This document provides a table of contents for a document on aerodynamics. It discusses various topics related to aerodynamics including mathematical notations, basic laws of fluid dynamics, boundary conditions, airfoil design methods, compressible flow, shock waves, and linearized flow equations. Specifically, it summarizes the conical flow method and singularity distribution method for obtaining the theoretical solution for pressure distribution on a finite span wing in supersonic flow. The conical flow method assumes the potential and other flow properties are constant along rays through a common vertex, modeling conical flow patterns seen in supersonic flows.
2. 2
Table of Content
AERODYNAMICS
Earth Atmosphere
Mathematical Notations
SOLO
Basic Laws in Fluid Dynamics
Conservation of Mass (C.M.)
Conservation of Linear Momentum (C.L.M.)
Conservation of Moment-of-Momentum (C.M.M.)
The First Law of Thermodynamics
The Second Law of Thermodynamics and Entropy Production
Constitutive Relations for Gases
Newtonian Fluid Definitions – Navier–Stokes Equations
State Equation
Thermally Perfect Gas and Calorically Perfect Gas
Boundary Conditions
Flow Description
Streamlines, Streaklines, and Pathlines
A
E
R
O
D
Y
N
A
M
I
C
S
P
A
R
T
I
3. 3
Table of Content (continue – 1)
AERODYNAMICS
SOLO
Circulation
Biot-Savart Formula
Helmholtz Vortex Theorems
2-D Inviscid Incompressible Flow
Stream Function ψ, Velocity Potential φ in 2-D Incompressible
Irrotational Flow
Aerodynamic Forces and Moments
Blasius Theorem
Kutta Condition
Kutta-Joukovsky Theorem
Joukovsky Airfoils
Theodorsen Airfoil Design Method
Profile Theory by the Method of Singularities
Airfoil Design
A
E
R
O
D
Y
N
A
M
I
C
S
P
A
R
T
I
4. 4
Table of Content (continue – 2)
AERODYNAMICS
SOLO
Lifting-Line Theory
Subsonic Incompressible Flow (ρ∞ = const.) about Wings
of Finite Span (AR < ∞)
3D Lifting-Surface Theory through Vortex Lattice Method (VLM)
Incompressible Potential Flow Using Panel Methods
Dimensionless Equations
Boundary Layer and Reynolds Number
Wing Configurations
Wing Parameters
References
A
E
R
O
D
Y
N
A
M
I
C
S
P
A
R
T
I
5. 5
Table of Content (continue – 3)
AERODYNAMICS
SOLO
Shock & Expansion Waves
Shock Wave Definition
Normal Shock Wave
Oblique Shock Wave
Prandtl-Meyer Expansion Waves
Movement of Shocks with Increasing Mach Number
Drag Variation with Mach Number
Swept Wings Drag Variation
Variation of Aerodynamic Efficiency with Mach Number
Analytic Theory and CFD
Transonic Area Rule
A
E
R
O
D
Y
N
A
M
I
C
S
P
A
R
T
I
I
6. 6
Table of Content (continue – 4)
AERODYNAMICS
SOLO
Linearized Flow Equations
Cylindrical Coordinates
Small Perturbation Flow
Applications: Nonsteady One-Dimensional Flow
Applications: Two Dimensional Flow
Applications: Three Dimensional Flow (Thin Airfoil at Small Angles of Attack)
Applications: Two Dimensional Flow (Thin Airfoil at Small Angles of Attack)
Drag (d) and Lift (l) per Unit Span Computations for Subsonic Flow (M∞ < 1)
Prandtl-Glauert Compressibility Correction
Computations for Supersonic Flow (M∞ >1)
Ackeret Compressibility Correction
A
E
R
O
D
Y
N
A
M
I
C
S
P
A
R
T
I
I
7. 7
SOLO
Table of Contents (continue – 5)
AERODYNAMICS
Wings of Finite Span at Supersonic Incident Flow
Theoretic Solutions for Pressure Distribution on a
Finite Span Wing in a Supersonic Flow (M∞ > 1)
1. Conical Flow Method
2. Singularity-Distribution Method
Theoretical Solutions for Compressible Supersonic Flow (M∞ >1)
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing
in a Supersonic Flow and Subsonic Leading Edge (tanΛ > β)
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing
in a Supersonic Flow and Supersonic Leading Edge (tanΛ < β)
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β)
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β)
Arrowhead Wings with Double-Wedge Profile at Zero Incidence
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having
Straight Leading and Trailing Edges and the same dimensionless profile in
all chordwise plane [after Lawrence]
8. 8
Table of Content (continue – 6)
AERODYNAMICS
SOLO
Aircraft Flight Control
References
CNα – Slope of the Normal Force Coefficient Computations of Swept Wings
Comparison of Experiment and Theory for Lift-Curve Slope of Un-swept Wings
Drag Coefficient
10. 10
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
The essential physical difference between Subsonic and Supersonic Flow is:
- Subsonic Flow: The disturbances of a sound point source propagates in all
directions.
- Supersonic Flow: The disturbance of a sound point propagates only within a
cone that lies downstream of the sound source. This so-called Mach-Cone has
the apex semi-angle μ
Supersonic
V > a
a t
V t
M
1
sin 1
Sound
waves
Mach
waves
1
1
tan
1
sin
1/:
2
MM
aVM
11. 11
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Mach Cone
Wing
Leading Edge
Mach Cone
Wing
Leading Edge
Mach Cone
Wing
Leading Edge
Mach Cone
Wing
Leading Edge
If the Mach Line lies before
the Wing Edge, the component vn
of the incident Flow Velocity U∞
normal to the Wing Edge is
smaller than the Speed of Sound
a∞. Such a Wing Edge is called
Subsonic.
Conversely, if the Mach Line
lies behind the Wing Edge, the
component vn of the incident
Flow Velocity U∞ normal to the
Wing Edge is larger than the
Speed of Sound a∞. Such a Wing
Edge is called Supersonic.
Subsonic Edge vn<a∞ μ>γ m<1
Supersonic Edge vn>a∞ μ<γ m>1
12. 12
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Mach Line
Wing
Leading
Edge
Mach :Line
Wing
Trailing
Edge
Mach LineWing
Leading
Edge
Mach :Line
Wing
Trailing
Edge
Mach Line
Wing
Leading
Edge
Mach :Line
Wing
Trailing
Edge
Subsonic Leading Edge
Subsonic Trailing Edge
Subsonic Leading Edge
Supersonic Trailing Edge
Supersonic Leading Edge
Supersonic Trailing Edge
Subsonic Leading
Edge Flow
Subsonic Trailing
Edge Flow
Supersonic Leading
Edge Flow
Supersonic Trailing
Edge Flow
13. 13
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Mach Line
Wing
Leading Edge
Mach :Line
Influence
Range of A
Wing
Trailing Edge
Consider a point A’ (x,y,z) on a Wing in a
Supersonic Flow (V∞/a∞ > 1). The points
on the Wing that, by perturbing the Flow,
influence the Flow properties at A’ are
only downstream to A’, bounded by the
Wing Leading Edges and the Mach Lines
(ML) passing through A’ (see Figure).
Mach Line
Wing
Leading Edge
Mach :Line
Influence
Range of A
Wing
Trailing Edge
Subsonic Leading Edge
Supersonic Leading Edge
Return to Table of Content
14. 14
SOLO Wings in Compressible Flow
Theoretic Solutions for Pressure Distribution on a Finite Span Wing
in a Supersonic Flow (M∞ > 1)
We present here two solution methods for PDE equation:
1. Conical Flow Method
2. Singularity-Distribution Method
1&1:0 2
2
2
2
2
2
2
MM
zyx
This method was proposed by Busemann in 1943 and was
extensively used before high speed computers were available.
