NASA Technical Memorandum 102271

Analysis of Small-Scale Rotor

Hover Performance Data

Cahit Kitaplioglu

March 1990

(NA_A-TM-10227i) ANALYSIS OF _NALL-SCALE

R_TQR HOVPR PERFORMANCE DAT_ (NASA) 27 p

CSCL 20D

NO0-203_5

unclds

G3/34 0772346

National Aeronautics and

Space Administration

NASA Technical Memorandum 102271

Analysis of Small-Scale Rotor

Hover Performance Data

Cahit Kitaplioglu, Ames Research Center, Moffett Field, California

March 1990

I_IASA

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000

SUMMARY

Rotor hover-performance data from a 1/6-scale helicopter rotor are analyzed and the data sets compared for the effects of ambient wind, test stand configuration, differing test facilities, and scaling. The

data are also compared to full-scale hover data. The data exhibited high scatter, not entirely due to ambient wind conditions. Effects of download on the test stand proved to be the most significant influence on

the measured data. Small-scale data correlated reasonably well with full-scale data; the correlation did

not improve with Reynolds number corrections.

NOMENCLATURE

Cd

cf

ct

CDo

Cp

cr

FM

K

Mtip

r,_,

R

Rt

Re

S

VW

airfoil-section drag coefficient

skin-friction drag coefficient

airfoil-section lift coefficient

rotor drag coefficient

rotor power coefficient

rotor thrust coefficient

rotor figure of merit

scaling factor

rotor hover tip Mach number

spanwise variables of integration

rotor radius

flat-plate Reynolds number

Reynolds number

equivalent blockage area of test apparatus

wind speed

rotor-blade angle of attack (pitch angle)
a rotor solidity

Subscripts

c corrected for download effects

fs full scale

ms model scale

I measured value

INTRODUCTION

Three hover tests of a small-scale helicopter rotor were performed in the facilities of the National

Full-Scale Aerodynamic Complex (NFAC). Detailed descriptions of the tests were reported elsewhere

(refs. 1 and 2). This report compares the rotor performance data sets obtained during these tests and

briefly discusses their differences, which result from the effects of (1) ambient wind, (2) the test stand

configuration, (3) the test facility, and (4) scaling.

The three tests, all of which u_ed the same model blades (see table 1 for blade characteristics) and

Rotor Test Rig (RTR), were

1. RTR/40 x 80 Test (fig. 1): This hover test of a 1/6-scale, four-bladed Sikorsky S-76 rotor was

performed in the test section of the 40- by 80-Foot Wind Tunnel. The rotor was mounted 10 ft above the

wind tunnel floor in a thrust down, wake up mode. No foam (for acoustic purposes) was applied to the

test rig. The blades did not have any boundary layer tripping devices. The performance data were measured using the RTR internal six-component balance and the torque load cell.

2. RTR/OARF I Test (fig. 2): This hover test of the same l/6-scale rotor was performed at the

Outdoor Aerodynamic Research Facility (OARF). The rotor was mounted 20 ft above the ground in the

standard thrust up, wake down mode. A foam cover (to serve as an acoustic anti-reflection treatment)

was installed on the external surface of the RTR during part of the test. During two runs, a simple

boundary layer tripping device was installed on the blades. Performance data were measured using three

vertical load cells installed directly below the RTR, and a torque load cell. Data were obtained over a

range of ambient wind conditions.

3. RTR/OARF II Test (fig. 3): For this test, the 1/6-scale rotor was mounted in the same thrust

down, wake up mode as in the 40 x 80 test, 10 ft above the ground. The foam covering was in place

during the entire test, but no boundary layer tripping device was used. Performance data were measured

using three vertical load cells, again installed directly below the RTR, and a torque load cell. Data were

obtained over a range of ambient wind conditions.

2
EFFECT OF AMBIENT WIND

During both OARF I and OARF II tests, data were obtained over a range of ambient wind conditions

from 0 _sec to approximately 3.6 m/see. A previous full-scale rotor test performed outdoors (ref. 3)

indicated that ambient winds of less than 1.3 m/see have only a small effect on rotary wing performance.

Thus, the data were sorted such that those obtained at ambient winds of less than 1.3 m/see were flagged

as low-wind data, and those obtained at greater than or equal to 1.3 m/see were flagged as high'wind

data.

