曝光台 注意防骗
网曝天猫店富美金盛家居专营店坑蒙拐骗欺诈消费者
ao,c - -- (1.59)
M~
where Cp is the pressure coefficient and ao is the lift-curve slope of the airfoil.
The suffixes r and c denote incompressible and compressible values, respectively.
It may be noted that the pressure coefficient is defined as
p - Pea
Cp = 1 2p.tl-2 (1.60)
According to Eqs. (1.58-1,60), which are based on potential fiow theory, the mag-
nitude of the pressure coe:fficient at any point on the airfoil surface and the section
lift-curve slope increases steadily with Mach number above their incompressible
values. These formulas are applicable as long as the flow ever)rwhere is subsonic
and shock free.
REVIEW OF BASIC AERODYNAMIC PRINCIPLES 41
~
M~ < Mcr
-
M =Mcr
--
M~ > Mcr
a) Subsoruc flow
<;==::: A--J(-:
b) Local soiuc flow
c) Supersonic flow on upper surface
-
M >Mcr
Wake
d) Supersonic flow on upper and lower surfaces
e) Airfoilin supersonic ftow
Fig.1.42 Flow over airfoil atlugh speeds.
When the local Mach number at point P on the airfoil surface reaches the value
of uruty, the corresponding freestream Mach number is called the critical Mach
number and is denoted by Mu as schematically illustrated in Fig. 1.42b. Elsewhere
on the surface of the airfoil, the local Mach number is below unity, and the flow
is subsonic. The value of Mcr depends on the geometrical shape of the body and
the angle of attack. For a given airfoil, the critical Mach number usually decreases
with increase in angle of attack.
The typical values of Mcr for airfoils at zero-lift lie in the range 0.6-0.85. For
example, at Ci = 0, the values of Mcr for NACA 2412, NACA 23012, and NACA
653-418 airfoils are, respectively, equal to 0.69, 0.672, and 0.656. For a thin airfoil
;:=::: -.
i:
':
h
:7
'~
1'
'4
r.
.'.:
'l'
. .:
k
:
-
."i
.i
.
t
o
,..
',
.i
.
..:
.-.
':
. ':;
. ':
.ri
. ';
' '.:.
o.
:t
j
:,
i
.:/
;}
..t
:i
-.
./
1
.=
..
.{
'/!
j
.\
:.
'l.
}
-
. ;.
':
: ::;
':.
,i'J.
. .,
, .\.
:<
.;!
'. i
t/
}:
..
t;
: ;-..
:..^:
.::,
}
':
t-
':.
,.-.:
::::
:t
::.
r~
I3
~
42 PERFORMANCE, STABILITY DYNAMICS,AND CONTROL
like NACA 0006, the a:itical Mach numberif 0.805 at Ci - 0. The values of Mc.r
for other airfoils may be found elsewhere.3,11
As the freestream Mach number increases beyond Mcr, the local velocity on the
surface of the airfoil exceeds the sonic velocity at more than one point In fact, a
small region appears where thelocal Mach number is either equal to or greater than
unity as shown in Fig. 1.42c. Because the flow Mach num~er behind the airfoil
has to be equal to the freestream value, which is subsonic, the region of supersonic
中国航空网 www.aero.cn
航空翻译 www.aviation.cn
本文链接地址：
PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL1(26)