A Conical Flow is defined by velocity, pressure , static
temperature, density constant along rays, through a common
vertex.
The Conical Flow can occur only at Supersonic Speeds.
Conical Flow are produced by passing over a conic body, but
It can be produced by small supersonic perturbations if the
Boundary Conditions satisfy the Conical Conditions.
In Supersonic Flow the disturbances are propagated only
downstream the Mach Cone.
Adolph
Busemann
(1901 – 1986).
This method is similar with that used in Incompressible Flow, but the
Singularities are Solutions of Supersonic Hyperbolic PDE.
Return to Table of Content
15. 15
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Finite Span Wing
in a Supersonic Flow (M∞ > 1)
1. Conical Flow Method
1&1:0 2
2
2
2
2
2
2
MM
zyx
Use for the Conical Flow the potential
Start with
x
z
x
y
fxzyx
:,:
,:,,
ff
f
x
2
222
2
2
2
2
1111
ff
x
f
x
ff
x
f
x
f
x
f
xx
Let compute
22
2
22
2
1
,
1
,
f
xz
f
z
f
xy
f
y
1&1:0/12/1 2
2
2
22
2
2
2
22
MM
fff
Mach Cone
16. 16
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Finite Span Wing
in a Supersonic Flow (M∞ > 1)
1. Conical Flow Method
Mach Cone
1&1:0/12/1 2
2
2
22
2
2
2
22
MM
fff
x
z
x
y
fxzyx
:,:
,:,,
Let compute
,,,,' fzyx
x
u
The equation of a ray starting at the origin is given by 2121,, cc
x
z
c
x
y
czyxr
We can see that for η = const., ζ = const., we have r (x,y,z) = const.
.,
2
1
,'.
,'
2,
.,'
2
constCUpconst
U
u
C
constu
pp
.
,'
1
2,'
1
,','
const
a
a
T
T
p
p
Isentropic Chain
17. 17
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Finite Span Wing
in a Supersonic Flow (M∞ > 1)
1. Conical Flow Method
Regions where Two-Dimensional Flow prevails
on Three-Dimensional Wings .
Shaded zones signify Two-Dimensional Flow.
Because in Supersonic Flows a perturbation is
felt only in the Mach cone downstream from the
source of disturbance, certain portion of the
Wings behave as though they were in the Two –
Dimensional Flow.
18. 18
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Inclined Rectangular Wing at
Supersonic Flow
(a) Planform
(b) Pressure disturbance at A-A Section
Conical Flow on Rectangular Wings
Propagations of Wing Edges (Leading and Side) on
the Supersonic Flow propagate over Mach Cones.
Looking at the Section A-E-A of the Wing, where E
is the intersection of Section A-A with the Mach
Line from the Wing Tip, we see that:
• Points on A-E (region II) are affected only by
the disturbances of the Wing Leading Edge. The
Flow is Conical and two dimensional on the
Wing, therefore the Pressure Coefficient is given
by
22
1
4
2/
MU
pp
cc plpp
• Points on E-A (region IV) are affected by
the disturbances of the Wing Leading Edge
and by the Side Edge. The Flow is Conical
and two dimensional on the Wing.
the Pressure Coefficient is given by
21
2
121cos
1
4
M
x
y
tt
M
cp
II
IV
A A
E
EdgeLeadingt
EdgeSidet
M
x
y
tt
c
c
plp
p
1
0
121cos 21
y
x
Area
Below Curve =
The mean value for is . 1,0t plpp cc 5.0
19. 19
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Mach ConeMach Cone
Wing
Leading Edge
Wing
Leading Edge
Wing
Leading Edge
Region II: AMNB
Region IV: ADM & BCN
Region II: ABE
Region IV: ADME & BCNE
Region V: MNE
Region II: ABE
Region IV: AME & BNE
Region V: MFNE
Conical Flow on Rectangular Wings
Propagations of Wing Edges (Leading
and Side) on the Supersonic Flow
propagate over Mach Cones.
Different Regions on the Wing are
affected by the Wing Edges.
Region II:
Flow over points on the Wing in
this region are affected only by
disturbances of Leading Edge.
Region IV:
Flow over points on the Wing in this
region are affected by disturbances of
both Leading Edge and one of Side
Edges.
Region V:
Flow over points on the Wing in this region are affected by disturbances of
Leading Edge and both Side Edges. The Flow is not Conical.
20. 20
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Conical Flow on Rectangular Wings
Aerodynamic Forces on Inclined Rectangular Wings of various Aspect Ratios at
Supersonic Incident Flow
(a) Lift Slope
(b) Neutral-point Position
(c) Drag Coefficient
21. 21
SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Conical Flow on Rectangular Wings
Pressure Distribution over the Chord
and Lift Distribution over the Span for
the Inclined Rectangular Plate of
Aspect Ratio AR = 2.5 at Supersonic
Incident Flow
89.1;41 2
MMARa
13.1;
3
4
1 2
MMARb
22. n
Mm
1
:1tan
tan
tan 2
1
4
2/ 22
MU
pp
c plp
tan
1
:
:
x
y
t
IRange
mtM
x
y
M
x
y
t
IIIandIIRange
1tan
tan
1
tan
'tan
1
:
22
10';sin11:'
2/
0
22
EdmmE
Basic Solution for Pressure Distribution of the Inclined Flat Surface in Supersonic
Incident Flow (Cone-Symmetric Flow) for Ranges I, II, III and IV
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Finite Span Wing
in a Supersonic Flow (M∞ > 1)
1. Conical Flow Method
23. SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Conical Flow on Swept-Back Wings
Pressure Distribution over in the
Wing Chord (schematic) for a
section of an Inclined Swept-Back
Wing
(a) Subsonic Leading and
Trailing Edges.
(b) Subsonic Leading and
Supersonic Trailing Edge.
(c) Supersonic Leading and
Trailing Edges.
24. SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Conical Flow on Swept-Back Wings
Pressure Distribution over in the
Wing Chord and Lift Distribution
over the Wing Span of Delta
Wings at Supersonic Incident
Flow
(a) Subsonic Leading Edge,
0 < m < 1.
(b) Supersonic Leading Edge,
m > 1.
n
Mm
1
:1tan
tan
tan 2
25. SOLO Wings in Compressible Flow
Wings of Finite Span at Supersonic Incident Flow
Conical Flow on Swept-Back Wings
Lift Distribution over the
Span of Delta Wings at
Supersonic Incident Flow for
several values of m:
• Subsonic Leading Edge,
0 < m < 1.
• Supersonic Leading Edge,
m > 1.
n
Mm
1
:1tan
tan
tan 2
Return to Table of Content
26. 26
Linearized Flow EquationsSOLO
Incompressible Flow (M∞ = 0)
Velocity Potential Equations:
A Particular Solution is
R
Q
zyx
Q
44 222
That can be rewritten as
Q – Source Strength
Compressible Subsonic Flow (0 < M∞ < 1)
0
11
2
2
2
2
2
2
2
2
zMyMx
Potential Equation:
A Particular Solution is
2222
14 zyMx
Q
That can be rewritten as
Q – Subsonic Compressible Source Strength
1
4/4/4/
222
Q
z
Q
y
Q
x
Sphere
1
4/
1
1
4/
1
14/
2
2
2
2
2
Q
M
z
Q
M
y
Q
x
Ellipsoid of Revolution
02
2
2
2
2
2
zyx
Elliptic Second Order Linear
Partial Differential Equation.
