either …
- Have a higher density/pressure
- Lower temp/viscosity
- Higher velocity
List compressibility categories and there related Mach numbers
Incompressible - M<0.3
Subsonic - 0.3
For inviscid compressible flow what is Cl a function of
Cl = f(Mach)
For viscid incompressible flow what is Cl a function of
Cl = f(Reynolds)
What is the centre of pressure
The location where the resultant of a distributed load effectively acts on the body
Equation for Centre of Pressure
Xcp = - Mle / n
n - normal force
Mle - Moment
For small incidence:
Xcp = Mle / L
How does centre of pressure move with reducing lift
Xcp moves downstream as lift reduces (can be outside aerofoil)
What is the aerodynamic centre
Fixed point about which the moment is independent of incidence (about 25% of chord length behind leading edge)
How do thickness and viscous effects change the aerodynamic centre
moves aerodynamic centre upstream
How does compressibility effects change the aerodynamic centre
moves aerodynamic centre downstream
Equation for vorticity
Vorticity = ∇ x V(velocity vector)
What does vorticity equal if flowfield is irrotational
vorticity = 0
Define circulation
the negative line integral of the velocity vector around a curce
For an irrotational flowfield, the circulation of a closed curve equals =
zero (contained in the flowfield)
What are the boundary conditions for inviscid flow
for inviscid flow, the flow cannot penetrate the solid surface i.e. velocity vector is tangent to the surface
What is Kelvin’s circulation theorem
circulation around a closed curve consisting of the same fluid elements remains constant as the fluid elements move through the flow
What is Kutta’s condition
Due to viscous effects, the flow has to leave the trailing edge smoothly. This results in circulation and therefore lift
The lift generation per unit span on an aerofoil is given as =
(Kutta-Joukowski theorem)
l = ρVΓ
Γ - circulation
What is the origin of the circulation around the aerofoil
- Before aerofoil moves there is no circulation
- As aerofoil movies a vortex rolls up and around the sharp trailing edge and is flushes downstream
- This starting vortex has a circulation with the same magnitude but opposite sign as that around the aerofoil
What is the thin aerofoil theory
A thin aerofoil can be simulated by a vortex sheet places along the camber line
Equations for a symmetrical thin aerofoil
Cl = 2πα dCl/dα = 2π Xcp = Xac = Xc/4
Equations for a chambered thin aerofoil
Cl = 2π(α-α(l=0)) dCl/dα = 2π Xac = Xc/4
Explain panel methods to solve for the incompressible inviscid flow field
- Aerofoil surface divided into ‘panels’
- Elementary solutions of unknown strength are linearly distributed on these panels
- velocity/boundary conditions give n+1 equations and n+1 unknowns
- Solutions of linear system gives functions for the potential flow fields
- From this, work out velocity vector and then pressure from Bernoulli;s equation
- Deduct forces and moments from surface pressure
What application can the panel method have
used for 3D geometries
Explain digits in NACA-4-series
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
Explain digits in NACA-6-series
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
Describe stall
- 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
As Re increases Cl …
Why
Clmax increases
This is because at high Re, the boundary layer is more resistant to separation
As Camber increases Cl …
Clmax increase
- shifts curve up and left
- stall angle reduces
- increased Cdmax too
As thickness increases Cl …
Why
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
