Wind-Tunnel Test of an Articulated Helicopter Rotor Model - NASA

93 Pages·2005·4.35 MB·English

https://ntrs.nasa.gov/search.jsp?R=19810005407 2019-03-28T11:44:55+00:00Z
NASA AVRADCOM
Technical Mcmonndurn 00110 Technical Rcport 79-49
Wind-Tunnel Test of an Articulated
Helicopter Rotor Model
With Several Tip Shapes
john U. Rerry and Raymond E. hineck
A wind-tunnel investigation has been conducted using an articulated heli-
copter model rotor systen to determine the effects of several blade tip designs
on both the perforaance m a acoustic ~h~ta~ttristoif ~tsh e totor. Cnly the
porfoasurco characteetsfics are presented in this paper. Six tip shapa rere
testtdr a squar* (baseline) tip, an ogee tip, a eubwing tip, a swep2 tip, a
winghet tip, and a short ogee tip. Teats were conducted at two rotor rotci-
tional apccds and five advance ratios including hover.
in hover at the lower rotational speed and a given lift cooffic5ent, the
swept, oqee, and short oget tips had about the saw torque cocff icitnt, and 'chc
subwing a d wingbet tips had a larger torque coefficient than the baseline
square tip blades. In simulated forward flight, tha ogee tip had valuea of
torque coefficient very close to or leas than the baseline square tip blades.
In simulated foruard flighk at 1200 rpm, the ogee tipa siumed a ckcraasc? in
torque coefficient relative to the baseline blades as drag a x f fi ciant
decreased. The swpt tip blades required less torque coefficient at lower rota-
tional sped8 than the baseline square tip blades and roughly equivalent totque
coefficient at higher rotational speeds in forward flight. The short ogee tip
was tested in limited forward flioht mnditions and it required higher torque
cxxfficient at higher lift coefficient than the baseline squace tlp blade.
Excessive noise has limited the acceptability of the helicopter near popu-
lated areas to such an extent that noise reduction has became an imlportant f%c-
tor in helicopter design. The effect on rotor performance of the changes in the
design to reauce noise must be evaluated. #my past tests of these designs Rave
include2 thorough investigations of either the prformance or acoustic char&.,--
teristics but not both. It is difficult to determine the eIfect of design
chsnoes on the robr acoustics and prformance because the past test.^ ze been
performed on different rotor systems with different scale factors. !iL,~?al
rotor blade tip shape designs have demonstreted improved acoustic character-
istics. Tc; evaiuate the effect cf tip shape on the perform&nce csaracteristics
of a rotor, a wind-tunnel test has been conducted using a single articulated
model rctor system with five interchangeable sets of blade tips. Subsequent
to the first test, a sixth tip shape was evaluated for performance in c limited
E-angeo f flight conditions. The first five tip shapes were evaluated fc;c acous-
tic performance. The acoustic results from the tests may be found in refer-
ence 1: the performance results are presented herein.
The six removable tips tested were a square tip (baseline), an ogee tip,
e swept tip, a subwing tip, a vinglet tip, and a short ogee kip. Each set of
tips was tested in hover at a rotational speed of 1200 r2m. Because tP,e radii
for all tip shapes were not identical, this speed corresponded to tip speeds
ranging from 19d Wsec (649 ft/sec) to 207 Wsec (680 ft/sec) depending on the
radius. The square tip, the w e e tip, ar,d the swept. tip were subsequently
tested in hover at 1340 rw and in simulated forward flight at 1200 L ~ I Ia nd
1340 rpm and four advance ratios of 0.2, 0.3, 0.35, and 0.40. The short ogee
tip wns tested in simulated forward ftigh: at 1200 rpm at an advance\ratio
of 0.3.
Th& units deed foe the pkysical q~antitiesd efined in this paper nre
given in the Intsrnational System of Units (811 and parenthetically in the
ly.."r.C ust:&isr,y Ysits. Meesuremeats and calculations were made in the U.S.
Custanary Unite. Conversion factors relating the two system are presented in
reference ?. The retot prfotxtnce data have been reselwd in the atabili ty
axis system with the BDtment reference cente: located at the center of khs
rotor hub.
