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
and Rocket Propulsion
Gordon, C. Oates 1984
Re-Entry Vehicle Dynamics
Frank J. Regan 1984
642 PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
Fig.T5 Schematic variations of airfoillift and drag coeffiaents with angle of attack.
resistant to autorotation. On the other hand, cross sections with fiat bottoms are
prone to autorotation.
To understand the aerodynamics associated with the autorotation ofthe ffiselage,
consider a noncircular cylindrical model mounted on a single-degree-of-freedom,
free-to-roll apparatus and held at an angle of attack in an airstream of uniform
velocity Uo as shown in Fig. 7.6. We assume that the modelis pivoted atits center
of gravity and is constrained to rotate about the velocity vector Uo. Let the origin
of the body-fixed coordinate system coincide with the center of gravity.
Suppose that the modelis ir;;parted a disturbance that makes it rotate in a clock-
wise direction (viewed from a~"end) with an angular velocity S2. The crossflow
angle 4 at an axial location x is given by
o = t--'(szv7)
(7.69)
We note that, for the cross sections ahead of the center of gravity, the crossflow
angle is positrve and, for the sections aft of the center of gravity, the crossflow
angle is negative.
The local dynamic pressure at the axial location x is given by
q, = -,pUo' [,+ (
S2x sin CL
)2]
(7.70)
For simplicity,-we assume that the side force coefficient Cy of the fuselage cross
sections depends only on the crossflow angle 4.
INERTIA COUPLING AND SPIN
C;
+>o
Section AA
Section BB
Cy
Fig. 7.6 Schematic diagram ofan autorotating fuselage.
643
rfhe moment developed by the section AA about the axis of rotation (velocity
vector) is given by
z\Nn = -,pU- [,+ (
or, in coefficient form,
where
ACfZ = -,, [,+(
C2x sin a
Uo
S2x sin ct
ACcz
)2lCy()bOxsinadx (7.71)
)2],y()xsiDCL-cIX (7.72)
ANrz
21p U02b0/2
(7.73)
Here, bo is the width of the body, and I is the total length of the body.
Given the variation of Cy with crossflow angle 4, an integration of Eq. (7.72)
gives the variation of the moment coefficient Cn as a function of the angular
velocity CZ. The schematic van:ation of Csz is shown in Fig. 7.7. It is possible that
this variation is of two types. For type I variation, the yawing moment coefficient
is initially positive and at some point crosses zero and becomes negative for higher
644 PERFORMANCE, STABILITY, DYNAMICS, AND CONTROL
+
0
┏━━━━━━━━━━━━━━━━━━━━━━━━━━━━━┓
┃ Auto- ┃
┃ ┃
┃[-~-.. ┃
┣━━━━━━━━━━━━━━━━━━━━━━━━━━━━━┫
┃\\ \Dm: - ┃
┗━━━━━━━━━━━━━━━━━━━━━━━━━━━━━┛
Fig. 7.7 Schematic variation of yawing moment coefficient of a rotating fuselage.
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