U.S. patent application number 14/954362 was filed with the patent office on 2016-05-05 for airfoil for rotor blade with reduced pitching moment.
The applicant listed for this patent is Sikorsky Aircraft Corporation. Invention is credited to Ashish Bagai.
Application Number | 20160122011 14/954362 |
Document ID | / |
Family ID | 48572141 |
Filed Date | 2016-05-05 |
United States Patent
Application |
20160122011 |
Kind Code |
A1 |
Bagai; Ashish |
May 5, 2016 |
AIRFOIL FOR ROTOR BLADE WITH REDUCED PITCHING MOMENT
Abstract
A rotor blade for a rotary wing aircraft includes a root region
extending from a rotor head to about 15% to 20% of a blade radius,
a main region extending from a radial extent of the root region to
about 80% to 95% of the blade radius, a tip region extending from a
radial extent of the main region to a blade tip, and a trailing
edge. A portion of one of the root region, the main region and the
tip region has an airfoil profile section including a trailing edge
reflex camber relative to a base airfoil shape of a remainder of
the one of the root region, the main region, and the tip region to
reduce pitching moment of the rotor blade while maintaining a
positive aerodynamic characteristic of the base airfoil shape. The
reflex camber extends 20% of a blade chord from the trailing
edge.
Inventors: |
Bagai; Ashish; (Vienna,
VA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Sikorsky Aircraft Corporation |
Stratford |
CT |
US |
|
|
Family ID: |
48572141 |
Appl. No.: |
14/954362 |
Filed: |
November 30, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
13315767 |
Dec 9, 2011 |
9284050 |
|
|
14954362 |
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Current U.S.
Class: |
416/242 |
Current CPC
Class: |
B64C 27/467 20130101;
B64C 27/463 20130101; Y02T 50/10 20130101; B64C 2003/147 20130101;
Y02T 50/12 20130101; B64C 27/473 20130101 |
International
Class: |
B64C 27/467 20060101
B64C027/467; B64C 27/46 20060101 B64C027/46; B64C 27/473 20060101
B64C027/473 |
Claims
1. A rotor blade for a rotary wing aircraft comprising: a root
region extending from a rotor head to about 15% to 20% of a blade
radius; a main region extending from a radial extent of the root
region to about 80% to 95% of the blade radius; a tip region
extending from a radial extent of the main region to a blade tip;
and a trailing edge extending from at least the main region toward
the tip region, wherein a portion of one of the root region, the
main region and the tip region has an airfoil profile section
including a trailing edge reflex camber relative to a base airfoil
shape of a remainder of the one of the root region, the main
region, and the tip region to reduce pitching moment of the rotor
blade while maintaining a positive aerodynamic characteristic of
the base airfoil shape, the reflex camber extending 20% of a blade
chord from the trailing edge.
2. The rotor blade according to claim 1, wherein the tip section is
defined by an outboard 20% of the rotor blade span.
3. The rotor blade according to claim 2, wherein the airfoil
profile section is disposed at least partially at an outermost 5%
to 7% of the rotor blade span.
4. The rotor blade according to claim 1, wherein the airfoil
profile section is at least partially disposed at the root
region.
5. The rotor blade according to claim 1, wherein the rotor blade is
a main rotor blade of a helicopter.
6. The rotor blade according to claim 1, wherein the reflex camber
extends from the trailing edge upward by a distance of about 6% of
the blade chord.
7. The rotor blade according to claim 1, wherein at least a portion
of one of the root region, the main region and the tip region
including an airfoil profile section defined by a scaled set of
coordinates in which a set of x/c coordinates defined as a ratio of
an x coordinate to a blade chord length, and y/c coordinates
defined as a ratio of a y coordinate to the blade chord length,
listed in Table I are scaled by a selected factor.
8. The rotor blade according to claim 7, wherein the tip section is
defined by an outboard 20% of the rotor blade span.
9. The rotor blade according to claim 7, wherein the airfoil
profile section is at least partially disposed at the root
region.
10. The rotor blade according to claim 7, wherein the rotor blade
is a main rotor blade of a helicopter.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation of U.S. application Ser.