Elliptic Second Order Linear
Partial Differential Equation.
2. Singularity-Distribution Method
27. 27
Linearized Flow EquationsSOLO
Compressible Supersonic Flow (M∞ >1)
1,0
11
2
2
2
2
2
2
2
2
i
zMiyMix
Velocity Potential Equation:
By analogy with the Subsonic Flow a Particular Solution is
2222
14 zyMx
Q
That can be rewritten as
Q – Supersonic Compressible Source Strength
1
4/
1
1
4/
1
14/
2
2
2
2
2
Q
M
z
Q
M
y
Q
x
Hyperboloid of Revolution
Only the part of the Flow lying downstream Mach Cone is physically significant.
Hyperbolic Second Order Linear Partial Differential Equation.
2. Singularity-Distribution Method
28. 28
SOLO Wings in Compressible Flow
2. Singularity-Distribution Method for Supersonic Flow (M∞ >1)
Velocity Potential Equation:
1&1:0 2
2
2
2
2
2
2
MM
zyx
Flow is Linear even without the assumption of Small Disturbances. This allows to combine Elementary
Solutions similar to Subsonic Incompressible Flow (I.e. Source, Sink, Doublet, Vortex, etc.) to obtain
General Solution for Supersonic Flow. Those Elementary Solutions are spread on the Aerodynamic
Bodies in such a way that satisfy the Boundary Conditions.
Example of Supersonic Elementary Solutions are:
c
S
r
q
4
Source
Doublet
c
c
V
r
vzq
4
Vortex
where
22
1
1
22
2/122
1
22
1
:
1:
:
zyy
xx
v
M
zyyxxr
c
c
H. Lomax, M.A., Heaslet, F.B., Fuller, “Integrals and
Integral Equations in Linearized Wing Theory”,
Report 1054, NACA 1951zr
zq
c
D
3
2
4
29. 29
SOLO Wings in Compressible Flow
2. Singularity-Distribution Method for Supersonic Flow (M∞ >1)
Four types of problems can be treated by the Singularity Distribution Method:
(a) Two Non-lifting Case (Symmetric Wing):
1. Given the Thickness Distribution and the Planform Shape, find the Pressure
Distribution on the Wing.
2. Given the Pressure Distribution on a Wing of Symmetrical Section, find the
Wing Shape (I.e. the Thickness Distribution and the Planform).
(b) Two Lifting Case (Non-Symmetric Wing):
4. A Lifting Surface, find the Pressure Distribution on it. In the Subsonic Case it is
necessary to satisfy the Kutta Condition at the Trailing Edge.
3. Given the Pressure Distribution on a Lifting Surface (Zero Thickness)
find the Slope of each point on the Surface.
Direct Problems: Cases 1 and 3, because they involve Integrals with known Integrands.
Indirect Problems: Cases 2 and 4, because the Unknown is inside the Integral Sign.
Cases 1 and 2 are more conveniently solved using Source or Doublet Distributions,
while Cases 3 and 4 are most often treated using Vortex Distributions.
Return to Table of Content
30. 30
SOLO Wings in Compressible Flow
Theoretical Solutions for Compressible Supersonic Flow (M∞ >1)
52&8.001 2
2
2
2
2
2
2
MM
zyx
M
Velocity Potential Equation:
1&1:0 2
2
2
2
2
2
2
MM
zyx
By analogy with the Subsonic Flow the influence of the
Point Source q located at (ξ’, η’, 0) is given by
2222
''4
''0,','
,,
zyx
ddq
zyxd
The Point Source q must be such that whose boundary are defined by 2222
'' zyx
This is a Mach Cone, with apex at (ξ’, η’, 0) and angle μ = cot-1β
1 2
10 2222
''4
''0,','
,,
zyx
ddq
zyx
zx
zxy
zxy
1
222
2
222
1
/
/
0'' 2222
zyx
z
w
y
v
x
u
zwUyvxuUu
',','
1'1'1'
Elementary Source
OfStrength q dξ dη
Elementary Source
OfStrength q dξ dη
Hyperbola (ξ, η) :
31. 31
Elementary Source
OfStrength q dξ dη
Elementary Source
OfStrength q dξ dη
Hyperbola (ξ, η) :
SOLO Wings in Compressible Flow
Theoretical Solutions for Compressible Supersonic Flow (M∞ >1) (continue -1)
Let integrate for all Sources (ξ, η, 0) (on the Wing)
that are in the Front Mach Cone with the apex at
(x,y,z)
1 2
10 2222
4
0,,
,,
zyx
ddq
zyx
The boundary are defined by 222
2
222
11 /,/, zxyzxyzx
From this we can compute
1 2
10 2/32222
2
4
0,,,,
,,
zyx
ddzq
z
zyx
zyxw
We can see that w (x, y, z = 0) is zero everywhere, except at the source x = ξ, y = η where we have a
indeterminate value 0/0. This was solved by Puckett in his PhD Thesis at Caltech in 1946
For ϕ (x,y,z), integrate the second
integral by parts
222
1
2222
sin
1
4
1
/
4
,
zx
y
vd
q
ud
zyxddv
q
u
Note that
12
/
/
222
1
,,
8
1
sin,
4
1
222
12
222
11
qq
zx
y
qvu
zxy
zxy
36. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β)
Pressure Field for a Semi-Infinite Triangular Wing
with a Subsonic Leading Edge
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from P are PB
and PD.
The Parts of the Wing that influence the Flow at P
are located in the Area AEPBA.
CPB
AEPC
AEPBA
P
yx
dd
yx
ddU
yx
ddU
zyx
222
222
222
0,,
yxBPalong
ACBalong
:
tan:
tan
,
tan
tan yxyx
B
The Limits of Integrations are defined by the
points A, E, P, B, C. The Lines of Integrations are
37. 37
SOLO Wings in Compressible Flow
CPB
yx
yx
y
AEPC
yxy
P
yx
dd
d
yx
dd
d
U
zyx
tan 222
tan
tan 2220
0,,
yy
yx
yxy
y
x
d
U
y
x
d
U
yx
dd
d
U
0
1
0
1
0
tan
1
tan 2220
tan
cosh1coshcosh
tan
0
1
0
1tan
tan
1
tan 222
tan
tan
cosh1coshcosh
yxyx
y
yx
yx
yx
y y
x
d
U
y
x
d
U
yx
dd
d
U
Section aa
Mach Cone
From P
Mach Cone
From P
32
tan 1
0
1 tan
cosh
tan
cosh0,,
I
yx
y
I
y
P
y
x
d
y
x
d
U
zyx
Let compute
x
zyx
zyxu P
P
0,,
0,,
tan
tan
cosh
tan
10,,
0,,
2
1
22 yx
yxU
x
zyx
zyxu P
P
We obtain
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 1)
38. 38
SOLO Wings in Compressible Flow
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
tn
tn
n
U
x
zyx
zyxuP
1
cosh
1
10,,
0,,
2
1
2
Therefore on the Wing ( t = 0 – Side Edge to t = 1 - Leading Edge)
11
tan
tan 2
2
22
n
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing apex
t =0 (Side Edge), t = 1 (Leading Edge)
We found
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 2)
tan
tan
cosh
tan
10,,
0,,
2
1
22 yx
yxU
x
zyx
zyxuP
Section aa
Mach Cone
From P
Mach Cone
From P
tn
tn
nU
zyxu
Cp
1
cosh
1
120,,
2
2
1
2
39. 39
SOLO Wings in Compressible Flow
Let find how the disturbances of the Wing on the
Flow affect a point N (x,y,0) outside the Wing
between the Wing Side-Edge and the Mach Line
(see Figure). The Mach Line through N that
intersects The Wing between the points L and J
Determines the Wing area ALN that affects the Flow
at N.