Figures 4 and 5 compare rotor performance data obtained at low- and high-wind conditions during

__: the two tests. For the data in these figures, an acoustic anti-reflection foam cover was in place around the

!-!: _ test rig; no boundary layer trip devices were in use. In each figure, second-order polynomials are fitted

to the data. These curves indicate a small (<5%) difference between the low-wind and high-wind data;

however, in view of the significant data scatter evident in the figures, little influence on rotor perfor-

_¢_0 f amblent wind is-indi_-cated when the wind speed is below the maximum 3.6 m/sec encountered.

: A detailed examination - o fthedata shows that approximately the same amount of data scatter is present

at both low and high winds, so it is not clear that the scatter is entirely the result of ambient wind. Data

presented in reference 3 demonstrated that data scatter is reduced when a wind-correction procedure is

applied; however, for winds less than 3 knots (1.5 m/sec) only a 3% correction in the rotor power coefficient was obtained. (The 40 × 80 test data, obtained with no ambient wind, also displays scatter; however, this is most likely due to recirculation effects within the enclosed test section.) It should be kept in

mind that these performance plots are a measure of the overall behavior of the rotor, not its detailed

aerodynamics.

Corresponding aeroacousfic data show significant variations with ambient wifid. Figur_:6_tira_e_'_ _ .....

the acoustic differences arising from different ambient wind conditions. A prominent feature is the presence of high-amplitude subharmonic spectral components (labeled 1-6 in fig. 6(b)) at the high-wind

condition;-these are completely absent at low winds. (Note that the "high-wind" condition here is

0.67 m/sec, only half the i.3 m/sec criterion stated previously.) The subharmonics gradually merge with

the blade-passage harmonics as the frequencies increase. The blade passage frequency (BPF) harmonic

peak levels are up to 2 dB higher at high ambient winds than at low winds, and there is noticeable frequency broadening. The fact that no significant performance differences were observed between highand low-wind tests while acoustic spectra differences were clearly present implies that unsteady blade

surface-pressure data, rather than mean blade pressures based on overall rotor performance, are required

for accurate noise predictions.

. - EFFECT OF THE TEST STAND CONFiG_URAT[_- -_:,-::s_-_--- -_-

The OARF I and II tests Were performed with the same rotor at the same facility but in different ........

=

!

i

5

configurations. In bot_ _estsl the rotor thrust was measured with three load cells mounted so as to

transmit only vertical forces. The load cells were located between the RTR base and the top of the

mounting tower, approximately 2 m below the rotor hub. Check-load calibration during testing demonstrated the high accuracy of the data obtained using this arrangement.
In theOARFI test the rotor had an unobstructed inflow, but the wake impinged on the test hardware

located below the rotor. Therefore, during the OARF I test the thrust measurements included download

effects. In the OARF II test, the wake was directed to open air, not impinging on the test hardware.

Although the flow into the rotor passed over test hardware of significant blockage area, this was

assumed to have minimal impact on rotor performance because of the low inflow velocity.

To correct the OARF I data for the download effect, the drag force on the mounting base was estimated (ref. 4, p. 3-16), based on wake velocity calculated from simple momentum theory. The corrected

rotor thrust coefficient is given by

(1)

where S is the blockage area of the mounting base and fairing, and CD0 is the drag coefficient of a

rectangular cross-section normal to the flow. The value of the drag coefficient (CD0 = 1.18) is approximately constant for the aspect ratio and the range of Reynolds numbers encountered. The rotor power

coefficient is assumed to be unaffected and the rotor figure of merit is corrected according to the

definition

(2)

Figure 7 compares performance data obtained during the two tests. Both sets of data are for low

ambient wind conditions (<1.3 m/sec) with the foam covering in place; boundary layer trips were not in

use. Clearly, the download corrections lead to a significant improvement in correlation between the two

sets of data. The thrust coefficient at which the peak figure of merit occurs appears to shift slightly

between the two tests. Figure 7 illustrates the basic validity of download corrections, and the data in the

figure could be used in future tests of similar configurations to make download corrections to data.

EFFECT OF THE FACILITY

Previous tests have indicated that enclosures, even when large compared to rotor size, influence rotor

performance. Piziali and Felker (ref. 5) reported the presence of wake re-ingestion during a hover test of

a rotor in a chamber having height and cross-section dimensions of 15 and 6 rotor diameters,

respectively.