Q
conventional rotor torque coefficient,
~>TIR~R%
conv~ntionalr otor thrust coefficient
rotor blade irfcal chord, m (ft)
rotor biade reference chordp 0. V3 a
torque-eighted eEfective blade chord, m (fc)
I
ft4
---"-0- -
thrust-veightctd effective blade chord, P a Ifk?
w3
rotor drag, N (Ibf)
rotor If ft, el (1Bf)
rotor rotational sped, sp!
rotor torque, WE iibf-ft)
rotor radius, E i'ft:
rotor thrust, W (lbb)
~otort ip npeed, RR, w'wc f ft/wc]
free-strcm velocity, ~ ' 9 4(f~t/s ec)
rotor snafc: a ~ ~olf eat tack, deg
-v, -
rotor advance ratio,
ree-strem density, kg/m3 (slugs/£ t3)
f
be
rectangular rotor :;clidltyr .-
n n
-b c ~
torque-uai j'?t;ed sol j c,.ry,
x R
rotor rotational speed, rad/sec
,
rotor blade natural frequency rad/eec
11008, Am APPARATUS
The general rotor -1 system (CRWS) was uaed in the trngley V/STOL Tunnel
,
for this investig? ticm. A sketch of the rodel appears in figure 1 vRich a180
present* the axis system. Details of the rod.1 may be found in references 3
and 4. The rotor hub ausd in the inwstigatiocr ms fully artisulated. The
flapping an8 1-iq hinges were coincident at 7.6 cn (3.0 in.) f r a the center
of rotatla. The pitch-flap coupling angle uas set at -2O. (Negative pitch-
ti* cccrgliry snglss mrrespasrd to a decrease in blade pit& uiUl an inctease
in blade fl.pping.1 Position wtentiomttrs were munted on the flapping and
lagging hingco of thc refcrencc blades to provide a readout of the flap~inga nd
lagq5ng aotims. T k ratot control system was instrumented to provide a readout
of the rotor control positions. The pitch link of the reference blade had a
athein gage to aeasure control loads.
The rotor was driven by tw 67-kW (90-hp) electric motors driving a cow
aon transmission. The rotor, transmission, and motor ass@ab;y was nounted on
a six-coqmneni: strain-gage balance. The balance was supported by a gimbal
system of springs and dampers to insure that the rotor would !?e free from ground
resonance.
The rotor blades used in the investigation had an NACA 0012 airfoil sec-
:ion. Each blade had provision for a removable tip. Six tip shapes were
tested in this investigation: a square tip (baseline) , an ogee tip, a swept
tip, a subwing tip, a winglet tip, and a short ogee tip. (See fig. 2 and
table X. ) The square tip, the swept tip, and the winglet tip were all the same
lenqeh so that the radii of these blade sets were equal. The wee tip was
longer than the baseline square tip so that its blade radius was larger b u ~th e
blad? erea was the same as that of the square tip. The subwing tip ccnsisted
of n basic square tip with a subwing added so that both the radius and area were
larger for the subwing tip than for the square tip. The short ogee tip had the
same pl?nform as the wee tip except that the radius was reduced by beginning
the q c e curvat)- re a t the point oL tip attachment and shortening the extreme
tir
slight!.y which resulted in a radius only 1.12 cm (0.44 in.) longer t5an the
baseiins sqo9re tip. Details of the rotor blades are given in table I.
The !?lades and tips were designeci LJ keep the change in mass between the
differerit types cf tips small. The change is reflected in the blade flapping
mettias sh~wni n t3ble I. The natural frequencies of the rotor blade with
each tip w?re det.ermined by the method described in reference 5 and the fre-
quencies for all six tips were very similar. (See table 11.) The effect of
blade dynamics on the performance of the rotor should be the same for each tip.
The short ogee tip shape was tested subseqi~entt o the main test. The
rotor system is identical to that used in the main test, but the helicopter
model bcdy murilted on the GlWS during testing of the sixth tip shape was that
of the YAH-64 attack helicopter .e '' he1 i.-opter model used during the main
test was a research helicopter ba2y designed to be configuration-indwndant
and easily modeled analytically.
