F18 – Lecture Notes
1. Prandtl-Meyer Waves
2. Shock-Expansion Theory (Supersonic Airfoils)
Reading: Anderson 9.6, 9.7
Prandtl-Meyer Waves
Wave flow relations
An expansion fan, sometimes also called a Prandtl-Meyer expansion wave, can be considered
as a continuous sequence of infinitesimal Mach expansion waves. To analyze this continuous
change, we will now consider the flow angle θ to be a flowfield variable, like M or V .
Across each Mach wave of the fan, the flow direction changes by dθ, while the speed changes
by dV . Oblique-shock analysis dictates that only the normal velocity component u can
change across any wave, so that dV must be entirely due to the normal-velocity change du.
�
1
1
2
2
1 2
M M
V
V
dV
du
u
V
u
µ
V
� �
V
dV
du
µ
Mach wave
d
�
dV
µ
tan
From the u-V and du-dV velocity triangles, it is evident that dθ and dV are related by
dV 1
dθ =
tan µ V
assuming dθ is a small angle. With sin µ = 1/M , we have
1 cos µ
1
− sin
2
µ
1
− 1/M
2
√
M
2
− 1= = = =
tan µ sin µ sin µ 1/M
and so the flow relation above becomes
dV
dθ =
√
M
2
− 1 (1)
V
This is a differential equation which relates a change dθ in the flow angle to a change dV in
the flow speed throughout the expansion fan.
Prandtl-Meyer Function
The differential equation (1) can be integrated if we first express V in terms of M .
V = M a = M a
o
1 +
γ −1
M
2
−1/2
2
1
ln V = ln M + ln a
o
−
1
2
ln
1 +
γ −1
2
M
2
dV
V
=
dM
M
−
1
2
1 +
γ −1
2
M
2
−1
γ −1
2
2M dM
dV
1
dM
=
V
1 +
γ−1
2
M
2
M
Equation (1) then becomes
dM
dθ =
√
M
2
− 1
(2)
1 +
γ−1
M
2
M
2
which can now be integrated from any point 1 to any point 2 in the Prandtl-Meyer wave.
θ
2
M
2
√
M
2
− 1
dM
dθ =
θ
1
M
1
1 +
γ−1
M
2
M
2
θ
2
− θ
1
= ν(M
2
) − ν(M
1
) (3)
γ +1
where ν(M ) ≡
γ −1
arctan
γ −1
(M
2
− 1) − arctan
√
M
2
− 1 (4)
γ +1
Here, ν(M ) is called the Prandtl-Meyer function, and is shown plotted for γ = 1.4 .
v(M) deg.
50
45
40
35
30
25
20
15
10
5
0
1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0
M
Equation (3) can be applied to any two points within an expansion fan, but the most common
use is to relate the two flow conditions before and after the fan. Reverting back to our
previous notation where θ is the total turning of the corner, the relation between θ and the
upstream and downstream Mach number is
θ = ν(M
2
) − ν(M
1
) (5)
2
ν(
)
θ
�
�
M
θ
1
M
M
2
M
M
1
M
2
This can be considered an implicit definition of M
2
(M
1
, θ), which can be evaluated graphi-
cally using the ν(M) function plot, as shown in the figure.
Shock-Expansion Theory
The combination of oblique-shock relations and Prandtl-Meyer wave relations constitutes
Shock-Expansion Theory , which can be used to determine the flow properties and forces
about simple 2-D shapes in supersonic flow.
Flat-plate supersonic airfoil
A flat plate is the simplest supersonic airfoil. When set at an angle of attack α, the leading
edge point effectively is a convex corner to the upper surface flow, with turning angle θ = α.
The upper flow then passes through the resulting Prandtl-Meyer expansion, which increases
its Mach number from M
1
= M
�
, to a larger value M
2
= M
u
. The latter is implicitly
determined via the Prandtl-Meyer relation (5).
α = ν(M
u
) − ν(M
�
) → M
u
(M , α)
The corresponding upper-surface pressure is then given by the isentropic relation.
1 +
γ−1
M
2
γ/(γ−1)
2
�
p
u
= p
1 +
γ−1
M
2
u
2
Conversely, the leading edge is a concave corner to the bottom surface flow, which then sees
a pressure rise through the resulting oblique shock. The lower-surface Mach number M
and
pressure p
are obtained from oblique-shock relations, with M
1
= M
�
, θ = α as the inputs.
M
p
p
u
p
l
α
L’
D’
R’
c
The pressure difference produces a resultant force/span R
�
acting normal to the plate, which
can be resolved into lift and drag components.
R
�
= (p
− p
u
) c
L
�
= (p
− p
u
) c cos α
D
�
= (p
− p
u
) c sin α
3
It’s worthwhile to note that this supersonic airfoil has a nonzero drag even though the flow is
being assumed inviscid. Hence, d’Alembert’s Paradox definitely does not apply to supersonic
2-D flow. The drag D
�
computed here is associated with the viscous dissipation occurring
in the oblique shock waves, and hence is called wave drag . This wave drag is an additional
drag component in supersonic flow, and must be added to the usual viscous friction drag,
and also the induced drag in 3-D cases.
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