Two tests conducted in this series, the 40 x 80 test and the OARF II test, provided an opportunity to

examine this effect. The tests differed from each other in two respects: (1) force measurements during

the OARF II test were made with three load cells, whereas the forces during the 40 x 80 test were measured with the RTR internal balance; and (2) the OARF H test was carded out in an open environment
with minimal wakerecirculation,whereasthe40 × 80testwasin aclosedchamber,albeitlargerelative

_otoi:_S]_.w_ch ma_yh_iveld to wakerecirculation.

-: - - OfDifferent Force Measurement Methods ..... _ _:_--_=-_:-:

- - The RTR internal balance, used to measure the aerodynamic forces on the small-scale rotor during

the 40 × 80 hover test, is subject to interactions between its various components. Empirical corrections,

based on static calibration data, are normally applied to measured data to account for interaction effects.

_ _Static calibratibri0-f-th-e balance indicates that torque/thrust interaction is the largest of these effects; it

also indicates s_0ng_stel:esis behavior under combined-load conditions. A satisfactory explanation of

:-:{ _ the hysteresis i-s not presently available. When the rotor is rotating, these interaction and hysteresis

=-- effe__ch_ge; a dynamic calibration of the internal balance with a rotating load on the rotor head

"==_ _ _hasno(l_ehp_ed to date.

Figure 8 compares the data obtained in the two tests. Two sets of 40 × 80 dafa areshown: ofie set

with the combined-load interaction corrections, one set without the corrections. Note that these data

include enclosure effects, discussed below, as well as differences that may arise from different force

measurement methods. It is not immediately clear whether combined-load corrections should be applied

to the balance data. The curves indicate that combined-load interaction and hysteresis effects may not be

-:--: _ as large under rotating loads as under static loads.

- _ Effect of Test Enclosures

• As mentioned before, the two tests were carded out in different environments. In the 40 × 80 test, the

wake may have been reingested by the rotor. This was not the case for the OARF Ii tes t, where the wake

was directed into the open atmosphere with no obstructions.

=

A hovering rotor experiencing wake reingestion can be considered to be in vertical flight at low

= speed. Basic helicopter theory (ref. 6) shows that additional rotor power is required for vertical velocity

at essentially fixed-th_st. Therefore, with wake reingestion one expects lower rotor performance than

.__..... when wake reingestion is not present; i.e., for the 40 x 80 test, the rotor power curve should be higher

: ::: := :: and _he figure _fmerit Curve should be lower than for the OARF II test. Figur e 8 shows that this is the

caste When the 40 x 80 data are not corrected for combined-load interactions. However, when the 40 X 80-

data are corrected for combined-load interactions, the opposite trends are observed.

_= ..... Shiv_anda (ref. 7) noted a degradation of rotor performance when testing in a small chamber, in

!i! i- i::_(ordahce _;vS_ffs_-theory, although it should be noted that the test chamber in Shivananda s experiment was not totally closed, there being some small wnt holes both above and brow=the rotor. P_ail ......

and Felker (ref. 5), on the other hand, who tested a small rotor both in a large chamber and outdoors,

observed a small improvement (-1.5%) in performance when testing inside the chamber.

Given these uncertainties, the rotor performance data in figure 8 cannot be used to determine if recirculation was present during the 40 × 80 test.

5
.................. RE-YNOLDS NUMBER AND SCALING EFFECTS

During the three small-scale tests, the range of blade-tip Mach numbers was typical of full-scale

helicopter rotors, so the Reynolds numbers differed by the scale factor of the model blades. The

Reynolds numbers, based on blade chord and tip speed, that were encountered during these tests are of

the order of 106 and are smaller inboard of the tip, thus failing in the critical boundary layer transition

region. Full-scale rotors, on the other hand, typically operate at fully turbulent Reynolds numbers of the

order of 107. Therefore, there is a possibility of laminar separation of the boundary layer on the smallscale blades, particularly at inboard stations.

This can have important effects on both the performance and the acoustics of the rotor. If the laminar _ -

boundary layer remains attached, skin-friction drag and rotor power wj!l_ lo_e_than for a full-scale

rotor. If, on the other hand, there is laminar separation, the large high-pressure separated region could .... :_: _ :

cause a sharp increase in drag and power. Also, laminar vortex shedding is an important source of noise.

Boundary Layer Trips

During the OARF I test, several data runs were made with simple devices installed on the top and

bottom surfaces of the rotor blades for the purpose of modifying the boundary layer characteristics.

