A large scale tunnel spanning wing was built and tested. The model can be operated as either a swept or unswept wing and can be tested in steady state or oscillated sinusoidally in pitch about its quarter chord. Data is taken at mid-span with an internal 6-component balance and is also obtained from miniature pressure transducers distributed near the center span region. A description is given of the system and a brief discussion of some of the steady and unsteady results obtained to date. These are the steady load behavior to Mach numbers of approximately 1.1 and unsteady loads, including drag, at a reduced frequency of approximately 0.1

Carta, F. O.

Jepson, W. D.

Rorke, J. B.

St.hilaire, A. O.

English

NASA Technical Reports Server

if
N79-20059
- .q
29
PROGRESS REPORT: SOME STEADY AND OSC%LLATING AIP_OIL TEST RESULTS,
INCLUDING THE EFFECTS OF SWEEP, FROM THE TUNNEL SPANNING WING
Franklin O. Carta and Arthur O. St. qilaire
United Technologies Research Center
James B. Rorke and W. Donald Jepson
Sikorsky Aircraft Division
United Technologies Corporation
SUMMARY
A large scale tunnel svanning wing has been built and tested. The model
can be operated as either a swept or unswept wing and can be tested in steady
state or oscillated sinusoidally in pitch about its quarter chord. Data is
taken at mid-span with an internal 6-component balance and is also obtained
from miniature pressure transdmcers distributed near the center span region.
This paper presents a description of the system and a brief discussion of
some of the steady and unsteady results obtained to date. These are the
steady load behavior to Math numbers of approximately i.i and unsteady loads,
including drag, at a reduced frequency of approximately 0.i.
INTRODUCTION
It has long been recognized that conventional two-dimensional aerodynamic
testing is not adequate for helicopter rotor blade applications. A typical
section of the blade is simultaneously subjected to wide variations in Math
number, skew angle, and incidence angle, and all of these variations are both
spatial and temporal. It is obvious that a traditional two-dimensional steady
state test program can only satisfy a few of the quasi-steady needs of the
designer, and even a sophisticated two-dimensional unsteady test is severely
limited to incidence angle variations of the typical section. In recognition
of this need for a test facility specifically geared to helicopter appli-
cations, United Technologies Corporation, through the combined efforts of its
Sikorsky Aircraft Division and its Research Center (UTRC), has developed a
minimum wall interference approach to obtailling airfoil aerodynamic data in
the wind tunnel. Data are being obtained for both steady and unsteady pitch-
ing motions, both in oblique and conventional flow, over a wide range of Mach
numbers and at representative full scale Reynolds numbers, using the same
airfoil model and its associated measurement systems. The purpose of this
paper is to present a brief description of the facility and its mode of
operation and to discuss some of the results obtained to date.
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1979011859-437
https://ntrs.nasa.gov/search.jsp?R=19790011888 2020-03-20T18:41:02+00:00Z
I
J
SYMBOLS _i
Values are given in both SI and U.S. Customary Units. The measurements
and calculations were made in U.S. Customary Units.
b semichord, m
CL, CD, CM lift, drag, and pitching moment coefficient
CN, C normal and chord force coefficientc
f frequency, cycles/sec
k reduced frequency
M Mach number
r/R span ratio
RN Reynolds number
V velocity, m/sec
instantaneous incidence angle
aM mean incidence angle
A_ amplitude of motion
A wing sweep angle
rotor advance ratio
rotor azimuth angle
frequency, rad/sec
( )N value taken normal to span
( )® free stream value
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1979011859-438
FACILITY DESCRIPTION
The wide range of parameter variations encountered in rotor craft
is indicated by the plets in figs. i and 2. In fig. i a typical spanwise
variation of lift coefficient as a function of Mach number for several blade
attitudes (including hover) is superimposed on two operating conditions of
