faculty of engineering and information technology 1.1 mechanics of flight/fundamentals of flight...
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Faculty of Engineering and Information Technology 1.1
Mechanics of Flight/Fundamentals of Flight
COURSE NOTES
John Baird
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Faculty of Engineering and Information Technology 1.2
MILESTONES IN FLIGHT
GEORGE CAYLEY (1773 - 1857): WAS THE FIRST TO EXPLAIN HOW THE FORCE ON THE WIN CAN BE RESOLVED INTO TWO COMPONENTS OF
LIFT PERPENDICULAR TO THE FLIGHTDIRECTION
DRAG PARALLEL TO THE FLIGHTDIRECTION
WING SURFACEDRAG
LIFT
CAYLEY ALSO UNDERSTOOD THE FIRSTPRINCIPLES OF STABILITY AND CONTROL
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Faculty of Engineering and Information Technology 1.3
WE TRY TO ANSWER THE THREE BASIC QUESTIONS
1. WHY AN AEROPLANE FLIES?
AERODYNAMICS
2. WHY DOES IT FLY - SO FAST? SO FAR? SO HIGH?
PERFORMANCE
3. WHY DOES IT BEHAVE THE WAY IT DOES AND HOW TO CONTROL IT?
STABILITY & CONTROL
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Faculty of Engineering and Information Technology 1.4
Visualisation Examples
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Faculty of Engineering and Information Technology 1.5
VISCOSITYTWO FEATURES OF LIQUIDS AND GASES ARE RESPONSIBLE FOR EXISTENCE OFVISCOSITY
LIQUIDS
COHESIVE OR ATTRACTIVE FORCESDOMINATE OVER INERTIA FORCES ANDLARGER THE COHESIVE FORCES (MORECLOSELY PACKED MOLECULES, GREATERTHE VISCOSITY.
GASESTHE BASIS OF VISCOSITY IS THE INTERNALRESISTANCE DUE TO COLLISION ANDTRANSFER OF MOMENTUM.
Fast drift
Slow drift
Fast molecules exchange with slow ones and vice versa. Slowing of fast molecules and vice versa is viscosity.
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Faculty of Engineering and Information Technology 1.6
VISCOSITY WITH TEMP. BECAUSE BONDS BETWEEN MOLECULES WEAKEN OR CAN EVEN BREAK IN OTHER WORDS, COHESIVE FORCES WEAKEN.
VISCOSITY WITH TEMP. BECAUSE INCREASE IN TEMPERATURE CAUSES INCREASED MOLECULAR ACTIVITY WHICH IN TURN LEADS TO MORE COLLISIONS AND MORE TRANSFER OF MOMENTUM, THEREFORE MORE VISCOSITY.
FOR LIQUIDS
DECREASES
FOR GASES
INCREASES
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Faculty of Engineering and Information Technology 1.7
NEWTON’S LAW OF VISCOSITY
THIS ESTABLISHES THE RELATION BETWEEN SHEAR FORCE (FRICTION FORCE) AND THE VISCOSITY.
IT STATES
SHEARING FORCE
A AREA OF INTERFACE
u VELOCITY DIFFERENCE BETWEEN ADJACENT
LAYERS OF FLUID.
h SEPARATION DISTANCE BETWEEN LAYERS OF
FLUID
PROPORTIONALITY CONST. CALLED THE
COEFFICIENT OF VISCOSITY.
Fs
Fs u
h
A
Fs
A
u
h
h
u1
Fs
u = u2- u1
u2
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Faculty of Engineering and Information Technology 1.8
Moving plate
uA
BA’
B’
u
u=0 gradient
Fixed plate
Couette or channel flow
Fixed plate
y
Flow
velocity profile
A B
velocitygradient
h
DEFINITION OF VELOCITY GRADIENT
Boundary layer flow
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Faculty of Engineering and Information Technology 1.9
UNITS OF SHEAR STRESS ARE N/m2 in SI.
FLUIDS OBEYING NEWTON’S LAW, ARE CALLED Newtonian Fluids
EXAMPLES : AIR, WATER, ALL FLUIDS WHICH HAVE SIMPLE MOLECULAR STRUCTURE.
THINK OF SOME FLUIDS WHICH ARE NOT NEWTONIAN!