These tripping devices consisted of plain cellophane tape 0.005 cm (0.002 in.) thick with their forward

edges located 0.2 blade chords behind the leading edge of the rotor blade.

Because the rotor blades did not contain pressure instrumentation, and flow visualization was not

available, there was no way to ascertain either the flow characteristics over the blades or the effect of the

trip strips. However, examination of the performance data with and without the trip strips gave a good

indication of their effect.

Figure 9 shows a discernible loss of rotor performance (at low-wind conditions) with the addition of

the trip strips. When the trip strips were applied, the drag increased, as indicated by the higher rotor

power coefficient curve. With reference to the well-known skin-friction drag plots (fig. 10 and ref. 4,

p. 2-6), it can be inferred that the boundary layer over the blade was probably not experiencing laminar

separation; if it were, tripping the boundary layer would lead to reattachment and an improvement of

rotor performance. On the contrary, the addition of the trip strips caused a large enough increase in surface roughness to lead to a jump to the "turbulent" boundary layer curve. This, in turn, led to an increase

in the skin-friction drag, which resulted in the measured loss of performance.

In view of these observations, it is hypothesized thai without boundary layer trips, these small-scale

rotor blades were operating in the laminar (or perhaps transition) but unseparated regime. With the addition of boundary layer trips, the--fl0w jumped to the fully turbulent regime, where full-scale rotors operate, but with an effective skin-friction drag component much larger than that of full-scale rotors.

6
Comparison with Full-Scale Data

For comparison, data was used from a hover-stand test performed by Sikorsky Aircraft (ref. 8), of a

full-scale S-76 rotor similar to that used during this series of small-scale tests. The small-scale tests

employed blades with a rectangular tip planform, whereas the Sikorsky test blades had the standard

_:--_-_-: :: : svi_t:fa_er_.J:i tips, but this difference in planform should have minimal effect in hover since the

maximum Mach numbers are lower than transonic. Also note that 3% corrections were applied to the

full-scale test data to account for test stand interference and ground effects. Of the three small'scale

' tests, the OARF I test has the most similarity to the full-scale test; both were performed outdoors in a

- flarust up, wake down mode with the rotor wake impinging on the test stand.

__Z- _ _ :::: _=: - .......

..... Differences in Reynolds numbers between the small- and full-scaieiests- _e:dx-p_tedto influence

the data significantly. Yamauchi and Johnson (ref. 9) have presented an extensive study of rotor airfoil

scaling in which they developed easily implemented scaling laws to account for Reynolds number

effects in helicopter airfoil analyses. These can be extended, in a straightforward manner, to rotors (see

appendix). The rotor thrust and power coefficients are scaled as follows:

_'_ .... -:= (CT_ =K(CT) _ -K_. O )ms !_U5--_:: _:?:::: \ ff Jfs \ ff /ms fs

_:-::-_:;--_: ; where

(4)

The rotor figure of merit in hover is defined by equation (2) on page 4. Combining equations (2) and (3)

to calculate the rotor figure of merit scaling relation gives

K5/2._ fs FMms fs: / -Zms (5)

Figures 11 and 12 show the effect of making these Reynolds number corrections, and figure 13

compares the two OARF data sets corrected for Reynolds number effects with the full-scale data. (The

OARF I data were also corrected for download.) No improvement in the correlations among the three

data sets is noted with the Reynolds number corrections. Again, the data scatter hinders a more definite

conciusion.

CONCLUSIONS

A 1/6-scale helicopter rotor was tested in hover, in a wind tunnel and outdoors, in several different

:: _' _nfigurations. The influence on rotor performance of the test facility and test configuration were inves-

[:=# =_:s= .:#:s :5 (

7
tigated.Theeffectsof ambientwind andReynoldsnumberwerestudied.Thefollowing canbe

concluded:

1. Measureddatadisplayedhigh scatterthatis notentirelydueto ambientwinds.

2. Ambient winds had only a small effect on rotor performance. However, the data scatter may be

masking more significant effects of wind. Acoustic data, more sensitive to rotor unsteadiness, showed

definite variation with small changes in ambient wind conditions.

3. Download effects of the wake on the test stand led to about a 15% change in the rotor thrust

when the rotor was in the thrust up, wake down configuration.

