Effect of twist on helicopter performance and vibratory loads

THIRTEENTH EUROPEAN ROTORCRAFT FORUM

J..j

Paper No.

m

EFFECT OF TWIST ON HELICOPTER

PERFORMANCE AND VIBRATORY LOADS

Charles Keys

Frank Tarzanin

Frank McHugh

BOEING HELICOPTER COMPANY

PHILADELPHIA, PENNSYLVANIA

USA

September 8-ll , 1987

ARLES, FRANCE

EFFECT OF TWIST ON HELICOPTER PERFORMANCE AND VIBRATORY LOADS

Charles Keys

Engineering Specialist, Aerodynamics Boeing Helicopter Company

Frank Tarzanin

Supervisor, Dynamics Boeing Helicopter Company

Frank McHugh LHX Technology Manager Boeing Helicopter Company

1. 0 ABSTRACT

High twist is desirable to provide improved hover, vertical climb and nap-of-the-earth performance capability; however, the high negative angles of attack encountered on the advancing blade tip can adversely affect forward flight performance and vibratory loads. To quantify these effects, a 10 ft. diameter model rotor was evaluated in the Boeing Vertol low speed wind tunnel. This paper summarizes the test results and makes comparisons with theoretical predictions.

2. 0 NOTATION b = Number of Blades c = Thrust-Weighted Chord, ft = Rotor Rotor Rotor Rotor Thrust Coefficient, T/pnR2_VT2_crT Power Coefficient, P/pnR2vTs = De

=

Diameter, ft Effective Drag, v - x ) , l b p L = Rotor Lift, lb

M1 so = Advancing Blade Tip Mach Number P = Rotor Power, ft lb/sec

j5 q

= =

Nondimensional Power, P/qD2aTV Dynamic Pressure, ~pv2

R = Rotor Radius, ft

V = Forward Flight Speed, ft/sec VT = Tip Speed, ft/sec

X

=

Rotor Propulsive Force, lb

X

Rotor Propulsive Force Coefficient, X/qD2crT = Advance Ratio, V/VT

= Nth Flap Natural Frequency = Nth Chordwise Nat~ral Frequency wNT = Nth Torsional Natural Frequency p

=

Air Density, lb sec2/ft4

crT = Rotor Thrust-Weighted Solidity, b c/nR

Subscript

REF = Referred to baseline level (CT/crT = 0.08 or 180 knots for twist= 11.5°)

3.0 INTRODUCTION

A review of available literature indicates that increasing twist improves hover and low speed performance. Higher twist results in a more uniform downwash velocity in the far wake and a corresponding decrease in induced power required as described in the various texts on rotary wing performance. In forward flight very little data has been pub-lished on the effect of blade twist. Reference 1 indicates that higher blade twist reduces forward flight power required based on flight test data published in 1948. The speed range of these early aircraft was less than 70 kts. as compared to the design speeds for 1980/1990 helicopter designs of 180 kts to 200 kts.

A test was conducted to quantify the effect of twist on performance and aircraft vibra-tion on an existing 10 ft. diameter rotor evaluated previously as described in Refer-ence 2. The rotor was a four bladed rotor with Mach scaled composite blades designed to be dynamically representative of a typical full scale rotor. A sketch of the blade is presented in Figure 1. It had a thrust-weighted solidlty (aT) of 0.1383 and a tip taper of 0.6 starting at 95% of the rotor radius. The last 5% bf the blade tip had a quarter chord sweep of 30° aft. The twist of the blade initially tested was approxi-mately 11.5° linear, as shown in Figure 2. Since the blades were fabricated from C0m-posite materials i t was possible to retwist the blade to a twist of 17.3° from 11.5°

by a process developed by the Boeing Vertol Company that has been used successfully in the past to retwist model rotor blades. There is a small deviation in twist between the various blades, as illustrated by the bands in Figure 2, but it was no greater than the tolerance when originally fabricated.

The measured blade natural frequences at the nominal test rotor speed are presented in Table 1. As notej:l the frequencies are 0. 25 to 0. 5 per rev higher or lower than any even multiple of rotor RPM. Since both twist values were tested on the same blade, differences in measured vibratory characteristics are due entirely to twist and not variations in the rotor blade properties .

• 228R .28R .85R .95R

I I

CHORD= 6.77 IN.