concern to the designer, the drag divergence region and the potentially
dangerous condition involving severe stall. (Note that the CL peak rear the
blade tip in hover is caused by a trailing vortex encounter). Similarly,"
typical contours of constant skew an_e (solid lines) and constant Mach number
(dashed lines) are superimposed cn a rotor disk in fig. 2. It is seen that
the high load conditions are generally encountered at moderate Mach numbers on
the retreating side of the disk, but thgt a need exists for high Mach number
data as well. Furthermore, the effoctive blade sweep angle is zero only at
azimuth angles of _= 90° and 270 ° a1:d is 15° or greater over 60 percent
of the rotor disk (as shown by the shaded region in fig. 2).
The first step in responding to these clearly defined needs was to build
a versatile steady-state facility. The resulting Tunnel Spanning Wing system
(TSW) was constructed by Sikorsky Aircraft in 1971. iu consists of _ basic
tunnel spanning rectangular steel spar approximately 2.44 m (g ft) in length,
plus additional end pieces to complete the spar for installation at several
available sweep angles or for installation in different wind tunnels with a
test section that has one dimension at leas_ 1.83 m (b ft_. Interchangeable
airfoil-shaped shells with 40.64 cm (16 in) chord are mounted in sections
onto the spar to provide the test configuration. A sc_,ematic view of this
system is shown in fig. 3. It is seen that the model shell consists of two
sets of pieces: the upper portions which surround the spar fore and aft, and
the lower cover plates which complete the airfoil profile. Also shown in
fig. 3 is the center span metric section, 20.32 cm (8 in) in width, which
mounts to a pair of completely enclosed strain gage balances. This allows the
airloads to be measured far from the tunnel side walls and ceiling. With this
system it is possible to test several airfoil profiles in a single installa-
tion by replacing one set of model shells with another. The nominal chord
length of 40.64 cm (16 in) allows data to be obtained at representative full
scale Reynolds numbers and at a favorable tunnel height to chord ratio of 5.25
or greater for the tests conducted to date.
In addition to the internal balance in the metric section, two surface
pressure tap systems are incorporated, arrayed principally in a chordwise
dicection. One system was originally installed along the centerline of the
metric sectlon to measure only steady state pressures for comparison wi_h
balance results. A second system was later installed in the model shell
immediately adjacent to the metric section to measure both steady and
oscillatory loads. As part of this system, some additional spanwise taps were
included to measure spanwise loading and are so arrayed to relate the effects
447
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............ -.................. 1979011859-439
of chordwise and streamwise surface pressure distributions when testin_ in
oblique flow. Thus, the TSW incorporates dual airloads measurement systems
which are redundant over a wide range of the test envelope and, together, are
used as a data self-checking system. Hot film skin friction _ages have also
been included to assess the surface flow zonditions in steady and oseilla_ory
situations. A description of models and results of steady tests are presented
in reference I; the unsteady tests, not yet published, were performed by F. O,
Carta at United Technologies Research 6enter (UTRC) under NASA Contract
NAS 1-14873.
The TSW has been tested by the Navy and Sikorsky Aircraft in the NSRDC
2.13 m x 3.05 m (7ft x 10ft), 12 percent permeable transonic wind tunnel at
zero sweep angle up to a maximum Mach number of M _ i.I (ref. l). The 3.06 m
(10ft) span model installation is pictured in fig. 4. In this figure are
shown a pair of part span supports from the tunnel floor to the model pressure
surface. Tests with and without the struts installed showed conclusively that
these supports had a n,gligible effect on the measured data. (The round
object in the lower riKht corner of the tunnel was a part of the NSRDC sting