Fs
A shear stress ' tau'
u
h shear rate
uh
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Faculty of Engineering and Information Technology 1.10
SeaLevel
10
20
30
Altitude(km)
Temperature (degrees Centigrade)-56.5
Troposphere
Tropopause
Stratosphere
Mesosphere Ozone Layer
Balloons
Concorde
Large Jet Liners
General AviationHelicoptersBirdsInsects
U2 Spyplane
International Standard Atmosphere
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Faculty of Engineering and Information Technology 1.11
THE SPEED OF SOUND IS RELATED TO THE PRESSURE & DENSITY BY THE ISENTROPIC RELATION
WHERE IS THE RATIO OF SPECIFIC HEATS IN AIR (CP/ CV) AND IS EQUAL TO 1.4
NOTE THAT THE SPEED OF SOUND IS A FUNCTION OF TEMPERATURE. THE SPEED OF SOUND DECREASES WITH INCREASING ALTITUDE (IE DECREASING TEMPERATURE)
R IS THE GAS CONSTANT FOR AIR AND IS 287.05 J/kg K
SPEED OF SOUND
a p
RT
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Faculty of Engineering and Information Technology 1.12
PROPERTIES OF FLUIDS
FLUIDS CAN BE CLASSIFIED AS OR
.
IN A COMPRESSIBLE FLUID, PRESSURE & VELOCITY CHANGES ARE ACCOMPANIED BY SIGNIFICANT DENSITY CHANGES.
IN AN INCOMPRESSIBLE FLUID, PRESSURE & VELOCITY CHANGES DO NOT CAUSE ANY APPRECIABLE CHANGES IN DENSITY.
LIQUIDS ARE GENERALLY INCOMPRESSIBLE (Eg. WATER).
GASES ARE GENERALLY COMPRESSIBLE (Eg. AIR).
COMPRESSIBLE
INCOMPRESSIBLE
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Faculty of Engineering and Information Technology 1.13
THE FLOW PARAMETER THAT BECOMES IMPORTANT UNDER SUCH CIRCUMSTANCES IS THE MACH NO. DISCUSSED EARLIER.
AT HIGH SPEEDS THEREFORE, WE CAN TALK IN TERMS OF:
SUBSONIC 0.3 < M < 0.7
TRANSONIC 0.7 < M < 1.4
SUPERSONIC 1.4 < M < 5
HYPERSONIC M > 5
PROPERTIES OF FLUIDS
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Faculty of Engineering and Information Technology 1.14
FLOW REGIMES IN AIR
1 2 3 4 5 6 7 8 9 10 0
Su
bso
nic
Tra
nso
nic
Supersonic Hypersonic
TERMINOLOGY
0
Inco
mp
ress
ible
Compressible
No
n L
ine
ar
Linear
1 2 3 4 5 6 7 8 9 10
Oxygen dissociates (Chemical reactions important)
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Faculty of Engineering and Information Technology 1.15
IN VISCOUS FLOWS THE EFFECTS OF VISCOSITY (WHICH PRODUCE FRICTIONAL OR SHEAR STRESSES) ARE CONFINED TO A VERY THIN LAYER OF FLUID CLOSE TO THE SURFACE. THIS THIN LAYER NEAR THE SURFACE IN WHICH VISCOSITY EFFECTS ARE CONFINED IS CALLED BOUNDARY LAYER, BECAUSE OF INTERNAL FRICTION DUE TO VISCOSITY, THE LAYER OF AIR CLOSEST TO THE BODY ‘STICKS’ TO THE SURFACE AND THE VELOCITY GRADUALLY INCREASES TILL AT THE ‘EDGE’ OF THE BOUNDARY LAYER, IT IS EQUAL TO THE VELOCITY IN THE ADJACENT EXTERNAL FLOW.
Fixed plate
y
Flow
velocity profile
A B
velocitygradient
BOUNDARY LAYERBOUNDARY LAYER DETAIL
BOUNDARY LAYERS
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Faculty of Engineering and Information Technology 1.16
BECAUSE THE EFFECTS OF VISCOSITY ARE CONFINED TO THE BOUNDARY LAYER, IT HAS OFTEN BEEN POSSIBLE TO ANALYSE AERODYNAMIC PROBLEMS BY TREATING THE AIR AS IDEAL FLUID AND THIS HAS YIELDED QUITE ACCEPTABLE RESULTS ESPECIALLY AS REGARDS
. FOR TREATING THE PROBLEM OF AERODYNAMIC DRAG,
HOWEVER, WE NEED TO CONSIDER THE
WE ALSO NOTE THAT IN AN IDEAL FLUID, THAT HAS NO VISCOSITY, THE FLUID EXHIBITS NO SHEAR FORCES AND THERE WOULD BE NO RELATIVE MOTION BETWEEN ADJACENT LAYERS OF FLUID. SUCH A FLUID IS SAID TO
PAST THE SURFACE.