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 3)
Section aa
Mach Cone
From N
Mach Cone
From N
ALN
N
yx
ddU
zyx
222
0,,
yxNLJalong
AJalong
:
tan:
0,/
tan
,
tan
tan
yxL
yxyx
J
The Limits of Integrations are defined by the
points A, L,J. The Lines of Integrations are
tan
0
1
0
1tan
0
tan
1
tan 222
tan
0
tan
cosh1coshcosh
yxyx yx
yx
yx
N
y
x
d
U
y
x
d
U
yx
dd
d
U
tan
0
1 tan
cosh0,,
yx
N d
y
xU
zyx
40. 40
SOLO Wings in Compressible Flow
Let find how the disturbances of the Wing on the
Flow affect a point N (x,y,0)
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 4)
Section aa
Mach Cone
From N
Mach Cone
From N
tan
0
1 tan
cosh0,,
yx
N d
y
xU
zyx
tan
tan
cosh
tan
tan
tan
tan
tan
cosh
tan
1
tan
1
2
1
tan
0
22
22
22
2
1
22
tan
0 222
yx
yxU
yx
yx
U
yx
d
U
x
u
yx
yx
N
N
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
11
tan
tan 2
2
22
n
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing
apex , t = 0 (Side Edge), t =- n (Mach Line)
41. 41
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 5)
Section aa
Mach Cone
From N
Mach Cone
From N
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing
apex , t = 0 (Side Edge), t =- n (Mach Line)
tn
tn
n
U
x
zyx
zyxuN
1
cosh
1
10,,
0,,
2
1
2
Therefore between t = 0 (Side Edge )
to t = -n (Mach Line)
tn
tn
nU
zyxu
Cp
1
cosh
1
120,,
2
2
1
2
42. 42
SOLO Wings in Compressible Flow
Let find how the disturbances of the Wing on the
Flow affect a point L (x,y,0) outside the Wing
between the Wing Leading-Edge and the Mach
Line(see Figure). The Mach Line through L that
intersects The Wing between the points J and G
Determines the Wing area AJG that affects the
Flow at L.
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 6)
Section aa
Mach Cone
From A
Mach Cone
From A
AJG
L
yx
ddU
zyx
222
0,,
yxGJNalong
AJalong
:
tan:
0,
tan
,
tan
tan
yxG
yxyx
J
The Limits of Integrations are defined by the
points A, L,J. The Lines of Integrations are
tan
0
1
0
1tan
0
tan
1
tan 222
tan
0
tan
cosh1coshcosh
yxyx yx
yx
yx
y
x
d
U
y
x
d
U
yx
dd
d
U
tan
0
1 tan
cosh0,,
yx
L d
y
xU
zyx
43. 43
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 7)
Section aa
Mach Cone
From A
Mach Cone
From A
tan
0
1 tan
cosh0,,
yx
L d
y
xU
zyx
Let find how the disturbances of the Wing on the
Flow affect a point L (x,y,0)
xy
yxU
xy
yx
U
yx
d
U
x
u
yx
yx
N
N
tan
tan
cosh
tan
tan
tan
tan
tan
cosh
tan
1
tan
1
2
1
tan
0
22
22
22
2
1
22
tan
0 222
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing
apex , t =1(Leading Edge) to t = n (Mach Line)
1
1tan
tan
tan
tan
1tan
tan
tan
tan 2
2
2
tn
tn
x
y
x
y
x
y
x
y
xy
yx
11
tan
tan 2
2
22
n
44. 44
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 8)
Section aa
Mach Cone
From A
Mach Cone
From A
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing
apex , t =1(Leading Edge) to t =+ n (Mach Line)
1
cosh
1
10,,
0,,
2
1
2 tn
tn
n
U
x
zyx
zyxuL
Therefore between t = 1 (Leading Edge )
to t = +n (Mach Line)
1
cosh
1
120,,
2
2
1
2 tn
tn
nU
zyxu
C L
pL
45. 45
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Subsonic Leading Edge (tanΛ > β) (continue – 9)
Pressure Field for a Semi-Infinite Triangular Wing with a Subsonic Leading Edge
Return to Table of Content
46. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β)
Pressure Field for a Semi-Infinite Triangular
Wing with a Subsonic Leading Edge
Consider the Point L (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from L are LB
and LC. B and C are on the Wing Leading Edge.
The Parts of the Wing that influence the Flow at L
are located in the Area LBC.
LBC
L
yx
ddU
zyx
222
0,,
tan:
:
:
CBalong
yxBLalong
yxCLalong
tan
,
tan
tan
tan
,
tan
tan
yxyx
C
yxyx
B
The Limits of Integrations are defined by the
points C, L and B. The Lines of Integrations are
Mach Line
yx
yx
y
yxy
yxL
yx
dd
d
U
yx
dd
d
U
tan 222
tan
tan 222
tan
Section aa
47. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) (continue – 1)
Pressure Field for a Semi-Infinite
Triangular Wing with a Subsonic Leading
Edge
Section aa
Consider the Point L (x,y,z=0) on a Single Wedge
Triangular Wing.
y
x
y
x
y
x
y
d
yx
d
yx
yx yx
tan
cosh1coshcosh
1
1
0
1
tan
1
tan tan 2222
y
x
y
x
y
x
y
d
yx
d
yx
yx yx
tan
cosh1coshcosh
1
1
0
1
tan
1
tan tan 2222
tan
tan
1
tan
1
tan
1 tan
cosh
tan
cosh
tan
cosh
yx
yx
yx
y
y
yx
L
y
xU
y
x
d
U
y
x
d
U
yx
yx
y
yxy
yxL
yx
dd
d
U
yx
dd
d
U
tan 222
tan
tan 222
tan
48. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) (continue – 2)
Consider the Point L (x,y,z=0) on a Single Wedge
Triangular Wing.
tan
tan
1 tan
cosh
yx
yx
L
L d
y
x
x
U
x
u
tan
tan
222
0
1
1
0
1
1
tan
tan
tan
tan
cosh
tan
1
tan
tan
tan
cosh
tan
1
yx
yx yx
dU
yx
y
yx
x
U
yx
y
yx
x
U
tan
tan
222
tan
yx
yx
L
yx
dU
u
tan
tan
1 tan
cosh
yx
yx
L d
y
xU
Section aa
49. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) (continue – 3)
Consider the Point L (x,y,z=0) on a Single Wedge
Triangular Wing.
2
22
22
2
22
22
22
22222
tan
tan
tan
tan
tan
1
tan
tan
tan
tan
1
tan
yx
yx
yx
d
yx
d
tan
tan
22
22
2
22
1
22
tan
tan
tan
tan
tan
sin
tan
1
yx
yx
L
yx
yx
U
u
2
1
1
1
sin
uxd
ud
xu
xd
d
use
2/
1
22
22
2
22
1
22
2/
1
22
22
2
22
1
22
tan
tan
tan
tan
tan
tan
sin
tan
1
tan
tan
tan
tan
tan
tan
sin
tan
1
yx
yxyx
U
yx
yxyx
U
22
tan
U
uL
tan
tan
222
tan
yx
yx
L
yx
dU
u
Section aa
50. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) (continue – 4)
Consider the Point L (x,y,z=0) on a Single Wedge
Triangular Wing.
2
222
1:,
tan
:
1tan
Mn
n
UU
x
u
L
L
Section aa
51. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β ) (continue – 5)
Pressure Field for a Semi-Infinite Triangular Wing
with a Subsonic Leading Edge
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from P are PB
and PD.
