Subsonic Aerodynamics Flashcards
What are the conventions for questions?
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.
List SI units.
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
State and interpret Newtons laws
1st - Same speed if no resultant force
2nd - F=ma, so acceleration is proportional to resultant force
3rd - Equal but opposite reaction force
Explain air density
Mass per unit volume
List atmospheric properties that affect density
Temperature and pressure, gas constant
P=RHORT
Define static Pressure
Pressure at any point in fluid not associated with its motion
Define dynamic pressure, and its formula
Pressure attributed to a fluid’s movement
q = 1/2 * RHO * V^2
Apply formula for given altitude and speed
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
State Bernoulli’s equation
Conservation of energy in closed system.
At throat, Velocity increases so pressure decreases, dynamic pressure increases
Remember converging and diverging nozzle
Define total pressure. Apply it to a venturi
P+q=P0
AKA stagnation pressure.
In venturi, along streamlines total pressure is constant. Static and dynamic vary.
Describe how IAS is gotten from Pitot-static tube
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
Describe relationship between P,T,RHO for air
Explain continuity
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
Define IAS,CAS,EAS,TAS
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
Describe steady and unsteady airflow
Steady - straight streamline. smooth, predictable. Takes path of least resistance
Unsteady - broken streamline
Steady - P,T and RHO are independent of time
Explain concept of streamline
The path a bit of fluid takes. Has constant total pressure
Describe airflow through a stream tube
A streamtube defines a volume of air. No flow in and out of tube
Describe force resulting from pressure distribution round an aerofoil
High pressure below, low pressure on top, so resultant force upwards
Resolve resultant force into lift and drag
Lift = Perpendicular to free stream flow
Drag - Parallel to free-stream-flow
Describe all parameters of a wing
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
Explain difference of 2D and 3D wing
Spanwise flow.
Both are ideal fluid and we will consider viscosity
Explain streamline pattern around aerofoil
Top - close together, fast, less static pressure
Bottom - further apart/same as free stream. Higher static pressure
Describe stagnation point
V = 0. Static pressure - Total pressure. So high pressure
Explain the effect of AOA changes on stagnation point
Will change its position accordingly…
Explain local pressure changes with camber and AOA
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.
Describe effect of AOA increase and camber increase on aerofoil
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
Explain the centre of pressure
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.
Explain aerodynamic centre
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.
List two physicsal phenomena that cause drag.
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
Why do drag and wake cause loss of momentum?
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.
Explain influence of AOA on lift. Symmetrical and Cambered
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
Why are coefficients used in general?
They make the equations fit reality. Cl describes the natural ability to produce lift.
Describe the lift formula and perform simple calcs
Lift = 1/2 * RHO * S * V^2
Compare high and low speed aerofoils
Low speed - high camber/ high thickness ratio for max lift
High speed - low camber/ low thickness ratio for max lift
Describe drag formula
Same as lift except it has a coefficient of drag
Discuss effect of shape of body on drag
More frontal area - more drag
Define the following terms
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
Explain causes of spanwise flow
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
Downwash
Heavy near tips due to vortices
Boundary layer
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
What are 3 points of flow on win. What happens when change AOA
Stagnation point moves down
Transition point moves forward
Separation point moves forward
Due to adverse pressure gradient. Bad question no. 15544
Explain 2D drag
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!!