AERODYNAMIC LIFT
BOUNDARY LAYER
SLIP
BOUNDARY LAYERS
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Faculty of Engineering and Information Technology 1.17
WITH REAL FLUIDS, ON THE OTHER HAND, THERE IS NO RELATIVE MOTION AT THE SURFACE. IN OTHER WORDS, AT THE SURFACE WE HAVE CONDITION.
THE CONCEPT OF THE BOUNDARY LAYER WAS FIRST PROPOSED BY THE GERMAN AERODYNAMICIST LUDWIG PRANDTL (1875 - 1953) IN 1905.
IT WAS TRULY A MILESTONE IN AERODYNAMICS
NO SLIP
BOUNDARY LAYERS
y
VelocityZero velocity at wall, the no slip condition
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Faculty of Engineering and Information Technology 1.18
TYPICAL BOUNDARY LAYER PROFILE
Free stream velocity, U
0.99U
Velocity gradient
Velocity Profile
Surface
Boundary layer thickness
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Faculty of Engineering and Information Technology 1.19
WHEN THE BOUNDARY LAYER IS SUBJECTED TO
INCREASING PRESSURE IN THE FLOW DIRECTION, IT
BECOMES MORE AND MORE SLUGGISH AS IT HAS TO FLOW
AGAINST AN ADVERSE PRESSURE GRADIENT. IT
EVENTUALLY COMES OFF THE SURFACE. THE BOUNDARY
LAYER IS THEN SAID TO BE SEPARATED.
WHEN THE BOUNDARY LAYER ON AN AEROFOIL OR WING
SEPARATES, WE SAY THAT THE AEROFOIL OR WING HAS
ONCE THE BOUNDARY LAYER SEPARATES THE DRAG
INCREASES DRAMATICALLY AND WE CAN NO LONGER
ASSUME THAT FLOW OVER. THE AEROFOIL IS SMOOTH
(IDEAL).
STALLED
BOUNDARY LAYER BEHAVIOUR
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Faculty of Engineering and Information Technology 1.20
A Separated Boundary Layer
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Faculty of Engineering and Information Technology 1.21
B.L. BEHAVIOUR IN ADVERSE PRESSURE GRADIENT
Pressure force
Flow decelerating
Increasing Pressure
Point of separation Flow separated
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Faculty of Engineering and Information Technology 1.22
PRESSURE DISTRIBUTION ON AN AEROFOIL
Negative pressure
Adverse Pressure Gradient
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Faculty of Engineering and Information Technology 1.23
STALL ON AN AEROFOIL
Negative Pressure
NOTE RECIRCULATION REGION
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Faculty of Engineering and Information Technology 1.24
DRAG ARISES DUE TO
SKIN FRICTION ON SURFACE OF BODY
PRESSURE DISTRIBUTION OVER BODY
SKIN FRICTION DRAG
PRESSURE DRAG
NATURE OF DRAG
(ALSO CALLED FORM DRAG)
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Faculty of Engineering and Information Technology 1.25
REAL FLOWS AND AERODYNAMIC DRAG
VV
INCISCID FLOW VISCOUS FLOW
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Faculty of Engineering and Information Technology 1.26
Pressure Drag on a Cylinder
degrees
p p01
2U2
-3.5
-3.0
-2.5
-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
0 30 60 90 120 150 180 210 240 270 300 330 360
Theoretical (potential flow)
Measured (R = 670000)
Measured (R=190000)
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Faculty of Engineering and Information Technology 1.27
Skin Friction Drag
yU
y
wall
Dskin friction Awall AU
y
wall
Skin Friction Drag= Area x Shear Stress at wall
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Faculty of Engineering and Information Technology 1.28
Total Drag
DRAG ON A TWO DIMENSIONAL OBJECT (PROFILE DRAG) IS A COMBINATION OF PRESSURE DRAG (ALSO CALLED FORM DRAG) AND SKIN FRICTION DRAG
PROFILE DRAG = PRESSURE DRAG + SKIN FRICTION DRAG
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Faculty of Engineering and Information Technology 1.29
VELOCITY IS INVERSELY PROPORTIONAL TO THE CROSS-
SECTIONAL AREA. IN OTHER WORDS,
WHENEVER THERE IS ACCELERATION OF FLUID FLOW, THE
CROSS-SECTION IS NARROW AND THE STREAMLINES
CONVERGE.
WHEN THERE IS A DECELERATION, THE CROSS-SECTION IS
WIDER AND THE STREAMLINES DIVERGE.
Stream Tube
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Faculty of Engineering and Information Technology 1.30
BERNOULLI’S EQN. RELATES CHANGES IN VELOCITY TO
CHANGES IN PRESSURE IN STEADY INCOMPRESSIBLE
INVISCID FLOW ALONG A STREAM LINE. IT IS THE MOST IMPORTANT EQUATION IN FLUID MECHANICS.
p 1
2V2 p0 constant
where p is the (static) pressure is the density, V is the local velocity and p0, the constant, is called the total pressure. The term is sometimes called the dynamic pressure.