4. The effect of the test facility could not be clearly discerned because of different forcemeasurement techniques used during different tests. Also, the RTR balance requires careful calibration

under rotating loads.

5. Rudimentary boundary layer tripping increased skin-friction drag and resulted in loss of performance, rather than alleviating laminar separation (if present) to improve performance.

6. Small-scale data, although quite close to full-scale data generally, exhibit significant differences.

The correlation did not improve with Reynolds number corrections.

8
APPENDIX

DERIVATION OF ROTOR REYNOLDS NUMBER SCALING RULES

In general, the rotor thrust can be calculated from the airfoil section lift coefficient, c t, by integrating along the rotor radius. The resulting rotor thrust coefficient is given by

CT_ - _ 1 if2 ct(r ) •r 2 dr (A 1)

.... Similarly, the rotor power coefficient can be obtained from the airfoil drag coefficient, cd (not considering induced drag, which should not be influenced by Reynolds number):

cp= ._ cd(r ) . r3dr (A2)

::::_ : effects as follows (ref. 9):

The drag and lift coefficients of helicopter rotor airfoil sections are scaled for Reynolds number

ct = Kct I --ff (A3)

- Cdt (A4)

where the subscript 1 indicates a measured value and no subscript indicates a scaled value; the measured

---v lift coefficient, c t , is a function of a scaled pitch angle given by 0q/K; and the scale factor, K, is

" _ defined by K = (l_e/Rel) 1/5.

: 2 .i

..............................

In the present study, the measured values were obtained with a Small-scale model, and w e a_re interested in calculating the corresponding full-scale values for comparison with measured full-scale data.

i =_= --=-/ : Accordingly, substituting equation (A3) into equation (A1) and using the geometric scaling

- transformation

r = K5_

_:, gives the Scaling relation for the rotor thrust:

1 [ Rms

(C_-C_)fs = 2R_ms J0 KC/ms_2d_

(A5)

(A6)

9
.... Usingihe generaldefinhionof thrust coefficient (eq. (A 1)) set up for model scale, this becomes

fs k. c Jms

(A7)

where the model-scale quantity (CT/O)ms is evaluated at the scaled pitch angle O_ms]K. Similarly, the

scaling relation for the rotor power coefficient becomes

= _'k"_-jm s

(A8)

10
REFERENCES

1. Kitaplioglu, C.; and Shinoda, P.: Hover and Forward Flight Acoustics and Performance of a SmallScale Helicopter Rotor System. 11 th European Rotorcraft Forum, Paper 98, London, Sept. 1985.

2. Kitaplioglu, C.: Scaling of Helicopter Main Rotor Noise in Hover. AIAA Paper 86-0393,/an. 1986.

3. Felker, F. F.; Maisel, M. D.; and Betzina, M. D.: Full-Scale Tilt-Rotor Hover Performance.

J. American Helicopter Soc., vol. 31, no. 2, Apr. i986, pp. 10-18.

4. Hoerner, S. F.: Fluid Dynamic Drag. Published by the author, Bricktown, NJ, 1965.

5. Piziali, R.; and Felker, F: Reduction of Unsteady Circulation in Hovering Model Helicopter Rotor

Testing. J. American Helicopter Soc., vol. 32, no. I, Jan. 1987, pp. 54-59.

6. Gessow, A.; and Myers, G. C., Jr.: Aerodynamics of the Helicopter. F. Ungar, 1952, p. 83.

7. Shivananda, T. P.: Pressure Measurements Near the Tip of a Hovering Model Rotor and a Preliminary Investigation of the Flow in the Rotor Wake. PhD Thesis, Georgia Institute of Technology,

School of Aeronautical Engineering, Dec. 1977.

8. Jepson, D'; Moffitt, R.; Hilzinger, K.; and Bissell, J.: Analysis and Correlation of Test Data From

an Advanced Technology Rotor System. NASA CR-3714, 1983.

9. Yamauchi, G. K.; and Johnson, W.: Trends of Reynolds Number Effects on Two-D_m_aTAir _:=_ _ _

foil Characteristics for Helicopter Rotor Analyses. NASA TM-84363, 1983.

11
Table 1.- Modelrotor characteristics

Scale......................... 0.16of full-scaleS-76

Radius................................................ 107m

Chord................................................ 63cm

Airfoil ...........................................SC1095

Numberof blades.....................................4

Twist..................................................... -10°

Sol!dity ................................................ 0.075

r

.m

12
ORIG tI',L_I_ PAGE

BLACK AND WHITE P.HOT.OGRAP.H

Figure 1.- Rotor Test Rig set up for the 40- by 80-Foot Wind Tunnel test.