No. 13/315,767 filed Dec. 9, 2011, the disclosure of which is
incorporated by reference herein in its entirety.
BACKGROUND
[0002] The subject matter disclosed herein relates to rotary-winged
aircraft. More specifically, the subject disclosure relates to an
airfoil section for at least partial use on a rotor blade of a
helicopter.
[0003] Conventional rotary-wing aircraft have a forward airspeed
limited by a number of factors. Among these is the tendency of the
retreating blade to stall at high forward airspeeds. As the forward
airspeed increases, the airflow velocity across the retreating
blade slows such that the blade may approach a stall condition. In
contrast, the airflow velocity across the advancing blade increases
with increasing forward speed producing high lift, but increasingly
higher drag that results in higher rotor power requirements.
Forward movement of the aircraft thereby generates a dissymmetry of
lift between the advancing and retreating sides of the rotor. This
dissymmetry may create an unstable condition if lift is not
equalized across the advancing and retreating sides of the rotor.
An important approach in alleviating this dissymmetry is to use
airfoils that are capable of producing high lift at high pitch
angles and low relative velocities over the retreating side, while
minimizing the increase in drag over the advancing side. However,
designing such airfoils results in conflicting requirements as
governed by the physics of the problem. That is, designing an
airfoil that is capable of producing high lift at low speeds and
low drag at high speeds typically results in the manifestation of
some other undesirable characteristics, such as pitching moments
that exceed the structural-dynamic tolerance of the rotor blades
and control hardware.
[0004] Many airfoil sections have been developed, for example, as
in US Patent Appl. Pub. 2007/0187549, that when applied to a main
rotor of a helicopter, alleviate the unstable condition by
addressing the lift and drag effects on the blade. These blades,
however, exhibit high pitching moments which are detrimental to
rotor blade dynamic characteristics.
SUMMARY
[0005] Disclosed is a rotor blade for a rotary wing aircraft
includes a root region extending from a rotor head to about 15% to
20% of a blade radius, a main region extending from a radial extent
of the root region to about 80% to 95% of the blade radius, a tip
region extending from a radial extent of the main region to a blade
tip, and a trailing edge extending from at least the main region
toward the tip region. A portion of one of the root region, the
main region and the tip region has an airfoil profile section
including a trailing edge reflex camber relative to a base airfoil
shape of a remainder of the one of the root region, the main
region, and the tip region to reduce pitching moment of the rotor
blade while maintaining a positive aerodynamic characteristic of
the base airfoil shape. The reflex camber extending 20% of a blade
chord from the trailing edge.
[0006] These and other advantages and features will become more
apparent from the following description taken in conjunction with
the drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] The subject matter, which is regarded as the invention, is
particularly pointed out and distinctly claimed in the claims at
the conclusion of the specification. The foregoing and other
features, and advantages of the invention are apparent from the
following detailed descriptions taken in conjunction with the
accompanying drawings in which:
[0008] FIG. 1 is a schematic view of an embodiment of a rotary wing
aircraft;
[0009] FIG. 2 is a schematic view of an embodiment of a rotor blade
of a rotary wing aircraft illustrating root, main and tip areas of
the rotor blade; and
[0010] FIG. 3 is a cross-sectional view of an airfoil shape of a
rotor blade compared to a prior art airfoil shape.
[0011] The detailed description explains embodiments of the
invention, together with advantages and features, by way of example
with reference to the drawings.
DETAILED DESCRIPTION
[0012] Shown in FIG. 1 is a rotary wing aircraft 10 having a main
rotor 12 with a plurality of main rotor blades 14. An airframe 16
supports the main rotor 12 and a propulsion system 18 which drives
the main rotor 12. The aircraft 10 may also include a tail rotor 20
having a plurality of tail rotor blades 22.
[0013] FIG. 2 illustrates a general exemplary plan view of a main
rotor blade 14. The main rotor blade 14 includes a root region 24
which extends from a hub 26 over about a first 15% to 20% of a main
rotor radius 28 and includes means by which the rotor blade 14 is
secured to a rotor head 48. A main region 30 extends from the root
region 24 to about 80% to 95% of the radius 28. The blade further
includes a tip section 34 that extends outboard of the main section
from about 80% to 85% of the radius 28 to the blade tip 36.