1
2V2
Bernoulli’s Equation
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Faculty of Engineering and Information Technology 1.31
Applications of Bernoulli’s Equation
One way of writing Bernoulli’s equation is:Total pressure = static pressure + dynamic pressure.
po p 1
2V2
If the velocity is zero the pressure is equal to the total pressure
p
To differential pressure gauged
4d
A PITOT TUBE
po p 1
2V2
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Faculty of Engineering and Information Technology 1.32
FROM BERNOULLI’S EQN. WE HAVE
p 1
2V2 p0
or
V 2 p0 p
THEREFORE, BY MEASURING THE DIFFERENCE BETWEEN
THE TOTAL PRESSURE (Po) AND STATIC PRESSURE (), WE
CAN CALCULATE THE VELOCITY IN A FLOW.
Measurement of Velocity:
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Faculty of Engineering and Information Technology 1.33
PITOT TUBE USED IN A WIND TUNNEL TO MEASURE VELOCITY
po
V
To micromanometer
Pitot-static tube
Test Section
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Faculty of Engineering and Information Technology 1.34
PITOT TUBE USED TO MEASURE BOUNDARY LAYER PROFILE
V
Boundary layer velocity
profile
Staticorifice
To manometer
Pitot tube fixed to manometer traverse
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Faculty of Engineering and Information Technology 1.35
po p 1
2V2
Total energy per unit volume
pressure energy per unit volume
kinetic energy per unit volume
ALTERNATIVE VIEW OF BERNOULLI’S EQUATION
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Faculty of Engineering and Information Technology 1.36
Pressure and Velocity
Negative pressure
How Lift is Generated
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Faculty of Engineering and Information Technology 1.37
THEREFORE, BY NOTING THE EQUIVALENT AIRSPEED FROMASI AT ANY ALTITUDE, WE CAN DETERMINE THE TRUE AIRSPEED BY THE KNOWLEDGE OF RELATIVE DENSITY .
USUALLY, THE AIRSPEED INDICATOR IS CALIBRATED SOTHAT IT READS DIRECTLY THE SPEED EITHER IN KNOTS ORkm/hr.
Static pressure measured by port on fuselage
Using Bernoulli’s Equation to Measure Velocity of Aircraft
Dynamic pressure measured by pitot probe
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Faculty of Engineering and Information Technology 1.38
Flow Examples
Flow accelerates and pressure reduced
Dividing streamlinep=p0 at stagnation point
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Faculty of Engineering and Information Technology 1.39
Flow Examples
All objects in a flow have a stagnation point
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Faculty of Engineering and Information Technology 1.40
Flow Examples
Distance along aerofoil
Pressureand
Velocityp
V
Free streamVp
s s’
s’s
p
V
Vp
T
FF’
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Faculty of Engineering and Information Technology 1.41
TRANSITION TO TURBULENCE
Smooth entry
(a) Laminar Flow
Dye filament Control valve
dWater
Filament becomes unstable
Filament breakup and turbulent flow
(b) Turbulent Flow
V
V
Osborne Reynolds in the 1880s investigated the behaviour of flow that was either direct (laminar) or sinuous (turbulent).
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Faculty of Engineering and Information Technology 1.42
Transition to Turbulence
Whether the flow is turbulent or laminar depends on the relative magnitude of the viscous forces and the kinetic forces (momentum) of the flow.
When viscous forces are large, small irregularities are removed by ‘viscous damping’. This is characterised by slow flow, and/or high viscosity.
When the flow has sufficient momentum such that the viscous forces are relatively small, it becomes turbulent.
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Faculty of Engineering and Information Technology 1.43
Example of a Turbulent Boundary Layer
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Faculty of Engineering and Information Technology 1.44
Stability of Shear Flow
Splitter Plate Dividing stream line
Consider the flow above in which viscosity is very small. Consider what happens if the dividing stream line is disturbed a small amount.
Note that, if the velocity difference is high enough the pressure differences will act to increase the divergence of the streamline.