Photograph8 of the mdels in the turgley V/mL Tunnel are sham in fig-
ure 3. The model vas mounted on a ,ecand strain-gag& balance attached to a
special model sting. 'Ihis sting -1 support system perritu the angre of
attack and angle of sideslip to be varied war wide ranges while the mcdel is
mintained at a fixed position in the tunnel.
This investigation was mducted in the hnyley 'V/STOL Tunnel, which is
a closed return, atrospher ic tunnel. The tunnel my be con£i gured with s
closed test section or an open test sectian with the walls and ceiling r.e moved.
The closed test section measures 4.42 m (14.50 it) by 6.63 m (21.75 ft) Test-
ing a t forward speed was conducted in the closed test section with the rotor hub
2.9 m (9.5 ft) above the tunnel f l c ~ ~ r!.io ver testing was conducted in tne open
test section at several heights above the tunnel floor: 4.5 m (14.7 ft), 2.9 m
,
(9.5 f tl and 1.6 m (5.4 ft) Five sets of rotor blade tips were tested in
hover, and after analysis of the hover results, it was decided that only three
- -
of those tips t.he square tip, the swept tip, and the ogee tip would be
tested at simulated forward speeds, The short g e e ti^ "as subsequently tested
in hover and a t one forward flight conditior..
The three tips initially tested a t si~nulatedf ofwarr! flight corldi2ions wc-.re
tested at rotational speeds of 1200 rpn and 1340 rpm and a t zdvance ~ a t i o sof
0.0 (hover), 0.20, 0.30, 0.35, and 0.40. These values vere chosen to cover Lhe
practical range used in helicopters. A t each combination of r.~tacir,nal speed
and advance ratio, tests were ,nade a t six nominal shaft angles: 0°, -3O, -6O,
-go1 -12O, and -15O. A t each shaft angle the rotor collective pitch was varied
in lo ar 2O increments until one or more of the following lim its was reached:
maximum temperatures in the electric drive ~ t o r i i ,a vailable collective and
cyclic pitch, lead-lag motion of the blades, and/or stress limits on the blades,
pitch liqks, and strain-gage balances. Combinations of rotor rotational speed,
advance ratio, and shaft angle are shown in table 1x1. A t each collective pitch
setting, the first harmonic longitudinal and lateral flapping with respect to
the shaft were reduced to zero by varying the cyclic pitch settings.
The model hub configuration used is representative of a true scale hub.
For comparison with the baseline blade shape, data presented have not been
adjusted for force and moment contributions due to aerodynamic forces acting on
the model rotor hub. The data for simulated forward flight have been corrected
for wall effects by using the methods described in reftorenca 6. The &--rectad
tunnel free-stream dynamic pressure and flow direction are uY?d to compute
thm
simulated flight speed and rotor shaft angle of attack.
The rotor balance forces and mments, rotor control positions, and tunnel
test conditions were recorded on the tunnel digital data acquisition system.
The instrumented rotor blade flapping angle, pitch link loadi and azimuth were
recorded on an analog tape.
PRESENTATION OF RESULTS
The results of this wind-tunnel investigation have been presented in
dimensional and nondimensional f ~ r min figures 4 to 14. The data are resolved
in the stability axis system wiL9 the moment reference center a t the center of
the rotor hub. Complete results are presented as follows:
-
1
-
Figure for data at p of
as, NI
Result rm
deq
0.0 0.20 0.30 0.35:(.0-
Effect of solidity in . . . . . . .
nondimensionalizing 1200 4Ia) 4 (bi
--- -
--
Rotor performance canparison
(thrust as a function of torque) :
Square, ogee, swept, sub-ding,
ana w.ing.le.t .tip.s. on. r.es.ea.rch. . .
body 0 1200 5(a)
1
Square, wee,, md .ew.e.pt .ti.ps. o.n . . I I
rerear,ch body -3 i200 6(.) 6(b)
-6 l20J 6(C) 6(d) 6(e) 6(f)
-9 1200 6(g) 6(h) 6(j)
-12 1200 6(j) 6(k)
-15 1200 6 (1)
0 1340 5(b)
-9 1340 6
-12 1340 6(h) O(O)
Square and short o.ge.e .ti.ps.a . . . .
on YAH-64 body 0 1200 7(a)
-6 1200 7 (b) - .