I

11.0R aT =.1383

-1E=E~~~~~~' ~~~~~~

I

RADIUS 5.05 FT

~

- -

9;

30• SWEEP '\,.- VR12 VR15-TAPER RATIO= .6 '\,. FLAP HINGE= 2.86% R '---~---~

B

UJ

9

....

!12

;::

....

Figure 7. Planform of Model Rotor Blade

BASELINE BLADE ( 11.5° LINEAR TWIST)

/

RETWISTED SLADE ( 17.3° LINEAR TWIST)

/

-~ ORIGINAL TOLERANCE _ ON TWIST 0

L _ _ L_ _ _JL__....l. _ _ ...L.. _

___J=]

VARIATION WITH RE-TWISTING

0 .2 .4 .6 .8 1.0

RADIAL STATION (r/R)

Figure 2. Model Blade Twist Distributions

Table 1. Blade Natural Frequencies

at

Nominal Rotor Speed

f.U:!F

WIF w4F W1c W2c w1T

PER PER PER PER PER PER

RIN RIN REV REV RIN RIN

4.0 TEST STAND/AIRFRAME

The model blades were tested on the powered model test stand shown in Figure 3. The rot?r and con~rols were m?unted on a five-component rotor balance. The rotor hub was

art~cu~ated Wlth a flap h1nge at 2.86 percent of radius and a lag hinge at 15 percent

of rad1us. ~h~ rotor was powered b¥ three 130 HP air motors located in a pod below the rotor and ~r1v1ng through a reduct1on gearbox. A fuselage shell was mounted on a six

component l!lternal fu~elage balance ~ttached to the chassis. The fuselage shell was represen~at1~e of a s1ngle rotor hel1copter. The rotor shaft angle, blade collective and cycl1c p1tch were remotely controlled.

Power Pod

I

Figure 3. Test Stand Configuration

5.0 INSTRUMENTATION

The model had a five-component main rotor balance and a six-component fuselage balance located as noted in Figure 3. The rotor balance measured rotor thrust, propulsive force, side force, pitching moment and rolling moment. Rotor torque was measured using strain gauges installed on the rotor shaft. Fuselage drag, lift, side force, pitching, rolling and yawing moment data was obtained from the fuselage balance.

One rotor blade was instrumented with flap, chord and torsion bending strain gauges distributed along the blade radius as shown in Figure 4. The gauges were installed inside the blade on the spar prior to final assembly to avoid the aerodynamic penalties associated with surface mounted gauges. Blade flap and lag motions were measured by transducers positioned on the flap and lag hinges of the instrumented blade. The pitch links were instrumented to measure control loads and blade collective and cyclic angle were obtained from swashplate position.

The rotor balance was calibrated statically using weights and pulleys to measure steady forces and moments. A dynamic calibration was obtained to measure 4 per rev vibratory hub loads. in the fixed system at the normal operating rotor speed. The calibration

consisted of shaking the model at 4/Rev through a fixture attached in place of the hub.

The shake fixture had bearings to penni t conducting the shake tests with the rotor shaft rotating in order to include the proper shaft dynamics. The balance sensitivi-ties were then assembled in a dynamic calibration matrix and used to determine

vibrato-ry hub loads during the testing.

FLAP .12R.19R CHORD .12R TORSION .30R .5 .83R .4SR .40R .52R .67R .85R

6.0 TEST CONDITIONS AND PROCEDURE

A summary of the test conditions is presented in Figure 5. Forward flight testing con-sisted of speed sweeps at rotor thrust coefficients (CT/oT) of 0.067 and 0.08 keeping the rotor trimmed at each point for zero one-per-rev flapping and maintaining a

propulsive force coefficient (X/qD2o'T') of 0 .1. Thrust sweeps were obtained at an

advance ratio of 0.4 and a propulsive- force coefficient of 0.1. The range of thrust

coefficient was. limited by model 'power available or blade loads. Propulsive force

sweeps and Mach number sweeps at a rotor thrust coefficient ( C /a ) of 0. 06 7 and an advance ratio (~) of 0.4 were also obtained to evaluate rotor pro~llive efficiency and compressibility effects. Hover out-of-ground effect testing was performed by conduct-ing thrust sweeps.