IT WILL BECOME TURBULENT
p,V1
p,V2V P
V P
V P
V P
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Faculty of Engineering and Information Technology 1.45
Stability of Shear Flow
p,U1
p,U2U P
U P
U P
U P
IF VISCOUS FORCES DOMINATE, DISTURBANCE ENERGY WILL DISSIPATE. IF KINETIC FORCES DOMINATE THE DISTURBANCES WILL GROW AND FLOW WILL BECOME TURBULENCE. WHETHER A FLOW IS TURBULENT OR NOT DEPENDS ON:
KineticForces
Viscous Forces
1
2U2
dUdy
U2
U
D
UD
D
= REYNOLDS NUMBER
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REYNOLDS NUMBER
= FLUID DENSITYV = A VELOCITY - USUALLY THE FREE STREAM VELOCITYD = A REPRESENTATIVE LENGTH SCALE = THE FLUID VISCOSITY
R UD
EXAMPLES
U
D
R UC
CHORD C
For a given configuration and definition the Re determines when the transition to turbulence occurs.
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LAMINAR AND TURBULENT FLOW
LAMINAR FLOW
SMOOTH STEADY SMALLER SHEAR STRESS
TYPICALLY STREAMLINEDBODIES
LESS SKIN FRICTION DRAG. MORE PRESSURE DRAGLaminar profile
Turbulent mean profile
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TURBULENT FLOW
HIGHLY DISORGANISED BASICALLY LARGEUNSTEADY SHEAR STRESS
MORE SKIN FRICTION DRAG.LESS PRESSURE DRAG
TYPICALLYBLUFF BODIES
Laminar profile
Turbulent mean profile
LAMINAR AND TURBULENT FLOW
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Transition to Turbulence
Smoke visualisation of a boundary layer. The laminar boundary layer on the left is ‘tripped’ by a grid and becomes turbulent
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Boundary Layer Profiles
Y
U
Laminar Profile
Streamlines
Momentum exchange by viscous forces only
Turbulent Profile
Momentum exchanged more effectively by mass transport into lower layers
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(a)
(b)
The laminar boundary layer (a) encourages separation and leads to a wide wake and high form drag. In (b) a trip is used to cause transition to turbulence in the boundary layer, separation is delayed, the wake narrows and the form drag decreases.
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Turbulent Boundary Layer on a Flat Plate
Free stream velocity V
Boundary layeredge
L
Vertical dimension exaggerated
t
s
tL
Laminar Transition Turbulent
turbulent eddies
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Four Fluid Phenomena: Number 1Pressure force
Flow decelerating
Point of separation Flow separated
1. Adverse pressure gradients cause separation
(Separation cannot occur in favourable gradients)
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Four Fluid Phenomena: Number 2
Point of separation Flow separated
2. Turbulence inhibits or delays separation
Pressure force
Laminar flow
Energy deep in boundary layer is resistant to separation
Turbulent flow
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Four Fluid Phenomena: Number 3
3. In a certain Re range adverse pressure gradients can encourage turbulence, and favourable gradients can relaminarise flow.
Pressure force
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Four Fluid Phenomena: Number 4
4. In a certain Re range surface roughness encourages turbulence.
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Cricket Ball Swing
Laminar Flow
Turbulent Flowtripped by seam and maintained by rough surface
Roughened side
Polished Side
Asymmetric wake
Angle (degrees)
Pressure
0 180
1/2 V2
TurbulentLaminar
LIFT
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REYNOLDS NUMBER & ITS SIGNIFICANCE
WE HAVE SEEN THAT THE TURBULENT MOTION IS MORE
VIGOROUS AND ENERGETIC. ALSO, THE VELOCITY
FLUCTUATIONS IN TURBULENT FLOW IMPOSE STRESSES
ADDITIONAL TO THOSE SHEAR STRESSES THAT RESULT
FROM MOLECULAR MOTIONS. THESE ADDITIONAL
STRESSES ARISING PURELY OUT OF TURBULENT
FLUCTUATIONS ARE CALLED OR
SOMETIMES REFERRED TO AS EDDY STRESSES.
THEREFORE:
TOTAL SHEAR STRESS VISCOUS SHEAR STRESS
IN TURBULENT FLOW
REYNOLDS STRESSES
REYNOLDS STRESSES
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More on Reynolds Number and Scaling
IN FLUID MECHANICS IT IS MORE CONVENIENT TO USE NON-DIMENSIONAL NUMBERS SUCH AS THE REYNOLDS NUMBER IN DESCRIBING FLOW. LET US FIND A NON-DIMENSIONAL NUMBER RELATED TO DRAG.