The Parts of the Wing that influence the Flow at P
are located in the Area AEPBA.
AEDBPD
AEPBA
P
yx
dd
yx
ddU
yx
ddU
zyx
222222
222
0,,
yxDEPalong
DACBalong
:
tan:
0,
tan
,
tan
tan
yxE
yxyx
D
The Limits of Integrations are defined by the
points A, E, P, B, C. The Lines of Integrations are
Mach Line
AED
yx
yx
BPD
P
yx
dd
dd
xU
zyx
0
tan
tan
22222
tan
0,,
52. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) ) (continue – 6)
Pressure Field for a Semi-Infinite
Triangular Wing with a Subsonic Leading
Edge
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from P are PB
and PD.
AED
yx
yx
BPD
P
yx
dd
dd
xU
zyx
0
tan
tan
22222
tan
0,,
y
x
y
x
y
x
y
d
yx
d
yx
yx yx
tan
cosh1coshcosh
1
1
0
1
tan
1
tan tan 2222
AED
yx
BPD
P
y
x
d
xU
zyx
0
tan
1
22
tan
cosh
tan
0,,
0
tan
222
0
1
1
22
tan
tan
tan
tan
cosh
tan
1
tan
yx
P
P
yx
d
yx
y
yx
x
U
x
u
53. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) ) (continue – 7)
Pressure Field for a Semi-Infinite
Triangular Wing with a Subsonic Leading
Edge
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from P are PB
and PD.
0
tan
22222
tantan
yx
P
P
yx
dU
x
u
0
tan
22
22
2
22
1
22
tan
tan
tan
tan
tan
sin
tan
1
yx
L
yx
yx
U
u
2/
1
22
22
2
22
1
2
1
22
tan
tan
tan
tan
tan
tan
sin
tan
tan
sin
tan
1
yx
yxyx
yx
yxU
tan
tan
sin
2tan
1 2
1
22 yx
yxU
uL
54. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) ) (continue – 8)
Pressure Field for a Semi-Infinite
Triangular Wing with a Subsonic Leading
Edge
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Triangular Wing. The Mach Lines from P are PB
and PD.
tan
tan
sin
2tan
1 2
1
22 yx
yxU
uL
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
2
2
22
1
tan
1tan n
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing apex
t = 0 (Side Edge), t = n (Leading Edge)
tn
tn
n
U
tn
tn
n
U
uL
1
cos
1
1
1
sin
21
1 2
1
2
2
1
2
55. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) ) (continue – 8)
Pressure Field for a Semi-Infinite
Triangular Wing with a Subsonic Leading
Edge
Section aa
Mach Cone
From P
Mach Cone
From N
Consider the Point N (x,y,z=0) between the Side
Edge of the Triangular and the Mach Lines from A
outside the Wing Planform. The f;ow disturbance on
N is due to Wing Surface AEC.
tan
tan
cos
tan
1 2
1
22 yx
yxU
x
u N
N
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing apex
t -n (Mach Line), t = 0 (Side Edge)
tn
tn
n
U
tn
tn
n
U
uL
1
cos
1
1
1
sin
21
1 2
1
2
2
1
2
ANC
N
yx
ddU
zyx
222
0,,
yxCEalong
ACalong
:
tan:
0,
tan
,
tan
tan
yxE
yxyx
C
The Limits of Integrations are defined by the
points A, E, C. The Lines of Integrations are
tan
0 tan
222
0,,
yx
yx
N
yx
dd
dd
U
zyx
56. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow and Supersonic Leading Edge (tanΛ < β) ) (continue – 9)
Mach Line
Pressure Field for a Semi-Infinite Triangular Wing with a Supersonic Leading Edge
Return to Table of Content
57. 57
SOLO Wings in Compressible Flow
Section aa
Mach Cone
From P
Mach Cone
From P
Consider the Point P (x,y,z=0) on a Single Wedge
Delta Wing. The Mach Lines from P are PB and PD.
The Parts of the Wing that influence the Flow at P
are located in the Area ADPBA.
CPB
AEPC
ADE
ADPBA
yx
dd
yx
dd
yx
ddU
yx
ddU
zyx
222
222
222
222
0,,
The Limits of Integrations are defined by the points A, D, E, P, B, C. The Lines of Integrations are
yxDEPalong
yxBPalong
ACBalong
ADalong
:
:
tan:
tan:
tan
,
tan
tan
tan
,
tan
tan
yxyx
D
yxyx
B
Based on: A.E. Puckett, “Supersonic Wave Drag of Thin Airfoils”, 1949, Caltech PhD Thesis
http://thesis.library.caltech.edu/2697/1/Puckett_ae_1949.pdf
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β)
58. 58
SOLO Wings in Compressible Flow
Section aa
Mach Cone
From P
Mach Cone
From P
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 1)
CPB
yx
yx
y
AEPC
yxy
AED
yx
yx
yx
dd
d
yx
dd
d
yx
dd
d
U
zyx
DE
AD
tan 222
tan
tan 2220
tan 222
0
tan
0,,
0
tan
1
0
1
0
tan tan
1
tan 222
0
tan
tan
cosh1coshcosh
yxyx
yx
yx
yx
y
x
d
U
y
x
d
U
yx
dd
d
U
yy
yx
yxy
y
x
d
U
y
x
d
U
yx
dd
d
U
0
1
0
1
0
tan
1
tan 2220
tan
cosh1coshcosh
tan
0
1
0
1tan
tan
1
tan 222
tan
tan
cosh1coshcosh
yxyx
y
yx
yx
yx
y y
x
d
U
y
x
d
U
yx
dd
d
U
59. 59
SOLO Wings in Compressible Flow
321
tan 1
0
1
0
tan
1 tan
cosh
tan
cosh
tan
cosh0,,
I
yx
y
I
y
I
yx
y
x
d
y
x
d
y
x
d
U
zyx
We want to compute
x
zyx
zyxu
0,,
0,,
We use LEIBNIZ THEOREM from CALCULUS:
)(
)(
)(
)(
),(
)),(()),((),(::
tb
ta
ChangeBoundariesthetodueChange
sb
sa
dx
s
sxf
sd
sad
ssaf
sd
sbd
ssbfdxsxf
sd
d
LEIBNITZ
yy
yx
d
y
x
d
xd
d
0 2220
1
tan
1tan
cosh
and
1
1
cosh
2
1
uxd
ud
xu
xd
d
tan
222
0
1
1tan 1
tan
1
tan
tan
tan
cosh
tan
1tan
cosh
yx
y
yx
y
yx
d
yx
y
yx
x
y
x
d
xd
d
0
tan
222
0
1
1
0
tan
1
tan
1
tan
tan
tan
cosh
tan
1tan
cosh
yxyx
yx
d
yx
y
yx
x
y
x
d
dx
d
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 2)
60. 60
SOLO Wings in Compressible Flow
2
22
2
22
22
2
222
tan
tan
tan
tan
tan
tan
yxyx
yx
1
tan
tan
tan
tan
tan
tan
tan
tan
tan
1
tan
2
22
22
22
2
22
22
22222
yx
yx
yx
d
yx
d
0
tan
22
22
22
2
1
22
0
tan
222
1
tan
tan
tan
tan
tan
cosh
tan
1