13

ORIGINAL PAGE IS

OF POOR QUALITY
ORIGtNA[ PAGE

BLACK AND _HITE PHOTOGRAPH

¥

Figure 2.- Rotor Test Rig set up for the OARF I test.

14
ORIGTNAL PAGE

BLACK AI,_D '_#HITg Pi-iOFOGR_PH

Figure 3.- Rotor Test Rig set up for the OARF II test.

15

. =;==W
.8

.6

0

-0.2

(.)

/ _ Low <1.3

7 >1.,,,,.,,, -- _' High 3

.015

_.010

=

.005

o/ °A

B

0 I l _ I • I t I J I J I l .

-.02 0 .02 .04 .06 .08 .10 .12

Figure 4.- Effect of wind on rotor performance--QARF I test; (a) figure of merit,

(b) power coefficient/solidity.

16

I

i

I

i

!

l

I

I

!

!

!

i
I

,8

.6

I .4

0

Eor, .2

i*

-.2

(=)

t_ _ _ Low <1.3

¢ ---@High >;.3

.015

(b)

[]

o.010

J

i

.005

a.

o

0 i I i I -_ I , 1__ J-_

-.02 0 .02 .04 .06 .08 .10 ,12

Coefficien/of thrus_soi!di_ ...............................................................

Figure 5.- Effect of wind on rotor performance (3ARF ii test; (a) figure_of me_t, ...............

(b) power coefficient/solidity.

]7
130

120

_110

_100

--1

(a)

90

..... V w = 0.04 m/sec

-- V w = 0.67 m/sec

70 _.,r_ "--

LJ t._

6O

50 f

31.5 125 500 2000 8000

1/3-Octave Band Frequency, Hz

130

120

110

.u 100

.=

= 90

_ 8o

"_ 70

0

_ 60

60

(b)

..... V w = 0.04 m/sec

-- V w = 0.67 m/sec

i_ i Subharmonics

ji 2 3 . BPFHarmonics

0 .250 300 .750 1.000 1.250

Frequency, k Hz

¢D

=,

Q.

"0

f130 (c)

120

110

-_ - V w = 0.04 m/sec

100 V w = 0.67 m/sec

9O

8O

60

50

0 1.0 2.0 3.0 4.0 5.0

Frequency, k Hz

Figure 6.- Effect of wind on rotor acoustics OARF I and II tests. Microphone is located 30 ° below the

rotor plane at two rotordiameters from the hubl Mtip = 0.55, CT/_= 0.06. All sound pressure levels

are given with respect to 2xlO -5 N/m 2. (a) l/3-octave-band spectrum; (b) narrow-band spectrum,

5.97-Hz bandwidth; (c) narrow-band spectrum, 23.9-Hz bandwidth.

18

i
o

E

E

-i

ii

.8

.6

.4

.2

0

--0,2

° ° /

j(jJ _out download corrections

_ith download corrections

-- - - 4) OARF II

___-._._ I I ! ' ! I I L-- i | I I

.Ol 5

.010

"5

&

U

E

.oo5

D.

o

(b) m_/.i/_ !

0

0 .02 .04 .06 .08 ,10 ,12

Coefficient of thrust/solidity

Figure 7.- Effect of test stand configuration on rotor performance; (a) figure of merit,

(b) power coefficient/solidity.

H

-7

19
.6

"_ .4

E

•-_ .2

U.

-0.2

(a)

/_ m OARF II-Lold cells

.L// _ ---- O40xSOBllence, wlth combinedload corredlon* ....

/_Ba ance, without combined-

- IO,,uco.ec,on.

ir_ I i 1 J 1 i I J 1 I I i

.015

(b)

I'11"_

iOlO yy

_ .11115 j []

G

• B

_ _'_-_ •

0 ----_ I 1 I ±_. 1 i i i I 1 J i

-.02 0 .02 .04 .06 .08 .10 .12

Coefficient of thrust/solidity

Figure 8.- Effect of test facility on rotor performance; (a) figure of merit,

(b) power coefficient/solidity.

1

i

20
;1¢

i

E

"6

g.