DRAG IS A FORCE WHICH, IN MANY CASES, IS RELATED TO THE KINETIC ENERGY OF THE FLOW (SEE BELOW)
Gauge pressure of 1/2 U2
Gauge pressure near zero
IN THE CASE SHOWN, THE DRAG IS LIKELY TO BE PROPORTIONAL TO THE DYNAMIC PRESSURE 1/2 U2
THEREFORE WE DEFINE THE DRAG COEFFICIENT AS
CD D
12U2A
DRAG
U
Area A
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More on Reynolds Number and Scaling
SIMILARLY THE LIFT COEFFICIENT IS
CL L
1
2U2A
WHERE L IS THE LIFT GENERATED BY AN AEROFOIL OF PLAN AREA A
Wing Planform (A)
Air Flow
(V)
Chord
(c)
Span (b)
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WHY USE COEFFICIENTS ?IMAGINE THAT YOU HAVE MEASURED THE LIFT ON AN AEROFOIL IN A WIND TUNNEL AT VARIOUS VELOCITIES
ANGLE OF ATTACK
LIF
T
ANGLE OF ATTACKLI
FT
NOW WE VARY THE SIZE OF THE AEROFOIL
U1
U2
U4
U3
C1
C4
C3
C4
THE RESULTS ARE FAR TOO COMPLEX
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WHY USE COEFFICIENTS ?
ANGLE OF ATTACK
LIF
T C
OE
FF
ICIE
NT
IF WE PLOT NON-DIMENSIONAL NUMBERS SUCH AS LIFT COEFFICIENT VS ANGLE OF ATTACK ALMOST ALL DATA COLLAPSES TO ONE LINE
THE RESULTS ARE NOW SIMPLIFIED
SLIGHT VARIATION CAUSED BY CHANGING REYNOLDS NUMBER
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ANOTHER EXAMPLEDRAG ON A CYLINDER
VELOCITY
DR
AG
VARIABLE DIAMETERS
Log
(DR
AG
CO
EF
FIC
IEN
T)
Log (REYNOLDS NUMBER)
ON A NON-DIMENSIONAL PLOT OF APPROPRIATE PARAMETERS ALL DATA COLLAPSES TO ONE GRAPH
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SCALING LAWS & DIMENSIONAL ANALYSIS
IN FLUID FLOWS, WE CAN ALWAYS IDENTIFY A
CHARACTERISTIC DIMENSION FOR A BODY OVER OR
THROUGH WHICH FLUID FLOWS. THIS CAN BE, FOR
EXAMPLE, THE CHORD OF AN AEROFOIL, DIAMETER OF A
PIPE, HEIGHT OF A CHANNEL, ETC.
IN A SIMILAR VEIN, WE CAN DEFINE A CHARACTERISTIC
VELOCITY SCALE WHICH IS TYPICAL FOR A GIVEN PROBLEM.
FOR EXAMPLE, THE FLIGHT SPEED OF AN AIRCRAFT, OR
TEST SECTION VELOCITY IN A WIND TUNNEL, OR FREE-
STREAM VELOCITY EXTERNAL TO THE BOUNDARY LAYER.
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REYNOLDS NUMBER
UL
REYNOLDS NUMBER IS THE MOST IMPORTANT PARAMETER IN AERODYNAMICS.
is the density of the fluid
U is a characteristic velocity, typically free stream velocity
L is a characteristic dimension, typically the diameter of a cylinder or pipe or the chord of an aerofoil
is the fluid viscosity.
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Drag & Lift CoefficientsThe main aim of aerodynamics is to predict drag and lift.
Drag has the units of force mass x acceration ML
T2
The drag is likely to be related to the pressure experienced by the aerofoil which is typified by the Stagnation Pressure 1/2 U2
1
2U2 has units of
M
L3
L2
T2
M
LT2
1
2U2A has units of
M
L3
L2
T2
L2
ML
T2
We can multiply by an area ( L2) to get the same units as Force
ThereforeD
1
2U2A
is dimensionless and is called the Drag Coefficient
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DRAG ON A CYLINDER
Re < 1 Viscous flow, drag is proportional to velocity, skin friction drag dominates
10 < Re < 5x105 Viscous flow, drag is proportional to the square of velocity, pressure drag dominates.
Re > 5x105 Boundary layer is turbulent and separation is delayed. Thus the wake is narrower and there is a greater area of pressure recovery on the rearward surfaces. Thus drag is reduced.
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0.01
0.1
1
10
100
1000
1 2 3 4 5 6 7 8 9
Series1
log10(Reynolds number)
Spe
ed (
m/s
)
dust particles
insects
models
birds & bats
human-powered aircraft
hang gliders
sailplanes
general aviationjet transports
dirigibles
0.001
0.0001
0.01
0.050.10.2
0.51.0
Mac
h nu
mbe
r
Approximate Reynolds number ranges of aerodynamic objects in nature and technology.
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0.05 90.05 5
10000 95 9
5
5.5
6
6.5
7
7.5
8
8.5
9
9.5
0.01 0.1 1 10
Take-offCruise
Vstol, cargo.