tan
1
yx
yx
yx
yx
yx
d
x
I
1
1
cosh
2
1
uxd
ud
xu
xd
d
use
0
tan
222
1
tan
1
yx
yx
d
x
I
EdgeLeadingSubsonictan
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 3)
Start with
61. 61
SOLO Wings in Compressible Flow
2
22
2
22
22
2
222
tan
tan
tan
tan
tan
tan
yxyx
yx
1
tan
tan
tan
tan
tan
tan
tan
tan
tan
1
tan
2
22
22
22
2
22
22
22222
yx
yx
yx
d
yx
d
y
yx
d
x
I
0 222
2
tan
1
y
y
yx
yx
yx
d
x
I
0
22
22
22
2
1
220 222
2
tan
tan
tan
tan
tan
cosh
tan
1
tan
1
1
1
cosh
2
1
uxd
ud
xu
xd
d
use
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 4)
62. 62
SOLO Wings in Compressible Flow
2
22
2
22
22
2
222
tan
tan
tan
tan
tan
tan
yxyx
yx
1
tan
tan
tan
tan
tan
tan
tan
tan
tan
1
tan
2
22
22
22
2
22
22
22222
yx
yx
yx
d
yx
d
tan
222
3
tan
1yx
y
yx
d
x
I
tan
22
22
22
2
1
22
tan
0 222
3
tan
tan
tan
tan
tan
cosh
tan
1
tan
1
yx
y
yx
yx
yx
yx
d
x
I
1
1
cosh
2
1
uxd
ud
xu
xd
d
use
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 5)
64. 64
SOLO Wings in Compressible Flow
tan
tan
cosh
tan
tan
cosh
tan
10,,
0,,
2
1
2
1
22 yx
yx
yx
yxU
x
zyx
zyxu
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
tn
tn
tn
tn
n
U
x
zyx
zyxu
1
cosh
1
cosh
1
10,,
0,,
2
1
2
1
2
Therefore
11
tan
tan 2
2
22
n
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 7)
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing apex
We found
65. 65
SOLO Wings in Compressible Flow
tn
tn
tn
tn
n
U
x
zyx
zyxu
1
cosh
1
cosh
1
10,,
0,,
2
1
2
1
2
We want to prove that
2
22
1
2
1
2
1
1
cosh2
1
cosh
1
cosh
t
tn
tn
tn
tn
tn
2
22
1
2 1
cosh
1
120,,
0,,
t
tn
n
U
x
zyx
zyxu
tan
:&tan:,: n
x
y
t
xd
zd
S
Finally we obtain
tn
tn
tn
tn
1
:cosh,
1
:cosh
22
Define
Let compute 2
sinh
2
sinh
2
cosh
2
cosh
2
cosh
tn
tnn
tn
tnn
tn
tnn
tn
tnn
12
1
2/1cosh
2
sinh,
12
1
2/1cosh
2
cosh
12
1
2/1cosh
2
sinh,
12
1
2/1cosh
2
cosh
2
22
2
22
2
22
112
1
12
1
2
cosh
t
tn
t
tn
n
n
t
tn
n
n
q.e.d.
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 8)
66. 66
SOLO Wings in Compressible Flow
1&
1
cosh
1
120,,
0,, 2
22
1
2
ttn
t
tn
n
U
x
zyx
zyxu
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Subsonic Leading Edge (tanΛ > β) (continue – 9)
Since the Pressure and Velocity are constant along
t = (y/x) tan Λ, i.e. along rays through the vertex of
the Delta Wing, the Solution is of Conical Flows.
For t = 1 we get the ray corresponding to the
Leading Edge. For t = n=tanΛ/β we get the ray
along the “Mach Line” from the vertex of the Delta
Wing.
1&
1
cosh
1
140,,2
2
22
1
2
ttn
t
tn
nU
zyxu
Cp
Pressure Coefficient
1:,
tan
:&tan:,: 2
Mn
x
y
t
xd
zd
S
67. 67Theoretical Solution for a Delta Wing
(a) Pressure Distribution for a Single-Wedge Delta Wing at α = 0 [From Puckett (1946)]
SOLO Wings in Compressible Flow
68. 68
Theoretical Solution for a Delta Wing
(b) Thickness Drag of a Double-Wedge Delta Wing with a Supersonic Leading Edge
and a Supersonic Line of Maximum Thickness [From Puckett (1946)]
SOLO Wings in Compressible Flow
69. 69
Theoretical Solution for a Delta Wing
(c) Thickness Drag of a Double-Wedge Delta Wing with a Supersonic Line of
Maximum Thickness [From Puckett (1946)]
SOLO Wings in Compressible Flow
Return to Table of Content
70. 70
SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β)
Consider first a Point P1 (x,y) on the Wing, lying between
the Wing Leading Edge and the Mach Line
(1 > t > n). This point have a Potential determined only by
the Sources lying in region (1) (Defined by Mach Lines P1A1
and P1C1 intersecting only the swept Trailing Edge OA1).
But this Potential must be the same as for an Infinite Sweep
(Λ) Wing, therefore is given by
2
2121 1:,
tan
:
1
,
1 1
Mn
n
U
x
Pu
n
xU
P
P
1
2
The Point P2 (x,y) on the Wing is lying behind the Mach
Lines from the Wing Tip. The Mach Line PA
intersects the Leading Edge OA and the Mach Line PC
intersects the Leading Edge OB. If only the Leading Edge
OA exists (no Leading Edge OB) than the Potential at P2
would be the Same as P1.
To consider the Leading Edge OB we must subtract the
disturbances in the area of region (2) OBC (no sources)
2
)2(
22222 1:,
tan
:,:,
1
Mn
x
z
yx
ddU
n
xU
P
S
2
71. 71
SOLO Wings in Compressible Flow
ODB
y
CDB
y
y
yx
yx
dd
d
U
yx
dd
d
U
yx
ddU
zyx
0 tan
tan 222tan 222
)2(
222
2
2
1
0,,
tan1
:tan
1
11
11
yx
y
MLyyxx
OALEyx
tan2
:tan
2
22
22
yx
y
MLyyxx
OBLEyx
y
x
y
x
y
x
y
d
yx
d
yx
yx yx tan
cosh1coshcosh
1
1
0
1
tan
1
tan tan 2222
y
x
y
x
y
x
y
x
y
d
yx
d tan
cosh
tan
coshcosh
1
11
tan
tan
1
tan
tan 2
tan
tan 222
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 1)
2
72. 72
SOLO Wings in Compressible Flow
)2(
222
yx
ddU
0
111
2
2
1
tan
cosh
tan
cosh
tan
cosh
y
y
y
d
y
x
y
xU
d
y
xU
0
1
0
1
21
tan
cosh
tan
cosh
yy
d
y
xU
d
y
xU
The u – velocity associated with this potential is given by
2
2
1
1
21
21
0
222
0
222
0
222
0
1
2
212
0
222
0
1
1
111
0
1
0
1
tantan
tan
tan
cosh
tan
tan
cosh
tan
cosh
tan
cosh
I
y
I
y
yy
yy
yx
dU
yx
dU
yx
dU
yy
yx
x
yU
yx
dU
yy
yx
x
yU
d
y
x
x
U
d
y
x
x
U
x
u
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 2)
2
73. 73
SOLO Wings in Compressible Flow
0
2221
