Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Radial flow devices
Reexamination Certificate
2000-05-19
2002-04-02
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Radial flow devices
C416S228000, C416SDIG002, C416SDIG005
Reexamination Certificate
active
06364615
ABSTRACT:
The present invention relates to a blade for the rotary wings of an aircraft, in particular a helicopter, said blade making it possible simultaneously to reduce noise and improve the performance at high load, especially at takeoff and at moderate flight speed.
It is known that, both in hovering flight and in forward flight, the performance of a rotor of a rotary-wing aircraft, especially a helicopter, is limited by the following phenomena:
the shockwaves which develop on the suction face of the advancing blades during high-speed flights;
the stalling that results from the detachment of the boundary layer on the suction face of the retreating blades when there is a demand for lift in translational flight;
the interaction of the vortex generated by the previous blade with the following blade, which leads, during hovering flight, to a substantial dissipation of energy in two forms: induced power and profile drag power.
In addition to being responsible for loss of performance, the shocks and the blade-vortex interaction are also responsible for acoustic problems in the form of pulsating noise, caused by shock delocalization (high-speed flight) and by pulsating changes in lift when the marginal vortex directly strikes the blade (decent), respectively.
It has been found that the performance of a blade for the rotary wings of an aircraft depends, to a large extent, on parameters associated with the construction of the blade, such as:
a) the radial distribution of the blade area;
b) the sweepback of the blade tip;
c) the change in relative thickness of the profiles;
d) the distribution of the twist of the profiles; and
e) the blade tip droop.
The influence that the three main parameters a), b) and c) have on the performance and on the noise of a rotary-wing blade is explained in detail below.
a) Radial Distribution of the Blade Area
For a rotor of a rotary-wing aircraft whose elementary profiles or sections all work with the same coefficient of lift Cz, the linear lift varies as the local chord length L(r) and as the square of the local speed, which is directly proportional to the radius (radial position)
r
of the section. This means that the total lift of the blade varies proportionally with the mean chord {overscore (L)} defined by a square-law weighting of the radius
r
:
L
_
=
∫
R0
R
L
(
r
)
r
2
ⅆ
r
∫
R0
R
r
2
ⅆ
r
in which RO represents the radius
r
at the start of the blade section at the root end of the blade and R is the total radius of the blade.
It is common practice for the performance of blades of various shapes to be compared, by referring them to this mean chord {overscore (L)}.
Compared with a conventional blade of rectangular shape, calculations show, and experience confirms this, that a reduction in the chord at the outboard end of the blade (tapered shape) improves performance over a wide speed range, including in hovering flight. In translational flight, this improvement is explained essentially by the reduction in the drag of the profiles, which is due to the reduction in the profile drag due to the reduction in the chord at the tip. Shocks in this region are exerted on a smaller area while the central part of the blade, not subjected to the shocks, provides most of the lift with a maximum efficiency: the lift/drag ratio here is a maximum. At low speed and in hovering flight, the tapering of the tip improves the lift efficiency-that is to say makes it possible to reduce the power needed for a given lift force-according to a different mechanism: a more homogeneous induced velocity distribution is obtained over the entire rotor disk, preventing the load from being too concentrated at the tip. The distribution thus approaches the optimum distribution for the lift efficiency, which consists of a uniform induced velocity over the entire disk.
A second known advantage provided by the outboard taper of the blade is a certain reduction in the noise. On the one hand, the volume of air displaced by the high-speed passage of the tip is reduced as the square of the chord (for the same relative thickness of the profiles). This results in a reduction in some of the noise still present, that corresponding to the noise called “monopole source”. On the other hand, the blade edge vortex, at the origin of the blade-vortex interaction noise, curls more slowly and the maximum velocity in the core of this vortex is lower the greater the distance between the maximum chord zone and the tip. This results in appreciable attenuation of the interaction noise, particularly during the decent phases of the aircraft.
However, the outboard tapering of the blade has the drawback of requiring an increase in the chord over the rest of the span, so as to maintain the constant mean chord {overscore (L)} and so as not to excessively increase the coefficient of lift Cz of the profiles. This increase in the chord may be significant because of the r
2
weighting in the expression for {overscore (L)} (see above) and this results in the rotor being somewhat heavier. Nevertheless, the tapering on the blade tip side over a moderate length, of the order of 5 to 6% of the rotor radius, is a means commonly employed for improving the performance of the latter, generally in combination with sweepback of the blade tip, as illustrated in Patents FR-2,755,941, FR-2,689,852 and FR-2,617,118.
The tapering of the chord on the inboard side of the blade, that is to say on the side where it is attached to the hub, is a known means of limiting the drawback of an increase in mass and of improving the performance at high speed, that is to say above 300 km/h, since, under these conditions, this zone of the blade contributes little to the lift and greatly to the power consumed by the rotor (see Patents FR-2,755,941 and FR-2,689,852). However, this arrangement proves to be unfavorable in the case of the performance at moderate speed and in hovering flight since it tends to excessively reduce the load in the central zone of the rotor and to make the induced velocity distribution less uniform, thereby resulting in a reduction in the lift efficiency.
b) Offset of the Profiles in the Plane of the Rotor, with Part of the Blade Swept Back
In addition, in order to push back the threshold at which shockwaves appear and to limit their intensity, it is advantageous for the blade tip to be curved nearward (Patents FR-2,755,941 and FR-2,689,852 and Patent Application FR-97/16227) or else for it to have a double curvature, alternately toward the front and the rear (Patent FR-97/11230). The sweep angle &Lgr;, defined by the line of aerodynamic centers (approximately at the front quarter of the chord) and the feathering axis, reduces the effective Mach number and thus sweeping back the blade tip constitutes an effective means of reducing the unfavorable consequences of the compressibility of air, especially the appearance of shockwaves.
However, it is new known (Patent Application FR-97/16227) that the sweep angle must remain modest, typically less than 35°, so as to avoid the formation of a three-dimensional ram's horn vortex—or apex vortex—similar to that observed on delta-type wings. This is because this type of very stable and concentrated vortex produces intense interactions with the following blades and therefore contributes to the noise of a helicopter as it descends.
Furthermore, the magnitude of the offset with respect to the feathering axis, and the span length of the zone in question, also limited by the torsional forces which result from the offset of the aerodynamic lift as well as of the center of gravity. A known means for limiting this unfavorable effect consists in shifting the profiles of the mid-part forward and those of the tip rearward in such a way that the blade remains balanced overall: see, in particular, Patents FR-2,755,941, FR-2,689,852 and FR-97/11230.
c) Change in the Relative Thickness of the Profiles
The relative thickness of a blade cross section is defined as the ratio of the absolute thickness
e
with respect to the chord length L of the profile
Rebut Mathieu
Stahl Peter Rudolf
Toulmay François Valentin Gaston
Eurocopter
Look Edward K.
Nguyen Ninh
Stevens Davis Miller & Mosher LLP
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