Large Jet transport
L= 4 m
Concorde L = 20 m
US SST L = 32 m
Hypersonic Aircraft L = 32m
ICBM
IRBM
Gemini GT-3L = 2.1 m
Space FerryL = 10.7 m
Velocity (km/s)
RL
Low Speed Compressible
Transonic
Supersonic Hypersonic
M - 0.3 0.8 1.2 5
Subsonic
Reynolds number and speed (Mach number) regimes for various vehicles (Poisson-Quinton, 1968)
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How does Streamlining work?
Re<1
A B
Consider two bodies of revolution A and B. At low Re, skin friction drag is much larger than pressure drag. Viscous forces dominate. Note that B has a much larger surface area than A.
The drag on B will be much larger than the drag on A. Streamlining will not work.
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Re>100
A
B
Pressure drag dominates.
The drag on A will be much larger than the drag on B. Streamlining reduces the drag
substantially
How does Streamlining work?
Angle (degrees)
Pre
ssur
e
0 180
1/2 V2
Distance from leading edge
Pre
ssur
e
1/2 V2
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The Effect of Streamlining
NACA 64-421 airfoil compared with a circular wire having the same drag. The diameter of the wire is one tenth of the thickness of the aerofoil.
Thickness
Wire
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MORE ON REYNOLDS NUMBER
If we want to model the flow over an object we should try to get the Reynolds number of the model and the full size object as close as possible.
Re mod el Re full size
UL
mod el
UL
full size
if model full size and model full size , then
Umodel L full size
LmodelUfull size
Lmodel Lfull size
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MORE ON REYNOLDS NUMBER
IF (Re )model= (Re )full size
Then (Cd )model= (Cd )full size
and (CL )model= (CL )full size
IN PRACTICE THIS CAN RARELY BE ACHIEVED
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V
W
L AF
DT
Primary forces acting on an aircraft in steady level flight.
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V
L
D
L/D=2
V
L
D
L/D=3.3
V
L
D
L/D=4
Lift to drag ratio for a flat plate, a cambered plate and an aerofoil at incidence.
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0
2
4
6
8
10
12
14
16
18
-2 0 2 4 6 8 10 12
Wright Brothers (1903)
A modern Aerofoil
Angle of Attack (degrees)
Lif
t to
Dra
g R
atio
Variation of Lift to Drag ratios.
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Wing Planform (S)
Span (b)
Air Flow
(V)
Leading edge
Trainling edge
Chord
(c)
Thickness
Camber
Camber line
Trailing edge
Leading edge radius
Air Flow
Chord line
AoA()
Chord (c)
A
A
Section A-AWing and aerofoil nomenclature
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Sweep angle (
b
c
Rectangular wing Swept wing
c
Root cord (cr)
Elliptic wing
Tapered wing
Tip cord(ct)
b
cr
cr
Delta wing
Ogive wing(Concorde)
Figure 8.4a Typical wing planforms
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Conventional low speed aerofoil
Low speed symmetric aerofoil
Laminar flow aerofoil
Transoinc ‘supercritical’ aerofoil
Thin supersonic ‘biconvex’ aerofoil
Multi-element ‘high lift’ aerofoil
Typical aerofoil sections
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-2
-1.5
-1
-0.5
0
0.5
1
0.0 0.2 0.4 0.6 0.8 1.0x/c
Cp
Pressure loss due to viscosity
Pressure distribution on a lifting aerofoil
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0
0.2
0.4
0.6
0.8
1
1.2
1.4
-5 0 5 10 15 20
A
B
C
Angle of Attack ( degrees)
Lif
t Coe
ffic
ient
(C
L)
Maximum lift coefficient
Moderate angle of attack
Lift curve slope(ideal value 2)
Linear or operationalsection of the lift curve
Lift curve showing changes in pressure distribution and the flow around the aerofoil
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Ways of Plotting Aerofoil Performance
Angle of Attack
Lift
Coe
ffic
ient
Angle of AttackD
rag
Coe
ffic
ient
or
Drag Coefficient
Lift
Coe
ffic
ient
Polar Plot
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Lift and Drag Curves
Angle of Attack
Lift
Coe
ffic
ient
Angle of AttackD
rag
Coe
ffic
ient
CLmax
dCL
d
CDmin
Shape of curve indicates sudden stall or gentle stall
Offset indicates asymmetric aerofoil
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Polar Plots
Drag Coefficient
Lift
Coe
ffic
ient
dCL
dCD
max
CLmax
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Measurement of Profile Drag
Up till now we have been talking about the behaviour of infinitely long aerofoil sections (or ‘profiles’). These are measured in an wind tunnel with the aeorfoils being terminated in the walls. Real wings are not infinite
Testing Aerofoil ProfilesTesting three dimensional wings
Wing Tip Vortex
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Induced Drag
Wing Tip
Low Pressure
High Pressure
Wing Tip Vortex
Note that a downwards velocity is generated
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Induced Drag
Airspeed
Lift
Drag
Two Dimensional Flow(infinite aspect ratio)
Local Airspeed
Local Lift
Drag
Three Dimensional Flow(finite aspect ratio)
Induced downwards airflow
Induced Drag
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Induced DragInduced Drag is proportional to the square of the Lift Coefficient.