Rotor performance
(aerodynamic. c.o.ef.fic.ie.n.ts).: . . . . . .
~qu--et ip 1200 8(a) 8(b) 8(c) #(dl 8(e)
1340 8(f) 8(g) 8(h) B(i1
. . . . . . . . . . . . . . .
Ogers tip 1200 9(a) 9(b) 9{c) 9(d) 9(e)
. . . . . . . . . . . . . . 1340 9(f) 9(9) 9(h) 9(i)
Swept tip 1200 10(a) 10(b) 10(c) 10(d) 10Ce)
1340 10(f) 10(g) 10(h) 10(i)
. . . . . . . . . . . . .
t- Subwing tip
. . . . . . . . . . . . .
winqlct tip
. . . . . . . . . . . .
Short wee tip 13(b) -
I 1
Corparison of tip sham pcrfor.ra.nc.e . . . I1
(aerodynamic mfficienta) 1200 I(.) 14(a) ld(Al l(!ci I4(d)
I
13 40 i 14 ) ! 14(f) 14(g)
*he data in figures 7 and 13 compare the perforlance of the short oget tip
with the baseline square tip when both are tested with the YW-64 helicopter model
shell muneed. A amparison of performance of the baseline squacc! tip blades at
the tuo short ogcc hlade test conditions with both the research and the YU-64
rode1 shells rounttd produced no significant deviation in the results.
DISCUSSION OF RgcULTS
Effect of Nondimensionalizing
Rotor force ar?d moment data are traditionally presented in coefficient
form n~ndimensianr~lizebdy the test medium density, rotor tip speed squared,
and either the rotor blade area or disk area. The different planforms tested
in the investigation complicate the choice of the proper values of radius and
area. A yoper nondincnsionalization will rzsult in a rorpazison of values
correspondino to forces and moments resulting frcm tip shape changes only.
Since changes in tip planform el= affect the rotor swept disk -rea, the tip
speed, an3 the rotor solidity, the effects of these quantities on the nondimen-
sionalization are examined.
The blade element analysis method of reference 7 represents the forces and
momentr, in the rotor system as the sums of incremental forces and rwrents acting
at different blade radial stations. These incremental forces and moments ace
assumed to be only a function of their radial location. Reference 7 gives a
technique for deterniniw weighted solidities for blades of arbitrary planform
based on blade element theory.
The basis Lor the weighted solidities found in reference 7 is an equivalent
effective chord for an imaginary rectangular blade of the same radial lengt;~a s
the nonrectanqular blade. Forces at a given radial station are proportional to
the quare of the radial 6istance from the center of rotation to that statim,
a d ;roaents generated at a given radial station about the center of rotation
are proportional to the ~364o. f the radial distance from the wnter of rotation
to that station. Effective chords for use in thrust (force) and torque (mmnt)
calculations are given by
Weighted solidities, based on the effective chords com~uted,a re given by
For comparison with this method of nondimensionalization a rotor solidity
for each tip was coarputed assuming a rectangular planform to the rull blade
radius, as given by
Values for these solidities for all tip shapes studied are given in table IV.
Comparison of the results of these methods of nondimertsionalizing for hover and
a forward flight condition can be found in f igute 4. The same data presented
.
ibi a dimensional form can be found in figures 5 (a) and ' (c)
It is apparent thzt one's perception of t..a effect of tip snape on rotor
performance is influenced by thz normalizing procedure chosen to aralyze the
data. Nondimensionalization by the weighted-solid ity technique, which, in
e'fect, distributes planform variations uniformly along the blade span, obscures
t9c ?£hence oL shape changes occurring in a highly bccalized area of dis-
in-ztly three-dimensional fl w e ffects because .he normalizing technique is
basi;d on the two-dimensional blade :lement theory. On the other hand, non-
dimensionalization of the data by a rectangular-solidity method, where the
basic blade chord is retained out to the full tip radius, yields a closer
approximation to the actual physical processes involved, since the planform
variations are confi,led to a c~mp,~ativcly small portion of the blade. Although
some variatior due to uneqcal blade planforn areas is present when the
rectangular-solidity technique is used, it is believed that this procedure is
more effective in representing tip shape change effects than the weighted-
solidity method. Accordingly, data presented in this report, unless otherwim,
specified, have been normalized by the rectangular-solidity technique.