mount syste located severa], meters behind the TSW, and had no effect on the
results). The M_ch number/incidence angle and Mach number/Reynolds number
envelope_ achieved in these tests are shown in fig. 5. For comparison, a
t:pical rotor envelope for a dive/pull-up maneuver is compared with the NSRDC
M/u test envelope in the left panel, and the test range (dashed region) is
compared with typical M/RN variations for three helicopters in the right
panel. It is believed that the test envelope represents some of the widest
combinations of M and a achieved to date in a wind tunnel with a model having
a full scale chord, and at conditions well above critical Reynolds number
range.
A cooperative effort by _JTRC and Sikorsky Aircraft has yielded the
oscillatory system shown schematically in its unswept position in fig. 6.
This system consists of a Hrive motor and transmission mounted beneath the
UTRC 2.44 m (8ft) octagonal wind tunnel, which actuates a pair of push rods
and oscillatory cranks to provide a sin_soidal motion of the model about its
quarter chord (ref. 2). A swept, oscillating installation in the UTRC tunnel
at A = 30° is pictured in fig. 7, looking upstream. In this test, one-
third span supports were employed from both ceiling and floor uf the tunnel to
pivots located in the model spar. These were installed to overcome oscil-
latory bending deflections resultinz from a small chordwise noncoincidence of
moael c.g. and pivot ax_s. (Note that the model was originally designed for
steady testing only.) Steady-state oil flow studies have demonstrated that
only a small portion of the airfoil surface area aft of the quarter chord was
affected in the immediate vicinity of the supports, and that the effect on the
center span flow was negligible. Oscillatory tests at ^ = 300 have been
conducted at freestream Math numbers as high as M® = 0.58, and at reduced
frequencies up to kN = b_/VcosA = 0.I over a wide range of incidence angles
and at amplitudes of ± 8 and ± I0 de£.
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1979011859-440
Testing was also conducted at A = 45° in steady state by Sikorsky
Aircraft in the UTRC tunnel. This installation is pictured in fig. 8 as
viewed from above. In this sequence of steady-state tests the model was i
.. subjected to a free stream Mach number as high as M, = 0.83. or a normal to
span Mach number of MN - M.cosA = 0.59.
TEST RESULTS
A typical set of steady-state test resul_s for the Sikorsky Aircraft
SC-I095 airfoil are found in figs. 9, i0, and Ii, obtained in the NSRDC
tunnel (ref. I). These figures contain the lift coefficient versus incidence
angle, and the drag and pitching moment coefficients plotted as a function of
lift coefficient for a wide range of Mach numb( q from M = 0.3 to 1.075. The
solid curves were obtained from balance measurements and the dashed curves
are from integrated surface pressure and wake rake measurements. The flagged
symbols represent repeat points. The internal balance measurement system was
initially included for this type of testing to obtain drag data at high Mach
numbers where compressibility effects preclude use of a wake rake. The
several curves in each figure are plotted relative to a staggered set of
origins, indicated by tic marks and zeroes along the left ordinate. Z_e
scale for the M = 0.3 curve is shown on the right ordinate. These figures
show that all data exhibited excellent repeatability and data from both
systems generally substantiated one another concerning the trends of he
force coefficients with incidence angle and Mach number. The lift data from
these separate measurement systems were in close agreement. The drag data
from both systems were very similar, but at 0.9 Math number the wake rake
values were lower due to turbulent flow and/or compressibility effects. The
pitching moment data from both systems were generally in agreement; however,
at some conditions the balance coefficient data were more positive by + 0.015.
After correcting for :he differences in lift curve slope that occur between a
ventilated and a solid wall wind tunnel, these data are in close agreement