Di12U2S
CL
2
ARwhere
Di
S AR
the Induced Drag
the wing area
the aspect ratio, span divided by chord
CL
Infinite wing
Finite wing
Note, curves cross at zero lift point
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Basic Flight Mechanics
V
W
L AF
DT
Note that for steady level flight, the Lift L is equal to the Weight W (=mg). If we assume that the lift is generated entirely by the wing, we can write
CL L
1
2U2S
mg
1
2U2S
or
U mg12S
.1
CL
To maintain steady level flight at low speeds the CLmust be increased (by increasing angle of attack). In high speed flight CL must be decreased (by decreasing angle of attack).
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Basic Flight Mechanics
L
D
T
W
Angle of attack Small ( 3-4 degrees)
W
L
D
Angle of attack Large ( 20-25 degrees)
Umin mg12S
.1
CL max
Note: an aircraft’s minimum speed (landing speed) is determined by CLmax and the wing area.
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Total Drag on an AircraftThe total drag on an aircraft is a combination of drag caused by appendages, skin friction drag, pressure drag (collectively called parasite drag) and induced drag. Induced drag is high at low speeds (high CL) and low at high speeds ( low CL).
Drag
Velocity (steady and level)
CD CD,e CL
2
eARCD
CL2
eAR
CD,e
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Total Drag on an Aircraft
CD CD,e CL
2
eAR
Total Drag
Parasite Drag: Includes profile drag, (pressure drag and skin friction drag on wing), skin friction and pressure drag on fuselage, empennage, engine necelles, landing gear etc.
Induced Drag
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High Lift DevicesUmin
mg12S
.1
CL max
The minimum controllable speed at which an aircraft can fly determines the landing speed. From the above equation, the landing speed can be reduced by increasing CLmax ( ie by delaying stall) or by increasing the wing area. Most modern high lift devices work on a combination of:• Delaying stall by increasing camber,• Delaying stall by re-energising the boundary layers inhibiting separation,• Increasing the total effective wing area.
Chord Increased
Boundary Layers re-energised
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The Effect of Flaps
CL
=0
=30
=15
CLmax increases with
Zero lift point changes with camber
Modern aerofoils have a CLmax of 1.4. Multi-element flaps can increase that to 3.2
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Elementary Flight Mechanics
V
1.
S21W
bV.S2
1aDVP
V
1.
S21W
bV.S2
1aD
drag induced drag parasite bCaC
SV21
DC
23
req
2
22
2LD
2D
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Power and Drag
Velocity
Stall SpeedP
D
Vmp VmD
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As Weight Increases
Velocity
Stall Speed
D
VmD
DragWeight W1
Weight W2
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Increasing Weight
• W=L and as W increases L increases
• If CL and the AoA are unchanged V must increase as W1/2
• At a given AoA, L/D is constant and D is proportional to W
• At a given AoA, Preq =DV and is proportional to W3/2
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Power Requirements
Velocity
Stall Speed
Preq
Power
Pavail
Max climb rate Max speed
Take-off and Landing speed
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Maximum Payload
Velocity
Preq
Drag
Pavail
Minimum workable manoevre range
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Pitching Moment
Pitching Moments: Cm=M/(1/2 V2Sc)=Cm0+ k CL
(Nose up is positive)
CL (or AoA)Cm
Aft point of reference
Forward point of reference
Aerodynamic Centre ref point
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Aerodynamic Centre
O A C
L
M
O A
L
M0= constant
O A C
L
C
Centre of Pressure
Position of C varies with
Position of Aerodynamic Centre(A) remains fixed
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Stability of Wing Alone
CG moving aft
Highly unstable
Unstable
Neutrally stable
Stable (desirable)
Unresponsive
Pitching Moment
CL
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Longitudinal StabilityLw
LT
WC
G
Lw
LT
W
G
M0= constant
A
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Stability of Wing and Tailplane
Wing Alone
Pitching Moment
Tail Alone
Aircraft
CL
Trim Point
• Movement aft of CofG makes aircraft less stable.• Neutral point is point of CofG where aircraft is neutrally stable (Cm constant)• Static margin is distance CofG is ahead of neutral point