1
tan
y
yx
dU
I
73
2
22
22
2
2
22
222
tan
tan
tan
tan
tan
tan
yxyx
yx
2
22
22
22
2
22
22
22222
tan
tan
tan
tan
tan
1
tan
tan
tan
tan
1
tan
yx
yx
yx
d
yx
d
0
tan
22
22
22
2
1
22
0
2221
1
1
tan
tan
tan
tan
tan
sin
tan
1
tan
yx
y
y yx
yx
U
yx
dU
I
2
1
1
1
sin
uxd
ud
xu
xd
d
use
2/
1
2
1
22
2
1
22 tan
tantan
sin
tan
1
tan
tan
sin
tan
1
yx
yxyxU
yx
yxU
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 3)
74. 74
SOLO Wings in Compressible Flow
74
2
22
2
22
2
22
222
tan
tan
tan
tan
tan
tan
yxyx
yx
2
22
22
2
22
22
22
22222
tan
tan
tan
tan
tan
1
tan
tan
tan
tan
1
tan
yx
yx
yx
d
yx
d
0
tan
22
22
2
22
1
22
0
2222
2
2
tan
tan
tan
tan
tan
sin
tan
1
tan
yx
y
y yx
yx
U
yx
dU
I
2
1
1
1
sin
uxd
ud
xu
xd
d
use
0
2222
2
tan
y
yx
dU
I
2/
1
2
1
22
2
1
22 tan
tantan
sin
tan
1
tan
tan
sin
tan
1
yx
yxyxU
yx
yxU
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 4)
75. 75
SOLO Wings in Compressible Flow
75
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 5)
2
2
1
1
0
222
0
222
tantan
I
y
I
y
yx
dU
yx
dU
x
u
21
22
2
1
2222
2
1
22
tan
1
2tan
tan
sin
tan
1
tan
1
2tan
tan
sin
tan
1
II
U
yx
yxUU
yx
yxU
S
x
z
yx
yx
yx
yxU
:
tan
tan
sin
tan
tan
sin
tan
1 2
1
2
1
22
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
tn
tn
x
y
x
y
x
y
x
y
yx
yx
1
tan1
tan
tan
tan
tan1
tan
tan
tan 2
2
2
2
2
22
1
tan
1tan n
Define 1:1
tan
:&tan: 2
Mn
x
y
t
y/x=t/tanΛ is the equation of a ray starting from Wing apex
tn
tn
tn
tn
n
U
x
u
1
sin
1
sin
1
1 2
1
2
1
2
76. 76
SOLO Wings in Compressible Flow
76
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 6)
tn
tn
tn
tn
n
U
x
u
1
sin
1
sin
1
1 2
1
2
1
2
We want to prove that
2
22
1
2
1
2
1
1
sin2
1
sin
1
sin
t
tn
tn
tn
tn
tn
2
22
1
2 1
sin
1
12
t
tn
n
U
x
u
2
22
1
222
1
sin
2
1
1
1
1
1
2
t
tn
n
U
u
n
U
x
Pu
P
tn
tn
tn
tn
1
:sin,
1
:sin
22
Define
Let compute
2
sin2
2
cos
2
sin2
2
cos
2
sin2
2
sin
2
cos
2
sin
2
cos
2
sin
2
cos
2
sin
2
cos
tn
tnn
tn
tnn
tn
tnn
tn
tnn
1
1
sin1
2
sin
2
cos,
1
1
sin1
2
sin
2
cos
12
1
sin1
2
sin
2
cos,
1
1
sin1
2
sin
2
cos
2
22
2
22
2
22
1
2
1
1
1
1
2
sin2
t
tn
t
tn
n
n
t
tn
n
n
2
22
1
2
1
2
1
1
sin2
1
sin
1
sin
t
tn
tn
tn
tn
tn
q.e.d.
77. 77
SOLO Wings in Compressible Flow
77
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic
Flow and Supersonic Leading Edge (tanΛ < β) (continue – 7)
2
22
1
22
22
1
222
1
cos
1
12
1
sin
21
12
1
1
2
t
tn
n
U
t
tn
n
U
u
n
U
x
Pu
P
1:,1
tan
:&tan:,: 2
Mn
x
y
t
xd
zd
S
For an Un-swept Wing (Flat Surface) (Λ = 0)
we have t = 0 & n = 0
SP
x
z
M
UU
x
Pu
122
2
2
22
1
22
22
1
2
2
2
1
cos
1
14
1
sin
2
1
1
1
2
2
t
tn
nt
tn
nU
Pu
PCp
78. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Semi-Infinite Triangular Wing in a
Supersonic Flow
Pressure Field for a Semi-Infinite Triangular Wing
with a Supersonic Leading Edge
Pressure Field for a Semi-Infinite Triangular Wing
with a Subsonic Leading Edge
Mach Line
Summary
79. SOLO Wings in Compressible Flow
Theoretic Solution for Pressure Distribution on a Delta Wing in a Supersonic Flow
Inclined Delta Wing with
Subsonic Leading Edge (0 < m < 1)
(a) Wing Planform (Triangular Wing)
(b) Pressure Distribution on a Section
Normal to the Flow Direction, m = 0.6.
22
2
1
2
1
cos
1 tm
t
m
m
22
2
1
2
1
cosh
1 tm
t
m
m
nMm /11tan
tan
tan
: 2
mtM
x
y
M
x
y
t 1tan
tan
1
1: 22
Inclined Delta Wing with
Supersonic Leading Edge ( m > 1)
(a) Wing Planform (Triangular Wing)
(b) Pressure Distribution on a Section
Normal to the Flow Direction, m = 1.5.
Summary
81. 81
Delta wing vortices
Delta wing pressure distribution (suction effect at the tip)
SOLO Wings in Compressible Flow
82. 82
(A)- Flow field in wing-tail plane, influence of angle of attack
SOLO Wings in Compressible Flow
83. 83
(B)- Flow field in wing-tail plane, influence of
control deflection for pitch
SOLO Wings in Compressible Flow
84. 84
(C)- Flow field in wing-tail plane, influence of
control deflection for roll
SOLO Wings in Compressible Flow
Return to Table of Content
85. 85
SOLO Wings in Compressible Flow
Arrowhead Wings with Double-Wedge Profile at Zero Incidence [after Puckett and Steward]
Nomenclature
86. 86
SOLO Wings in Compressible Flow
Arrowhead Wings with Double-Wedge Profile at Zero Incidence [after Puckett and Steward]
Thickness Drag for e = 0
87. 87
SOLO Wings in Compressible Flow
Arrowhead Wings with Double-Wedge Profile at Zero Incidence [after Puckett and Steward]
Thickness Drag for e = 0.5
88. 88
SOLO Wings in Compressible Flow
Arrowhead Wings with Double-Wedge Profile at Zero Incidence [after Puckett and Steward]
Thickness Drag for b = 0.2 Return to Table of Content
89. 89
SOLO Wings in Compressible Flow
Arrowhead Wings with constant Chord, Biconvex Profile, and Subsonic Leading Edge [after Jones]
(a) Platform
(b) Pressure Distribution at various Spanwise Stations
(c) Thickness Drag Coefficient at various Spanwise Stations
Return to Table of Content
90. 90
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwise plane [after Lawrence]
(a) Nomenclature and Geometrical Relationships. Note that e is negative if C lies aft of B.
