曝光臺 注意防騙 網曝天貓店富美金盛家居專營店坑蒙拐騙欺詐消費者

┃                                  ┃
┃-::".-,2--:  -::::--l---'?- 'L    ┃
┣━━━━━━━━━━━━━━━━━┫
┃       Aile                       ┃
┗━━━━━━━━━━━━━━━━━┛
a) Rectangular wing
Aileron
~g.1.53   StaD progression on rectangular and swept-back wings.
     Another consequence of the stall progressiort from tips to the root is the pitch-
up tendency exhibited by the swept-back wings. This pitch up occurs because tip
regions that have the largest moment arm lose lift. The only contribution comes
from unstalled sections closer to the root that have a smaller moment arm. As a
result, the stabilizing contribution of the root sections decreases and the swept
wing experiences the pitch-up as shown schematically in Fig. 1.56.
   One way of controlling the tip stall of swept-back wings is the application
of boundary-layer fences or vortex generators.14 Boundary-Iayer fences are obsta-
cles positioned at various spanwise stations to prevent the spanwise flow within
the boundary layer. The height of these fences is sufficiently small so that they do
not disturb the external flow. Vortex generators are devices that generate stream-
wise vorticity to energize the boundary layer and make it more resistant to fiow
separation. Vortex generators may be either submerged within the boundary layer
or may protrude outside of it_ The vort,ex generators that protrude outside'~he
boundary layer may produce considerable skin friction, whereas the submerged
type alleviate this problem because they are not exposed to the external airstream.
14
x
W
{4
"~z.
i.:/: :
:':t.
'J.
z:9
+5
;v
'c;
it
--.
   : .'!
 :.
':.s
.L:
54                 PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
Fig. 1.54    Spanwise lift distribution on straight and swept wings.
      De/ta wings.    In the late 1950s or the early 1960s, the use of sweep-back at
 supersonic speeds, wherein the wing leading edge was swept behind the Mach
 cone, provided an efficient method of reducing the wave drag. However, at low
 speeds, the increased sweep caused difficulties in maintajning attached flow even
 at moderate angles of attack. Flow separations occurred at quite modest values of
 angles of attack and tended to spread in an unpredictable way. This often caused
  serious stability and control problems. The most troublesome of these separations
  originated at the leading edge and rolled up into a vortex sheet. The point of origin
 of these leading-edge separations were difficult to tie down beca~se the swept-
 back wing had a round leading edge. Designers tried various fixes to alleviate the
-problem but not with much success. This led aerodynamicists to believe that the
 requirements for high  and low-speed fiights were in direct conflict.
    The,ecmergence of the thin slender, sharp-edged delta wirig solved this prob-
lem.15,16 The fiow over a sharp-edged delta wing differs significantly from that
 over a swept-back wing with round leading edge. The most important difference
  is that the flow separation occurs right at the sharp leading edges and, in this way,
 flow separation points are fixed for all angles of attack. With fixed flow separation
 points, the problems associated with unpredictable stall progression of the swept-
 back wings were eliminated. Along with this, many of the stability and control
 problems associated with the uncertain flow separation pattern of thve swept-back
 wing also disappeared.
        The separated fiow on thelee side rolls up to form a spiral vortex over the lee side
  of the wing as shown in Fig.  1.57a  The pressure in the vortex core is considerably
  lower and, as a result, substantial Jift increment is'obtained as shown in Fig.  1.57b:
 This incremental lift is called vortex lift and is associated with the large mass
REVIEW OF BASIC AERODYNAMIC PRINCIPLES                 55
　
中國航空網 www.k6050.com
航空翻譯 www.aviation.cn
本文鏈接地址：PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL1(32)