with the data obtained in UTRC 8 foot solid wall wind tunnel.
A few selected results from the unsteady UTRC tests are shown
in figs. 12, 13, and 14 at MN = 0.3 and 0.4 at A = 0o and 30°. Recall-
ing figs. I and 2, it can be seen that these conditions are of primary
importance to the rotor designer because they are representative of the
flow experienced on the retreating side of the rotor disk where unsteady
stall conditions are encountered. In figs. 12 and 13 the results were
obtained from integration of the unsteady pressure distributions at each
instant of time and in fig. 14 the loads were calculated from the
unsteady output of the internal balances.
449
1979011859-441
The effect of varying frequency at constant incidence is shown in ._i
fig. 12 for the unswept SC-I095 aicfoil. Here a mean incidence angle of
aM = 12° al,d an amplitude of ± 8° preduces periodic penetration of the
stall regime, and as expected (refs. 3, 4, and 5), an increase in frequency
causes a progressively larger overshoot of both normal force and moment in
the dynamic stall regime . In fig. 13 sweep and oscillatory effects are
combined. The solid lines represent the unswept configuration and the dashed
lines the swept, both for a mean incidence angle of aM = 15° and an
amplitude of ± 8° (measured in the plane normal to the wing leading edge).
Although a single sample such as this is insufficient to establish a trend,
it is seen that the effect of sweep is to reduce the magnitude of the
unsteady excursion through the dynamic stall regime (i.e., sweep appears to
"soften" the impact of dy._e,nic stall). It is also interesting to note that
for unsteady conditions there is no___tta further extension of the maximum lift
coefficient in oblique flow as is experienced in steady oblique flow.
Finally, fi B. 14 shows unsteady lift a,_d drag coefficient varia-
tions for the unswept wing at two mean incidence angles as obtained from the
balance system. The significant _bservation to be made here is that, for the
moderate Mach numbers tes£ed, the unsteady drag is generally greater for
increasin_ incidence angle and follows a clockwise loop in the high incidence
regime, similar to that of the unsteady normal force. This behavior is
easily explained by noting that the drag is made up of a vector sum of the
normal and chord forces, expressed in component form as
CD = CN sins - C cosac
For incidence angles in the range 15° to 20° _he cosine is approximately
three times the sine, but the normal force is at least I0 to 20 times the
chord force. Ccnseqvently, the CN sina term dominates the equation in the
dynamic stall regime and the unsteady Crag behaves as shown.
CONCLUDING REMARKS
The TSW concept has been used to demonstrate that a consistent airfoil
data bank can be obtained at representative full-scale Reynelds numbers using
the same model system for a wide variety of te_t conditions approaching those
that are encountered by rotary wing aircraft. This model system provides a
means to minimize wall interference effects and to eliminate the complicating
factors introduced by using different airfoil models and load 1_asuring
system8 required to cover this extensive aerodynamic environmeut, both steady
and unsteady. This paper illustrates the versatility of this system with the
few o×amples presented herein.
i
450
19790-1185924ff-
REFERENCES
i. Jepson, W. D: Two Dimensional Test of Four Airfoil Co1,_igurations
with an Aspect Ratio of 7.5 and a 16 Inch Chord up to a Mach Number
of i.I. Sikorsky Enmineering Report SER-50977, April 5, 1977,
performed under Contract No. N60921-73-C-0057.
2. Carta, _.O. and A. O. St. Hilaire: An Experimental Study of Sweep
Effects on the Unsteady Aerodynamics of a Pitching Airfoil.
United Technologies Research Center Report R76-411931, March 1976.
3. Carta, F. O. and C. F. Niebanck: Prediction of Rotor Instability
at High Forward Speeds. Vol. III. Stall Flutter. USAAVLABS
Technical Report 68-18C, February 1969.
4. Liiva, J., F. J. Davenport, L. Gray, and I. C. Walton: Two-Dimension=i
Tests of Airfoils Oscillatin_ Near Stall. US_%VLABS Technical Report
68-13, 1968.
5. Cart, L. W.. K. W. McAlister, and W. J. McCroskey: %nalysis of