[0014] The rotor blade 14 has a cross-sectional airfoil shape of
the present invention over at least part of the radius 28 which
alleviates the pitching moment of prior art blades while
maintaining the positive aerodynamic characteristics of the blade
14. This is accomplished by providing a reflex camber over about
the aft 20% chord of the blade 14, "aft" referring to a portion of
the blade 14 closest to a trailing edge 38 of the blade 14. Reflex
camber is imparted on the blade by deflecting the trailing edge 38
upward, in some embodiments by about 6% of chord over the prior art
blade. The addition of the reflex camber to the prior art airfoil
shape allows such an airfoil shape to be utilized over a larger
radius 28 of the blade 14 thereby further maintaining the good lift
and low drag aerodynamic characteristics of the airfoil shape but
with reduced pitching moment.
[0015] Adding reflex camber effectively reduces the net
overpressure on the lower surface near the trailing edge thereby
reducing the exceedingly large negative (nose-down) pitching moment
of the blade 14 that was produced by the original (prior art)
airfoil. Reducing the magnitude of the negative pitching moment is
a desireable effect and enables the use of the airfoil for
rotor-blade applications.
[0016] The airfoil cross-sectional shape is shown in FIG. 3 as
ratios of x and y coordinates to chord length, "C". The values are
then simultaneously scalable to any dimensional chord length.
Because of the difficulty involved on providing an adequate written
description of the particular airfoil section being described, the
coordinates of a particular embodiment of the current invention
reflex airfoil are set forth in Table I below with, as indicated, a
first set of coordinates corresponding to an upper surface of the
airfoil, as referenced by its attitude during normal upright flight
of the helicopter, and a second set of coordinates corresponding to
a lower surface of the airfoil.
TABLE-US-00001 TABLE I X/C Y/C Airfoil Pressure Surface: 1.000000
0.008466 0.995427 0.008495 0.989669 0.008577 0.982521 0.008760
0.973944 0.009084 0.964241 0.009560 0.953842 0.010191 0.943083
0.010961 0.932157 0.011851 0.921155 0.012850 0.910114 0.013947
0.899053 0.015124 0.887980 0.016378 0.876896 0.017700 0.865806
0.019079 0.854709 0.020515 0.843605 0.021997 0.832496 0.023521
0.821380 0.025082 0.810260 0.026673 0.799144 0.028288 0.788035
0.029919 0.776933 0.031560 0.765840 0.033201 0.754752 0.034833
0.743671 0.036445 0.732590 0.038031 0.721512 0.039586 0.710432
0.041099 0.699348 0.042573 0.688261 0.044001 0.677170 0.045382
0.666074 0.046713 0.654975 0.047994 0.643871 0.049224 0.632762
0.050408 0.621651 0.051547 0.610536 0.052641 0.599416 0.053693
0.588295 0.054706 0.577170 0.055679 0.566044 0.056616 0.554916
0.057515 0.543786 0.058378 0.532655 0.059206 0.521522 0.059996
0.510389 0.060750 0.499254 0.061466 0.488117 0.062145 0.476979
0.062786 0.465841 0.063388 0.454702 0.063952 0.443564 0.064477
0.432425 0.064961 0.421287 0.065405 0.410151 0.065808 0.399016
0.066166 0.387884 0.066478 0.376753 0.066743 0.365626 0.066958
0.354501 0.067121 0.343379 0.067232 0.332261 0.067285 0.321146
0.067281 0.310036 0.067213 0.298930 0.067082 0.287829 0.066883
0.276733 0.066616 0.265642 0.066275 0.254557 0.065861 0.243481
0.065366 0.232411 0.064788 0.221351 0.064124 0.210302 0.063371
0.199265 0.062521 0.188242 0.061570 0.177236 0.060512 0.166251
0.059339 0.155291 0.058040 0.144360 0.056604 0.133461 0.055016
0.122600 0.053266 0.111782 0.051336 0.101016 0.049212 0.090317
0.046880 0.079712 0.044325 0.069246 0.041530 0.059004 0.038487