.8

.6

.4

.2

-,2

(a)

/_ -- i3 Wlth°ut trip

J ......... --- _>I Withi trl_ ,

i

|

,015

"0

= .010 o

e-

.005

L

0

-,02

(b)

......-l__ll[l{l[I

0 .02 ._ ._ .08 .10 .12

Coefficient of thrust/solidity

Figure 9.- Effect of boundary layer nip on rotor performance OARF I test; (a) figure of merit,

(b) power coefficient/solidity.

,01

.001

104

_ \.%

Cf

-.. Forced Turbulence

Blasius "_ "-_/x_o, '

_'_ 1000

• /. -. o ....

2000/ i-i ,_i ,

. " k = 4000 ""'_--.. Schoenherr

I l I ', i l _l I i _ t I I

10 5 10 6 10 7 10 8 10 9 1010

R_

Figure 10.- Skin-friction drag coefficient of smooth and plane surfaces in incompressible flow (ref. 4).

21
1.2

.h,

,4

E

'3

0

U,,

-.4

--0.8

(a)

l _ Corrected for Reynolds number

l_d for Reynolds number

I /,_/ ---- e Full-scale (Sikorsky) ///, , .....

.015 /// (b) _//

_> ./ /_/

,/ /

_ /

oj_j

i°°'

-.05 0 .05 .10 .15

Coefficient of thrust/solidity

Figure 11.- Effect o_ Reynolds number corrections--OARF I test; (a) figure of merit,

(b) power coefficient/solidity. |

i

22
1.2

.8

Z

Ii .4

E

tw

_ 0

-.4

-0.8

{.)

/,_rrected for Reynolds number

/.I/_ ---- _ Uncorrected for Reynolds number

//_'._(Sikorsky)// ra -- - - @ Full-scale

,L , I .. I I J,

.015

1= .010

"0

E

i .11115

#

(b) i///1.1"

/_//i

,/

f I __ I .... i I i

-.05 0 ,05 .10 .15

Coefficient of thrust/solidity

Figure 12.- Effect of Reynolds number corrections--OARF II test; (a) figure of merit,

(b) power coefficient/solidity.

23
1.5

1.0

.5

E

"6

.__ 0

tl

-.5

-1.0

(a)

IP

,0"_ -- • OARF I corrected for download

./_ and Reynolds number

OARF II Corrected for

._' _> Reynolde number

/,_ " -" O- Full-ecale (Sikorsky)

.010

.008

Z"

_.006

LI

E

_.004

.002

I/"

(b) ;.

I.I

0

0.

•

/ !

-.05 0 .05 .10

Coefficient of thruePsolidity

i

.15

Figure 13.- Comparison of full-scale data with small-scale data corrected for Reynolds number effects;

(a) figure of merit, (b) power coefficient/solidity.

24
Report Documentation Page

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N alo*"_d A_oneu_os

$i;_ c_ Adraili1_'atim

1. Report No.

NASA TM- 102271

4.

2. Government Accession No.

Title and Subtitle

Analysis of Small-Scale Rotor Hover Performance Data

7. Author(s)

Cahit Kitaplioglu

9. Performing Organization Name and Address

Ames Research Center

Moffett Field, CA 94035-1000

t2. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546-0001

3. Recipienrs Catalog No,

5. Report Date

March 1990

6. Performing Organization Code

8. Performing Organization Report No.

A-90046

10. Work Unit No.

505-61-51

11, Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15, Supplementary Notes

Point of Contact: Cahit Kitaplioglu, Ames Research Center, MS T042, Moffett Field,

CA 94035-1000

(415) 604-6679 or FTS 464-6679

16. Abstract

Rotor hover-performance data from a 1/6-scale helicopter rotor are analyzed and the data sets compared for the effects of ambient wind, test stand configuration, differing test facilities, and scaling. The data

are also compared to full:'scale hover data. The data exhibited high scatter, not entirely due to ambient wind

conditions. Effects of download on the test stand proved to be the most significant influence on the

measured data. Small-scale data correlated reasonably well with full-Scale data; the correlation did not

improve with Reynolds number corrections.

17. Key Words (Suggested by Author(s))

Helicopter

Rotor

Aerodynamics

19. Security Classif. (of this report)

Unclassified

18. Distribution Statement

Unclassified-Unlimited

Subject Category - 02

20. Security Classif. (of this page)

Unclassified

21. No. of Pages

26

22. Price

A03

NASA FORM 1626 OCT86

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