Effect of Tip Shape in Hover
The nondimensional rotor hover performance at 1200 rpm is presented in
part (a) of figures 4, 5, 7 to 10, and 13 and figures 11 and 12. When the
results at practical nondimensional lift coefficients (Cua > 0.04) are com-
pared, both the ogee and the swegt tip blades require the same torque coez-
Elcient &t a given lift coefficient as the beseline square tip blades. The
subwing and the winglet tip blades required a larger torque coefficient at
a given lift coefficient than the baseline square tip blades. Because the
suhwing and the winglet tips experienced significantly higher torque coef-
ficients than the other tips during the 7200 rprn hover testing, they were
,
elipinated from further testing. The short wee tip showb less torque cw:
ficient required below a moderate lift coefficient (CL/o < 0.05) than t.. .-
line square tip blades. -
The hover performance at ? 340 rpm is presented in figures 8 ( f), 9 r f j , ar.d
lo(£) for the square, wee, and swept tip blades. &3 was wced at 1200 rpm.
tie ogee Lip requiros the torque coefficient at a given lift coefficistn'
as the squa.re tip blades. The w p t tip requires an increass in torque coef-
ficient for practical lift coefficients *en canpared with the? baseline square
tip blades.
Effect of 'Tip Shape at Simulated Forward Speed
The ntor per toraance data fr on the wind-tunnel investigation are preeented
in figure 8 for the baseline square tip blades, in figure 9 for the ope tip
blades, and in figure 10 for !-he swept tip blades. The rotor lift coefficient,
drag coefficient, torque coefficient, and control-axis angle of attack are pre-
sented for each combination of advance ratio and rotor rotational sp;, Fiji
hover comparison the baseline squape tip data are alao presented in figures 9(a),
9(f), IOia), and 10(f).
4 direct conparison of the results for tiie three tip shapes is complicated
by the small differences in rotor shaft angle of attack. Therefore, the results
for the three sets cf tip shapes have been interplated to several values of
rotor lift m f fi cient. These results are presented in figures ?4(a) to 14 (d)
for 1200 r p a nd in figures 14 le) to 14 (g) tor 1340 rpm.
Data at 1200 rpm indicate that the ogee tip blades required a torque coef-
ficient less than or very cl~:?t o that required by the baselirre square ,;o
blades except in the following two conditions: (1 1 low adv nce ratio (y = 0.20)
and low lift coefficient (Cv'a = 0.84) above a drag coefficient of -0.003 and
(2) advance ratio of 0.38 and li2 coefficients of 0.06 and higher at drag coef-
ficients above -0.001. When compared to the baseline square tip bladesn the
ogec tip blades demonstrated a significant reduction in torque required at an
advance ratio 0.35 with lift coeeficipnts of 0.06 and 0.07 below a d~agc oef-
ficient of -7.006. The data presented indicate a general performmce improve-
ment (reduction in torque required) for the ogee tip blades as the drag coef-
f icient decreases when cocapared with the data for .he baseline square tip
blades. At 1340 rpm, the data presented indicate that relative to the bseline
square ti blades, there is a smaller tcrque coeificient requited for the ogee
)
tip blades at the higher lift coefficient at all but the lowest advance ratio
where the torque requirements are very close.
Data at 1200 rpar for the swept tip blades indicate a torque coefficient
requirement of lass than that for the baseline square tip blades for all cases
except those where torque coeff icients were very close, as follows: (1) at an
ad-lance ratio of 0.20 and low lift coefficient (CL/a = 0.04) below a drag coef-
ficient of -0.083; (2) at th- advance ratio of 0.35 and (a) li,ft coefficient
of 0.04 and drag coefficien less than -0.002, (b) lift coefficient of 0.06
and drag coefficient less th.. -0.006, (c) lift coefficient of 0.07 andl drag
coefficient less than -0.007; and (3) at an advance ratio of 0.40 and lift