0.049136 0.035199 0.039876 0.031711 0.031535 0.028127 0.024381
0.024597 0.018505 0.021251 0.013799 0.018154 0.010060 0.015297
0.007087 0.012635 0.004724 0.010107 0.002865 0.007662 0.001449
0.005256 0.000468 0.002850 Airfoil Suction Surface: 0.000000
0.000000 0.000262 -0.001992 0.001347 -0.004210 0.003051 -0.006022
0.005216 -0.007567 0.007843 -0.008982 0.011028 -0.010350 0.014933
-0.011716 0.019783 -0.013104 0.025830 -0.014529 0.033254 -0.015977
0.042004 -0.017398 0.051774 -0.018736 0.062178 -0.019950 0.072925
-0.021036 0.083851 -0.022002 0.094871 -0.022869 0.105945 -0.023649
0.117055 -0.024362 0.128187 -0.025022 0.139335 -0.025639 0.150495
-0.026220 0.161666 -0.026771 0.172843 -0.027305 0.184022 -0.027819
0.195203 -0.028316 0.206382 -0.028797 0.217560 -0.029261 0.228734
-0.029705 0.239907 -0.030125 0.251077 -0.030524 0.262243 -0.030898
0.273408 -0.031244 0.284572 -0.031563 0.295735 -0.031851 0.306897
-0.032111 0.318058 -0.032340 0.329219 -0.032540 0.340378 -0.032709
0.351538 -0.032844 0.362698 -0.032951 0.373858 -0.033026 0.385017
-0.033071 0.396176 -0.033086 0.407334 -0.033067 0.418493 -0.033018
0.429652 -0.032938 0.440809 -0.032829 0.451967 -0.032687 0.463125
-0.032515 0.474281 -0.032311 0.485436 -0.032075 0.496592 -0.031805
0.507747 -0.031502 0.518901 -0.031166 0.530056 -0.030797 0.541211
-0.030394 0.552367 -0.029958 0.563524 -0.029493 0.574681 -0.028995
0.585840 -0.028468 0.596999 -0.027918 0.608158 -0.027339 0.619319
-0.026740 0.630479 -0.026119 0.641640 -0.025477 0.652802 -0.024815
0.663965 -0.024137 0.675128 -0.023444 0.686293 -0.022735 0.697458
-0.022005 0.708626 -0.021257 0.719794 -0.020485 0.730964 -0.019686
0.742136 -0.018862 0.753311 -0.018010 0.764486 -0.017130 0.775664
-0.016218 0.786843 -0.015274 0.798023 -0.014298 0.809204 -0.013293
0.820386 -0.012260 0.831565 -0.011203 0.842744 -0.010126 0.853920
-0.009030 0.865090 -0.007915 0.876257 -0.006789 0.887416 -0.005654
0.898563 -0.004516 0.909695 -0.003387 0.920805 -0.002274 0.931874
-0.001189 0.942861 -0.000144 0.953675 0.000838 0.964125 0.001734
0.973869 0.002511 0.982479 0.003138 0.989650 0.003606 0.995422
0.003942 1.000000 0.004188
[0017] To reduce the pitching moment, in some embodiments, the
airfoil shape of Table I is applied at the tip section 34, for
example, at an outer 5% to 7% of the blade span 28. It is to be
appreciated that the airfoil section of Table I may be applied to
other locations along the radius 28, or even the entire radius 28.
Further, while in the embodiments described above the airfoil shape
is applied to a main rotor blade 14 of a rotary wing aircraft 10,
in other embodiments, the airfoil shape described herein may be
utilized in, for example, the tail rotor 20 or in other
applications such as prop-rotor, propeller blades, turbomachine
blades, or the like.
[0018] While the invention has been described in detail in
connection with only a limited number of embodiments, it should be
readily understood that the invention is not limited to such
disclosed embodiments. Rather, the invention can be modified to
incorporate any number of variations, alterations, substitutions or
equivalent arrangements not heretofore described, but which are
commensurate with the scope of the invention. Additionally, while
various embodiments of the invention have been described, it is to
be understood that aspects of the invention may include only some
of the described embodiments. Accordingly, the invention is not to
be seen as limited by the foregoing description, but is only
limited by the scope of the appended claims.
* * * * *