>-ffi

.12 -§ u. u. ~- .08 . . . ~ . . . u >-... ~

"'>-~g

.04 f-

X

VARIATION/ M1.00 VARIATION >-a: 0 >-0 a: oL---~'---~~--~'----L-1~1 0 .1 .2 .3 .4 .5

ADVANCE RATIO (IJ)

Figure 5. Test Conditions 7.0 PERFORMANCE TEST RESULTS

7.1 Hover

Effect of Twist on Power Required- Increasing the blade twist from 11.5° to 17.3° re-duced the power requ~red as shown in Figure 6. At a typical design C /a of 0.08, the power required reduction is 2.4% or 0.018 in figure-of merit as noted In ~igure 7. For a typical 10,000 lb. single rotor helicopter this improvement corresponds to a 160 lb increase in hover gross weight capability or a 5% increase in useful load.

Effect of Twist on Download - Increasing the twist resulted in an increase in hover download as shown ~n F~gure 8. A higher twist redistributes the downwash velocity cre-ating higher downwash inboard where the fuselage width is maximum. At a CT/aT of 0.08 the 17.3° twist rotor had 6% higher download than at 11.5° of twist. Fo:r

a:

typical single rotor helicopter with a download to thrust of 3% to 4%, 6% corresponds to an increase in download of 20 lb for a 10,000 lb aircraft. The net ben~fit of twist would be 160 lb due to Figure of Merit improvement minus the 20 lb download penalty for a net benefit of 140 lb. >- 1.76

z

w § _1.60 u. u. w 0 _1.26 u a:~ w ~ 3:: ~ 1.00 0 a. a.U

-a: tl. .76

g3

11.5°TWIST

\

\

17.3°TWIST .08 CT/aT 0 a: 0 w a: a: w u. w a:

I

.20

I

0 ' - - - - . . . _ _ _ _ - ' - - - ' - ' - - - - ' - - - - ' 0 .0004 .0008 .0012 .0018 .0020

ROTOR THRUST COEFFICIENT (CT 3/2)

. Figure 6. Effect of Twist on Hover Power Required

1: a: w :=; u. 0 w a: ::J~ 1.06 1.00 ~

li!

.90 ·CC

l

o-6l

.86 a:~ 0 w a:

ffi

u. w a: .so .70 2.4% INCREASE (.O 18 INCREASE IN FM)

l

-17.30TWIST

\/1--/\

I

11.5° TWIST

I

I

I

I

I

TYPICAL DESIGN CTiaT .02 .04 .06 .08 .10 .12

ROTOR THRUST COEFFICIENT (CT/aT)

16

TYPICAL DESIGN CT/OT

---

1

I

oL-~---L--~~1-L-L~~

0 .02 .04 .08 .08 .10 .12

ROTOR THRUST COEFFICIENT (CTiOT)

Za: Qw 1.20

s~

1.1. za.~

w.-

~1.12 ::;z o g~~ 1.08 <~-t-o 1.04 ~0 a: 1- 1.00 :I:<

DATA SOURCE: REFERENCE 3

1- _{o .4 .a 1.2 1.a 2.0 2.4 2.a a.2 a.e}

HEIGHT OF ROTOR ABOVE GROUND (Z/R)

Figure 8. High Twist Increases Hover Download Figure 9. High Twist Reduces Hover IGE Thrust

Augmentation

Effect of Twist on In-Ground-Effect Hover Performance - Testing was not conducted

In-Ground-Effect ( IGE), however, data presented 10 Reference 3 and reproduced in

Figure 9 indicates that the thrust augmentation obtained IGE at constant power is de-graded substantially when the blade twist is increased.

7.2 Forward Flight

Effect of Twist on Power Required - In forward flight, increasing the twist from 11.5°

to 17.3

°

resulted 1n an 1.ncrease in power required as shown in Figure 10. At 180 kt (~ = 0.434, 700 ftjsec tip speed) and C~/cr~ of 0.08, the measured power increase is 5%. Assuming a power available corresponding ~o a CP of .0013, the performance penalty due to twist is approximately 4 kt.

The corresponding L/De variation due to twist is presented in Figure 11. A degradation in rotor lift to effective drag ratio of 0.11 is shown at 180 kt.

The predicted azimuthal variation in power required for rotors with the two twist val-ues tested is presented in Figure 12. As shown, the power increase is due primarily to the increase in profile power on the advancing blade. The tip of the advancing blade operates at higher negative lift coefficients {C~) at 17.3" than with 11.5" of twist. The VR12/15 airfoils are excellent high Macn number airfoils as indicated in Reference 2; however, the higher negative

cp_

at the outboard region o·f the blade h?s a high drag coefficient. For the advanced airloils, the increased power on the advancing side of the rotor disc is due more to negative C~ operation than Mach number because, as shown in Figure 13, the L/De difference does not vary significantly with advancing blade Mach numbers (M1 , 90 ).