91. 91
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
92. 92
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
93. 93
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
94. 94
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
95. 95
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
96. 96
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leading and
Trailing Edges and the same dimensionless profile in all chordwiseplane [after Lawrence]
(b), (c), (d), (e), (f) , (g) Wings with Double-Wedge Profiles
97. 97
SOLO Wings in Compressible Flow
Systematic Presentation of Wave Drag of Thin, Nonlifting Wings having Straight Leanding and
Trailing Edges and the same dimensionless profile in all chordwise plane [after Lawrence]
(h) Wings with Biconvex Parabolic Arc Profile Return to Table of Content
98. 98
SOLO Wings in Compressible Flow
λ – Taper Ratio, 12
M
CNα – Slope of the Normal Force Coefficient Computations of Swept Wings
ΛLE – Leading Edge Swept Angle
Wing Planform
S – Wing Area
2
1
2
1
2
b
c
b
c
c
c
b
ccS r
r
t
rtr
AR – Aspect Ratio
1
12
2
1
22
r
r
c
b
b
c
b
S
b
AR
LE
b
c tan
2
2/
0
2
322
2/
0 2
2
22
2/
0
2
2/3
1
2/
1
2/1
2
2/
1
2/
121
2/1
2
2
1
b
r
b
r
r
b
b
y
b
y
y
b
c
yd
b
y
b
y
c
bc
ydyc
S
c
2
0
2/
11
2/
b
y
b
y
c
b
y
cccyc rrtr
1
1
3
2
12
3
1
11
1
2
1
62
1
22/1
2 2
22 rrr ccbbb
b
c
2
22
1
14
1
1
3
2
b
Sc
c r
99. 99
SOLO Wings in Compressible Flow
λ – Taper Ratio,
12
M
ΛLE – Leading Edge Swept Angle
CNα – Slope Computation is done as follows:
1. Compute s = β/tan ΛLE.
If s<1 use the abscissa on the left side of the chart.
If s>1 use the right side of the chart with the
abscissa tanΛLE/β.
2. Pick the chart corresponding to the
Taper Ratio λ. If λ is between the values of
the given charts interpolation is needed.
3. Calculate AR tanΛLE for the given Airfoil.
This is the parameter in the charts. If λ is
between curves in the chart interpolation
is needed.
4. Read the corresponding value from the
ordinate; this value will correspond to
tanΛLE (CNα) if the left side of the chart is
used, and it will correspond to β(CNα)
if the right side of the charts is used.
5. Extract CNα by dividing the left ordinate
by tanΛLE , or by dividing the right ordinate
by β, as the case may be.
“USAF Stability and Control DATCOM Handbook” , Air Force
Flight Dynamics Lab. Wright-Patterson AFB, Ohio, 1965
CNα – Slope of the Normal Force Coefficient Computations of Swept Wings
100. 100
SOLO Wings in Compressible Flow
λ = 0 – Taper Ratio
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
101. 101
SOLO Wings in Compressible Flow
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
λ = 1/5 – Taper Ratio
102. 102
SOLO Wings in Compressible Flow
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
λ = 1/4 – Taper Ratio
103. 103
SOLO Wings in Compressible Flow
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
λ = 1/3 – Taper Ratio
104. 104
SOLO Wings in Compressible Flow
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
λ = 1/2 – Taper Ratio
105. 105
SOLO Wings in Compressible Flow
CNα – Slope of the Normal Force Coefficient, ΛLE – Leading Edge Swept Angle
λ – Taper Ratio, 12
M
λ = 1 – Taper Ratio
Return to Table of Content
106. 106
Comparison of Experiment and Theory for Lift-Curve Slope of Un-swept Wings
(after Vincenti, 1950)
SOLO Wings in Compressible Flow
107. 107
Comparison of Experiment and Theory for Lift-Curve Slope of Swept Wings
(after Vincenti, 1950)
SOLO Wings in Compressible Flow
109. 109
Thickness plus Skin-Friction Drag as a function of position of Maximum Thickness
(after Vincenti, 1950)
SOLO Wings in Compressible Flow
110. 110
Effect of radius of Subsonic Leading Edge on Pressure-Drag Ratio due to Lift
(after Vincenti, 1950)
SOLO Wings in Compressible Flow
111. 111
Effect of radius of Subsonic Leading Edge on Lift-to-Drag Ratio (after Vincenti, 1950)
SOLO Wings in Compressible Flow
112. 112
SOLO Wings in Compressible Flow
Lift Slope of Swept-Back Wings (taper λ = 1) at Supersonic Incident Flow,
0 < m <1; Subsonic Leading Edge; Supersonic Leading Edge.
113. 113
SOLO Wings in Compressible Flow
Drag Coefficient due to Lift versus Mach Number
for a Trapezoidal, a Swept-Back, and a Delta Wing
of Aspect Ratio Λ = 3.
Dashed curve: with suction force.
Solid curve: without suction force.
114. 114
SOLO Wings in Compressible Flow
Drag Coefficient (Wave Drag) at Zero Lift for Delta Wing (Triangular Wing) versus
Mach Number.
Profile I: Double Wedge profile.
Profile II: Parabolic Profile,
0 < m < 1: Subsonic Leading Edge,
m > 1: Supersonic Leading Edge
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115. 115
SOLO Wings in Compressible Flow
Drag Coefficient (Wave Drag) at Zero Lift of Swept-Back Wings (tape λ = 1) at
Supersonic Incident Flow.
0 < m < 1: Subsonic Leading Edge,
m > 1: Supersonic Leading Edge
116. 116
SOLO Wings in Compressible Flow
Lift Slope versus Mach Number for a
Trapezoidal, a Swept-Back, and a Delta Wing
of Aspect Ratio Λ = 3.
Total Drag Coefficient (Wave Drag +
Friction Drag) versus Mach Number for a
Trapezoidal, a Swept-Back, and a Delta Wing
of Aspect Ratio Λ = 3.
Double-Wedge profile t/c = 0.05, x c = 0.50
117. 117
Lifting Properties of Three Planforms
(after Jones, 1946)
SOLO Wings in Compressible Flow
Induced Drag of Three Planforms
(after Jones, 1946)
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119. 119
centre stickailerons
elevators
rudder
Aircraft Flight Control
Generally, the primary cockpit flight controls are arranged as follows:
a control yoke (also known as a control column), centre stick or side-stick (the
latter two also colloquially known as a control or B joystick), governs the
aircraft's roll and pitch by moving the A ailerons (or activating wing warping
on some very early aircraft designs) when turned or deflected left and right,
and moves the C elevators when moved backwards or forwards
rudder pedals, or the earlier, pre-1919 "rudder bar", to control yaw, which move
the D rudder; left foot forward will move the rudder left for instance.
throttle controls to control engine speed or thrust for powered aircraft.
SOLO
131. 131
The effect of left rudder pressure Four common types of flaps
Leading edge high lift devices
The stabilator is a one-piece horizontal tail surface that
pivots up and down about a central hinge point.
Aircraft Flight ControlSOLO
134. SOLO
134
Aircraft Flight Control
Specific Stabilizer/Tail Configurations
Tailplane
Fuselage mounted Cruciform T-tail Flying tailplane
The tailplane comprises the tail-mounted fixed horizontal stabiliser and movable elevator.
Besides its planform, it is characterised by:
• Number of tailplanes - from 0 (tailless or canard) to 3 (Roe triplane)
• Location of tailplane - mounted high, mid or low on the fuselage, fin or tail
booms.
• Fixed stabilizer and movable elevator surfaces, or a single combined stabilator or
(all) flying tail.[1] (General Dynamics F-111)
Some locations have been given special names:
• Cruciform: mid-mounted on the fin (Hawker Sea Hawk, Sud Aviation Caravelle)
• T-tail: high-mounted on the fin (Gloster Javelin, Boeing 727)
Sud Aviation Caravelle
Gloster Javelin
135. SOLO
135
Aircraft Flight Control
Specific Stabilizer/Tail Configurations
Tailplane
Some locations have been given special names:
• V-tail: (sometimes called a Butterfly tail)
• Twin tail: specific type of vertical stabilizer arrangement found on the empennage of
some aircraft.
• Twin-boom tail: has two longitudinal booms fixed to the main wing on either side of
the center line.
The V-tail of a Belgian Air
Force Fouga Magister
de Havilland Vampire
T11, Twin-Boom Tail
A twin-tailed B-25 Mitchell