Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Tined or irregular periphery
Reexamination Certificate
1999-09-24
2001-02-20
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Tined or irregular periphery
C416SDIG002
Reexamination Certificate
active
06190132
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a rotor blade for rotorcraft such as a helicopter.
2. Description of the Related Art
FIGS. 13A and 13B
are views showing the aerodynamic environment of a helicopter rotor in the forward flight case. As shown in
FIG. 13A
, when a helicopter
1
having rotor blades of a rotor radius R and rotating at an angular speed &OHgr; advances at a ground velocity V, an advancing blade in which the ground velocity V is added to a rotor speed &OHgr;R, and a retreating blade in which the ground velocity V is subtracted from the rotor speed &OHgr;R are largely different in airspeed from each other.
At the azimuth angle &PSgr; (an angle measured counterclockwise from the rearward direction of the helicopter
1
) of 90 degrees, the airspeed of the advancing blade reaches the maximum, and the airspeed of the tip of the advancing blade is &OHgr;R+V. At the position of &PSgr;=270 degrees, on the other hand, the airspeed of the retreating blade reaches the minimum, and the airspeed of the tip of the retreating blade is &OHgr;R−V. The airspeed at an intermediate portion of the blade has a value which is obtained by proportionally distributing &OHgr;R+V and &OHgr;R−V. Assuming that &OHgr;R=795 km/h and V=278 km/h, for example, the airspeed at the position of about 35% from the root end of the retreating blade is zero as shown in FIG.
13
A.
When a helicopter flies at high speed, particularly, the airspeed at the tip of the advancing blade is transonic and a strong shock wave is generated. In a drag divergence region shown as a hatched portion in
FIG. 13B
, this strong shock wave causes an abrupt increase of drag which acts on the blade. Noise which is generated by such a strong shock wave is called high-speed impulsive noise. At this time, in a coordinate system as seen from a rotating rotor blade, a phenomenon which is called delocalization of supersonic region occurs. The generated shock wave propagates to a distant place through a delocalization supersonic region. As a result, high level of noise is observed in the distant place.
In the retreating blade, since the airspeed thereof is significantly lowered, angle of attack &agr; of the blade must be increased in order to obtain a lift which is equivalent to that of the advancing blade. For this purpose, it is common to carry out pitch control in which a pitch angle of the blade is controlled in accordance with azimuth angle &PSgr;. The pitch angle of the blade is controlled using a sinusoidal wave which has a minimum amplitude at &PSgr;=90 degrees and a maximum at &PSgr;=270 degrees. At this time, as shown in
FIG. 13B
, the angle of attack a of the blade is changed in a span direction by flapping movement of the blade itself. In the case of &PSgr;=90 degrees, for example, the angle of attack a of the blade is about 0 degree at the root, and about 4 degrees at the tip. In the case of &PSgr;=270 degrees, the angle of attack &agr; of the blade is about 0 degrees at the root, and about 16 to 18 degrees at the tip, and exceeds the stall angle of attack. When the angle of attack &agr; exceeds the stall angle of attack, large changes of lift coefficient Cl and pitching moment coefficient Cm suddenly occur, resulting in that a large vibration of the airframe and fatigue loads in pitch links are generated.
In this way, the high-speed impulsive noise is used as an evaluation item for the advancing blade, and the maximum lift coefficient Clmax and the stall angle of attack are used as evaluation items for the retreating blade. The maximum lift coefficient Clmax is defined by the maximum value of the lift coefficient Cl when the angle of attack &agr; of a blade having a predetermined aerofoil is gradually increased to reach the stall angle of attack. Usually, as the high-speed impulsive noise and absolute value of the pitching moment coefficient Cm are smaller, or as the maximum lift coefficient Clmax and the stall angle of attack are larger, the blade is judged to be more excellent.
In order to improve the performance of the high-speed flight and to reduce the high-speed impulsive noise, a thin airfoil section may be employed in a blade tip portion. In this method, however, the stall angle of attack is small and also the maximum lift coefficient is small. Therefore, this method is not appropriate. For the above-mentioned purpose, another method in which a simple swept-back angle is formed in the blade tip portion as shown in
FIG. 14
is usually employed. A shape provided with the simple swept-back angle is obtained by sweeping back the blade tip portion by a constant swept-back angle. When the simple swept-back angle is formed in the blade tip portion, however, aerodynamic center of the blade tip portion is largely shifted in a rearward direction as shown in FIG.
14
. At a position where the aerodynamic center is rearward shifted from a pitch axis by a length &Dgr;X, a moment M about the pitch axis has a value which is obtained by multiplying a lift L with the length &Dgr;X. In the case where the simple swept-back angle is formed as described above, the pitching moment in the direction of the head-down is increased, with the result that control performance is lowered.
SUMMARY OF THE INVENTION
It is an object of the invention to provide a rotor blade for a rotorcraft in which, in combination with a high-performance third-generation airfoil for a helicopter of high maximum lift coefficient Clmax and drag divergence Mach number M
dd
(for example, Japanese Patent Application No. 11-45196(1999)), improvement of performance in high-speed flying, reduction of high-speed impulsive noise, less increase of pitching moment and high control performance are attained.
The invention provides a rotor blade for rotorcraft, comprising:
a blade root portion
10
which is to be attached to a rotor head for rotating;
a center portion
11
which linearly elongates from the blade root portion
10
; and
a blade tip portion
12
which has a shape and a predetermined aerofoil, the shape elongating outward from the center portion
11
and being defined by a leading edge
23
, a side edge
25
, and a trailing edge
30
,
wherein in the leading edge
23
of the blade tip portion
12
, a swept-back angle &lgr;(r) at a distance r from a rotation center of a rotor satisfies the following relational expression (1):
cos
-
1
M
dd
M
∞
+
M
TIP
r
R
≦
λ
(
r
)
≦
cos
-
1
0.985
×
M
dd
M
∞
+
M
TIP
r
R
(
1
)
where R is a rotor blade length of the rotorcraft, M
∞
is a maximum flight Mach number which is a flight limit speed of the rotorcraft, M
TIP
is a blade tip Mach number which is a tip speed during hovering, and M
dd
is a drag divergence Mach number determined from the aerofoil of the blade tip portion
12
.
FIG. 11
is a graph showing the drag divergence Mach number M
dd
. The abscissa of the graph indicates a flight Mach number M of a certain rotorcraft which is normalized by a normalization constant M
cr
, and the ordinate indicates a drag coefficient C
d
which is normalized by a normalization constant C
dinc
. The graph shows that, when the flight Mach number M is small, the drag coefficient C
d
is substantially constant, and, when the flight Mach number M exceeds a certain value, the drag coefficient C
d
is suddenly increased. When the inflow Mach number to the blade leading edge is increased, a supersonic region appears on a blade surface, and a shock wave is formed, with the result that drag is rapidly increased and loud noise due to the shock wave is generated. These phenomena are caused due to compressibility of the air. As an index indicating the influence of compressibility is used the drag divergence Mach number M
dd
. The drag divergence Mach number M
dd
is a value peculiar to the airfoil, and defined as the flight Mach number M in the case where dC
d
/dM is 0.1.
FIG. 12
is a view showing an effective Mach number at the leading edge
Murashige Atsushi
Tsujiuchi Tomoka
Yamakawa Eiichi
Advanced Technology Institute of Commuter-Helicopter, Ltd.
Look Edward K.
McAleenan James M
Wenderoth , Lind & Ponack, L.L.P.
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