1- CTiOT

=.oa

z

1.2 w 0 _{4 KNOT} PENALTY-;:;: u. 1.0 w 0

o_

••

a: •

17.3"TWIST w ~ ;:a.

••

'\..

oo

a.' _a.

'\.

a:o

o-

.4 1- 11.5" TWIST 0 a: Cl

••

180 KNOTS UJ

I

a:

a:

w 0

_I

u. w _{.1 .2} .3 .4 .5 a: ADVANCE RATIO (p)

Figure 10. Effect of Twist on Forward Flight Power Required Q ~ a: C!> < ~ 1.2 w 2: 1.0

b

~ -u. u.

~··

't'w

OCl 'I~

.e

tw

:Je

a: ...J .4

o-,_

~ ~ .2

2-7-5

Cl w a: 0

a:

UJ u. w a: 180 KNOTS .1 .2 .3 .4 .5 ADVANCE RATIO (p)

~ 0. J: ~3000 w

§

2000 0 w a: 1000 0:: w ;:: 0 0.

DATA BASIS: B-65 VORTEX THEORY

180 270 380 AZIMUTH (DEGJ w ~ 1.0

b

~-

.e

u. ~ w ~ I W 8

oo .

~::J

tU. .•

=:;jf2

0::-'

o-

.2

b

0:: 0 w 0 II= .4 CTiaT

=

.067 11.5°TWIST

\

0

° t::r----cJ-_

__,._0

\7.3°TWIST 0

Figure 12. Predicted Azimuthal Power Variation with Twist 0:: 0:: w u. w 0::

.78 .eo .s2 .84 .ae .as

ADVANCIIIG BLADE Tl' MACH NO. (M 1•90J

Q !;( 0::

"

<

gj

w ~ 1.0

b

w u.- .a u. ~ w • I < g~ .8 I ' !!;~ ..J ~ .4, ex:, 0-' ,_~ 0 .2 0:: 0 w 0 0::

x

=.1 11=.4 11.5°TWIST

Figure 13. Effect of Mach Number and Twist onUDe

~

0::

w .02 .04 .08 .08 .10 .12

u.

~ ROTOR THRUST COEFFICIENT (CT/OT)

Figure 14. Effect of Twist and CT/a on UDe at/1=-4

The variation in rotor L/De with twist was demonstrated during thrust sweeps as pre-sented in Figure 14 for an advance ratio (IJ) of 0.4. The penalty due to increased

twist grows with thrust coefficient. At .04 c~aT the increment in L/De is 5% and at

0.08 CT/aT the increment is

11%-Effect of Twist on Blade Stall - Because of model power constraints with the high solJ.dl. ty rotor conf1gurat1on the effect of twist on stall inception was not obtained during the 1986 testing. However, stall data for blades with 7° and 13° of twist was obtained on an earlier test for a lower solidity I three bladed rotor configuration

(Reference 4). Figure 15 shows there is no difference in the stall inception bound-aries between the 7° and 13° twist blades over the entire speed range from hover to ~

of 0.4. Stall inception is defined as the C /a where the blade root torsion or pitch link loads increase rapidly. This datu was bbtlined from thrust sweeps at a constant trim propulsive force coefficient of 0.1.

Effect of Twist and Propulsive Force on Power Required - Nond~rnt:;nsion~l Power

2 (P) is

plotted 1n F1gure 16 as a funct1on of propuls1ve force coeff1c1ent (X :;:: X/gD crT) foE

C ja ~ 0.067 and~ ~ 0.4. As noted at an

X

of 0.14 ~e 17.3' t~ist blade_has 0.008 P

high~r power required than the 11.5° twist blade. At X of 0 the 1ncrement 1ncreases to 0. 012. Data was not obtained at lower propulsive forces; however 1 as noted in

Refer-ences 5 and 61 the penalty should continue to increase at negative propulsive force

conditions approaching autorotation

(P :;::

0) due to stall related power on the inboard blade region.

-;.

~

_....