the Development of Dynamic Stall Based on Oscil]at_n_ Airfoil
Experiments. NASA TN D-8382, January 1977.
45!
L
....'....."......... ".................... _" -_- 1979011859-443
!
"I
1.2 _.
__/_i_ = 240 °
._ _I_,. - ".:_"_i_i_:_i_i_i__evere stall
1.0 ._ I:_ur_l_i!i!i_ili::_ and drag
• '----"_'-_!i_ div e rge nc•
Lift coefficient, 0.8- '_>_3i_:l _"_iiii!iiiii_".:_ifi!!i!!:_!;_iiiii}!_iiiii!ii!i!iiii_.. :::::::::::: ::::::::::::::::::::::::::::::::::::::::::
- _ Hover _::::ii_
CL _ii:,_ii0.6- *- • ":_i_i::i::ii_i_::!i!i!:Z::i':_::i::!i::.!:.i::i::_i_i::
. _" _i;'_{ili_i',',i_i:,',}
me_ R_Root r / R :)i:i!i;_!:_i:ii;!:!ii!!!i::::!:!'!:i:i:i:"Advancing¢e,_0.2I-I _, =_oo
_.oo, __,.
0 0.2 0.4 0.._ 0.8 1.0
Mach number
Figure i.- Operating envelope.
-_ Approech velocity
lSO_
--_ Skew angles ;_6!i_::'" o
/..;__ilili.3°...._::_:_:_:_:)._:.:-_:_:_:_:_:i_iii!iii_iiii
i!'i}iiq....iii
--__
a_5o" _i::i.3.o°_
0 o Azimuth position
Figure 2.- Contour,_ of skew angle and M_ch number.
4
452
1979011859-444
Upper metric section --_ Static "_
L1%%Task balances_ -_'_. _._'_ l _ J
' • section_\_'_." ._/
Pitch_/'\_ Tunne, a w
arm _.
Figure 3.- Tunnel-spanning wing assembly.
Figure 4.- NSRDC installation. A = 0°.
453
1979011859-44
": " 4
20 '_ _ /- SlK/NS_IDC test
15 _-_envelope
10 Rotor envelope> _
r_ in dive/l)u-li-_r) _ -_ 2 CH-53E_ ..._ "'" I
, ,-m.n.uve,':!
I I I I I I
-5(_ 0.2 0.4 0.6 0.8 1.0 1.2 0 0.2 0.4 0.6 0.8 1.0 1.2
Mach number
Figure 5.- Comparison of rotor and NSRDC test envelope. Steady s_ate.
ow Blade shaft
and support
Push rod
.tf"
Orive shaft j/ Eccentric cam
]Ljt f"
Figure 6.- UTRC main wind-tunnel es_:i]]atory mode] system.
454
1979011859-446
\Figure 7.- UTRC instal]ation. A = 30° .
455
1979011859-447
Balance --_- - Integrated pressure _0-----
IM Flagged symbols indicatezlaO75
I _ repeat points
0 - _ _o
c;-_M'O'98r_.
-j
Incidence angle, 0 - _- 18.0
a (deg)
0 16.0
12.0
0 8.0
J 4.0
L× "o z_" o¢ J , i _ l A -4.0
-0.2 0 0.2 0.4 0.6 0.8 1.0 1.2
Lift coefficient, C L
Figure 9.- Steady-state lift coefficient plotted against incidence angle.
Balance --A- - Integrated pressure ----0--
I Flagged symbols indicate
0.1 -
_. ___ repeat points
M=1.075
0(M=1.075)- b-._. _._&..b- ..,&'_M=0.98
E tDr,_q coefficient. C D _ -0,9 - 0,3
0 _ - 02
M=O.6 a_ \,.
0 o------<_ --------o-- -z__- _,_m,,, 0.1
M=0.3._,_._
o _ _-------_:__ o
I
/
l [ ; , _ l l l -G.1
-0,2 0 0.2 0.4 0.6 0.8 1.0 1.2
Lift coefficient, C L
Figure I0.- Steady-state drag coefficient plotted against lift
couf ficient.
456 I
1979011859-448
I'r- ,"r
8
Balance - -A - -- Integrated pressure _0_-
.... Flagged symbols indicate
0 - _ _=1.075 repeat points
o- o_
Pitchk,_l moment M-0.9
coefficient, CM 0 l_ _____M-OJ
0 0
_ .--o.3.o._I 1 i i l i i
-0.2 0 0.2 0.4 0.6 0.8 1.0 1.2
Lift coefficient, CL
Figure ii.- Steady-state pitching-moment coefficient plotted against llft
coefficient.
2.0
f J
1.8 _. , ", j_ 4 CPs O.1
1.2 ". _" 0 .<_i>'_ "...,
0.8 _- -0.1 _"
i '0.4 -0.2 i _/ ,,'/"10 CPS0 0 _ ' ' ' ' -0.3 _ , , '_;_../, (.4 8 2412 16 20 0 4 8 12 16 20 24
Angleof attack, o_,deg
Figure 12.- Effect of frequency on oscillatory normal-force and moment
coefficients from integrated pressures. SC-I095 airfoil; M = 0.4;
Aa = -+8°; c_M = 12°
,i
457
t 1979011859-449
A am=15 oMN=0-3 _ 0 o
kN=O.095 --- 30 ° _=+8 °
1 "
" 1.6 _" ,,_r),% _/_/ 3 0.1[ £_Ste&dy
_ -0.1_" Steady-_'_
0.8_ __/_ -0.211 Unsteady j: P
_ 0":_ d:,_Un_t:e? ?dy 2! _0.3__, i , i , i , i _j'_" I) 8 16 24
Incidence angle, _N,d_ q __ Incidence angle, CCN,deg
Figure 13.- Effect of sweep on osc:llatory lift and moment coefficients
from integrated pressures. SC-I095 airfoil.
2.0- '_
Lift 1.6 -
coefficient, 1.2
CL 0.8- ,t":_ Steady
0.4 -/,- _- stateOr I I I 1 i
1.0 (x M =15 ° !
0.8
Drag 90,!///_coefficient, 0.6
C D 0.4 Steady s,ate-_ _0.2 - _,-. .
0 _,.-.._,J:;-_"l 1 .
0 "-'-4 8 12 16 20 24
Incidence angle, (x (deg)
Figure 14.- Unsteady lift and drag coefficients from balance. SC-]095
airfoil; M = 0.3; A = 0°; A_ = +-8°; f = % cps; k = 0.096.
458
1979011859-450


Some steady and oscillating airfoil test results, including the effects of sweep, from the tunnel spanning wing

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