8.18

....

z w 0.16

u:

u. ~-14 0 ~ .12 ~ J: 1-.10 0:

~

0

x

=.1

C. 13°TWIST} DATA SOURCE:

e 7"TWIST REFERENCE 4

ONSET OF BLADE STALL

/

.1 .2 .3

ADVANCE RATIO (p)

.4

Figure 15. Increased Twist has No Effect on Stall Inception

8.0 VIBRATORY LOADS TEST RESULTS

.28 10: .24

-

_!Z

w

5

.20

u:

u. w 0 (,) .16 0: w

;:

~ .12 CT/oT=-067 p=.4 17.3° TWIST .04 .08 .1X 11.5° TWIST .12 .18

ROTOR PROPULSIVE FORCE <i()

.20

Figure 16. Effect of Twist and Propulsive Force on Power Required

The effect of blade twist on rotor vibratory loads was evaluated using the dynamically calibrated balance output to define 4 per rev hub vibratory loads as well as blade strain gages to measure vibratory blade loads. In general, increasing blade twist in-creased hub and blade 4/rev vibratory loads. A summary of the percentage increase in hub vibratory loads measured for the 17.3° twist blade compared to the 11.5° twist blade is presented in Figure 17 for a C /cr of 0.067 and 180 kt. The 4 per rev loads increased from 37.3 to 126% with the lar~esl increases observed for the in-plane longi-tudinal and lateral force components. This result is consistent with the increase in power required on the advancing blade. Rolling moment also exhibited a large increase

in loads. ·

A summary of the percentage increase in blade alternating flap bending, chord bending and torsion maximum peak-to-peak loads due to increasing the twist at 180 knots is pre-sented in Figure 18. The flap bending loads increased 11.3% to 32.4% with the largest increase occurring on the inboard blade station .19R location. The chord bending at 0.45R showed a 45% increase in loads and a 6% reduction inboard of the blade lag hinge. Blade torsion also increased 6% to 11%.

TWIST INCREASED FROM 11.5° TO 17.3° CT/OT=-067, VTIP= 700FT/SEC, X= .1

;;;

,g

-.,

>

180 KNOTS

~

"i

~

.5

-~

"S> c::

.3

95.4%

-c::

.,

g

:IE

"'

.5 .t: B

a:

~

E 0 :IE

"'

~

0

a:

Figure 17. Increase in 4/Rev Hub Vibratory Loads Due to Twist Increase

2-7-7

CTioT=.067. Vnp =700FT/SEC,

x

=.1 180 KNOTS FLAP BENDING MOMENT CHORD BENDING TORSION MOMENT MOMENT

f!loo

so ,....---.---!~-,---,

..

_g~

-w< 40 o-' <"'

_, ....

l!l!/1 ~~ 30 o-w <W zc: 20 O:C!l

ww

>-O

-'.,

< . 10

6;::

ww (/)J: <>- 0 WO: O:w 0> ~0_-10 1/1.

"'

cd

-BLADE STATION

Figure 18. Change in Alternating Blade Loads Due to Increased Twist

8.1 Airspeed Sensitivity

The 4 per rev hub longitudinal forces for the 11.5° and 17.3° blade twists are present-·ed in Figure 19 as a function of advance ratio. The increase in loads due to twist

grows with increasing airspeed up to an advance ratio of 0.35 and remains constant be-yond. Alternating blade chord moment at 0.45R showed an increase in load due to twist as noted in Figure 20. The alternating blade flap loads, shown in Figure 21, are con-sistently higher for the blade with 17.3 ° twist with the general trend indicating an increase in the increment with airspeed at all radial positions.

8.2 Thrust Sensitivity

The effect of rotor thrust coefficient CT/a'I' on the vibratory longitudinal hub forces is presented in Figure 22 for an advance ra~io (~) of 0.4. The 17.3° twist blade has consistently higher loads with the increment tending to grow with airspeed. The alter-nating flap and chordwise blade bending moments show trends similar to the 4 per rev hub loads.

9.0 IMPLICATIONS OF TWIST DATA FOR FUTURE ROTORS

The implication of this test is that geometric twist increases carmot be utilized to increase hover performance for advanced high speed rotors because of the inherent for-ward flight performance and vibration penalties. 1990 design requirements will proba-bly require cruise speeds of 180 kt to 200 kt. Analytical studies indicate that employing large non-linear geometric twist distributions is not effective for advanced rotors at the 180 kt to 200 kt airspeeds due to the large negative lift created on the advancing blade. One possible solution is the use of live twist to reduce the cruise penalties. Blades can have considerable live twist as described in References 7 and 8.

By tailoring the blade planform, static twist and blade _properties as suggested in Reference 7, the blade can potentially have an azimuthal variation in twist that will negate the advancing blade performance and vibration problems. Reference 8 describes some recent work utilizing sweep and tailored blade properties to minimize vibration.

Q ~ a: _{CT/OT= .067} 180 KNOTS w () _x=.• a: 0 2.0 u.. ..J < ;: 0 1.5 ::::> ~ 1:

_•

<!! ~

z

0 :1:

-1.0 ..J _u.. en _~

!i!

> w .fi a:

-

...

0 w a: 0 a: w u.. .1 .2 w .3

••

.5

a: _{ADVANCE RATIO (IJ)}

Figure 19. Effect of Twist on 4/Rev Longitudinal Hub Loads a:

"'

"": ~ _{180 KNOTS} Q 1.6

,_

_x=.1 < _a:

,_

1.4

z

w ::!< _1.2 0 ::!< w ~ ~ 1.0 0 ~

gj::!i

:1:~

.a

()::!; <!!() ; : -

_{,_}

••

<

z

••

a: w

,_

..J _.2 < 0 w a: a: 0 w u.. _{.1 .2 .3 .4 .5} w a:

ADVANCE RATIO (IJ)

Figure 20. Effect of Twist on Alternating Chordwise Moment

Q

,_

< _a:

,_

z UJ ::! 0 ::!

"

z .2 0 1.2 C5 1.0 z UJ

"'

a. c:

.a

"'"'

..J~

u....:

.6 < .4 .2 0 .1 .2 .3 .4 ADVANCE RATIO (11) .1 -2 .3 .4 ADVANCE RATIO (11) CT/OT =.067 x=.1 1.2 1.0 a: .6

"'

~

,_

.6 < .4 .2 .5 1.4 1.2 1.0 a: .6 Ol '":

,_

.6 < .4 .2 0 .5 .1 .2 .3 .4 .5 ADVANCE RATIO (11) 0 .1 .2 .3 .4 .5 ADVANCE RATIO (11) Figure 21. Effect of Twist on Alternating Flap Bending Moments

1.6 _{11 =.4} Q x = .1 ~ 1.6 a: UJ (.) a: u 0 u. ..J < 1.2 ;;:; c E~ " ~ 1.0 zu: 0;:: ..JU. m:X: .8

~-11.5°TWIST

iii

.6 a:

-....

c .4 UJ a: .08 CT/aT a: UJ .2

I

u. UJ a: 0 .03 .05 .07 .09 .11

ROTOR THRUST COEFFICIENT (CT/OT) Figure 22. Effect of Twist on 4/Rev Hub

Longitudinal Loads at J1. = .4

10.0 REFERENCES

1. A. Gessow and G. C. Myers, Jr., AERODYNAMICS OF THE HELICOPTER, Frederick Ungar

Publishing Co., NY, (Republished in 1967).

2. M.A. McVeigh and F .J. McHugh, RECENT ADVANCES IN ROTOR TECHNOLOGY AT BOEING

VERTOL, American Helicopter Society 38th Annual Forwn, Anaheim, CA, Paper No.

82-38, May 1982.

3. E.A. Fradenburgh, AERODYNAMIC FACTORS INFLUENCING OVERALL HOVER PERFORMANCE,

AGARD Specialists Meeting on the Aerodynamics of Rotary Wings, Marseille, France,

Paper No. 7, September 1972.

4. F.D. Harris,

Aerodynamics

No. 7-A2.

Paper Reprinted from AGARD Lecture Series No. 63 on Helicopter and Dynamics, 7 Rue Ancelle 92200 Neuilly Sur Seine, France, Paper 5. W. Johnson, HELICOPTER THEORY, Princeton University Press, Princeton, NJ, 1980.

6. R.W. Prouty, HELICOPTER PERFORMANCE, STABILITY, AND CONTROL, PWS Publishers,

Boston, MA, 1986.

7. F.J. McHugh, DESIGN OF THE 225-KNOT CONVENTIONAL ROTOR, Tenth European Rotorcraft

Forum, The Hague, The Netherlands, Paper No. 21, August 1984.

8. D.K. Young, F.J. Tarzanin, D.L. Kunz, USE OF BLADE SWEEP TO REDUCE 4/REV HUB