Analysis and application of compliant rotor technology

SIXTH EUROPEAN ROTORCRAFT AND POWERED LIFT AIRCRAFT FORUM

Paper No. 11

ANALYSIS AND APPLICATION OF COMPLIANT ROTOR TECHNOLOGY

J. G. Yen and W. H. Weller Bell Helicopter Textron Fort Worth, Texas U.S.A.

September 16-19, 1980 Bristol, England

ANALYSIS AND APPLICATION OF COMPLIANT

ROTOR TECHNOLOGY

by

J.

G. Yen and W. Ho Weller

Bell Helicopter Textron

Fort Worth, Texas U.S.A.

Abstract

Application of compliant rotor tech-niques to a four-bladed hingeless rotor indicates how this new technology can be used to control steady and one-per-rev

blade elastic twist. In tests of a blade

with positive camber airfoils, the steady and one-per-rev blade elastic twist increased with airspeed and gave rise to

large steady and oscillatory control loads

and stresses at blade midspan. Analysis indicated that a negative camber over 80-87-percent radius would have a beneficial effect on rotor loads without detrimental effects on performance, handling

qualiti-es, or cabin vibrations. These analytical

predictions were verified by flight test of blades with negative camber.

Correla-tion of the analysis with measured loads and performance is presented in this

paper. b c CG GW R

s

T Notations number of blades blade chord

helicopter center of gravity, stationline

gross weight

rotor radius

speed of ·sound rotor thrust

rotor thrust coefficient,

2TjbpcR(R0)2

maximum speed in level flight

with maximum continuous power

~ advance ratio

p air density

a' air density ratio

n

rotor speed

1. Introduction

The use of composite materials and

new manufacturing technology has allowed

rotor designers to incorporate modern airfoils, nonlinear twist, and planform variation in new rotor designs. Extensive research has recently been directed at devising a passive means of inducing

elastic twist of the blades to improve

rotor performance further and to reduce

oscillatory loads at high advance ratios. The analytical work described in Reference 1 examined the feasibility of improving helicopter performance and reducing flight loads by passive control of blade

tor-sional response. Design considerations such as reduced torsional stiffness, tip sweep and airfoil camber were studied.

Results suggested that tip sweep on a blade of reduced torsional stiffness

improved performance and reduced control

and blade loads and that negative camber reduced blade loads but generally degraded

performance. Wind tunnel tests and analy-sis of the low torsional stiffness,

four-bladed soft inplane hingeless model

described in Reference 2 demonstrated useful effects of tip sweep and negative

camber on blade loads. Similar results were obtained from wind-tunnel testing of

a four-bladed articulated rotor as

re-ported in Reference 3. Effects of blade tip geometry on rotor loads and

perform-ance were investigated using a

four-bladed articulated rotor model with

results documented in Reference 4. In another recent work (Reference 5), com-parison was made of the performance and

blade oscillatory loads for an articulated rotor system with four different tip geometries as predicted by analysis and as measured in a 1/5-scale model wind tunnel test, a full-scale model wind tunnel test, and flight test. Results suggested that blade tip sweep and tip planform taper

were effective in reducing rotor forward

flight power requirements and blade oscillatory loads. The objective of all

of these research efforts was to achieve

an optimal match between aerodynamic and structural designs such that the dynamic twisting response of a part or the full blade would be beneficial in terms of

This paper presents the results of applying compliant rotor technology to a full-scale, four-bladed, soft inplane rotor. The discussion focuses on how a state-of-the-art analysis was used to apply compliant rotor technology to the subject rotor and on corr-elation between theory and flight test data.

2 . Background

The subject rotor is that of the Bell Model 412 shown in Figure 1. The Model

Figure 1. Model 412 helicopter

8

-4

BUILT-IN TWIST, DEG

FLAT SIDED

SPECIAL PURPOSE AIRFOIL BHT 674

412 rotor is a four-bladed, soft inplane rotor incorporating advanced airfoils, nonlinear twist distribution, tapered planform and is of composite construction. Figure 2 shows the Model 412 blade plan-form, twist, and airfoil distributions. Natural frequency diagrams (fan plots) of the 412 rotor blade are shown in Figure 3.

Note that the first torsion frequency is located at 5.15/rev and that bending modes are well separated from exitation fre-quencies. The 412 blade airfoils (as designed) have a modest nose-down pitching moment below the critical Mach number.

Early development flight testing revealed two potential problems. First, nose-down steady pitch link loads (com-pression for the leading-edge pitch horn) were higher than anticipated, limiting the maximum up collective under boost-off operation. second, the beamwise and torsional oscillatory moments were higher than expected and indicated possible fatigue life limitations.

Analysis of the flight test data showed unexpectedly high steady and one-per-rev blade torsional moments. Further analysis of the measured moments indicated the underlying cause to be aerodynamic pitching moments.

In order to support the design of blade modifications to reduce the aero-dynamic pitching moments, a state-of-the-art flight simulation analysis, the C-81

BHT 674 AIRFOIL FX-080 MOD T/E ~TRANSITION----~~-BHT 674_{-~~---TRANSITION~}

I

AIRFOIL

_I

BASIC CHORD 15.9 INCH STATION 0.0 32.0 52.0 124.2 TRIM TAB 193.2 II 240.0 215 220.8 NEGATIVE CAMBER 276. 0 % RADIUS 0.0 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00 Figure 2. Blade planform, twist, and airfoil distribution

< 0 u > u _z

"

~ 0 ~

"

28 00 OPERATING RPM 24 00 ~" V;' ~I ~ ~~ o'<- <.o\ ~~

I

«~0

I

I

2000 _{' I}

y

/I

lST TORSION 1600

I

I I I I v V 1200 _{lST l"pvNE}

.V;..

/1

o< / <yl-~<f,. o~< /

,-.o

/

o'i • 800 I / 0-:::,<t '>l

/

/ > ~..,.

I

,./"?-;:_. ... / I / /

---I /

... I

t~G

400 I / I / J:V'-1?? :;;-p ... ... ::::--- -::::v?.-e.'IJ ,.,. _,..,. :::;;-.:::-=- I LEAD IJ\G I 0 0 100 200 300 400 ROTOR SPEED , RPM

Figure 3. Calculated frequencies of baseline 412 rotor

Rotorcraft Simulation Program described in Reference 6, was employed.

3. Analytical Approach

A review of the blade design support analyses revealed that the blade torsional moments and pitch link loads had been designed using an empirical method based on data from rotors having symmetrical airfoils (whereas the design beam and chord bending ·moments had been estimated using flight simulation program C81). Further review of the C8l predicted tor-sional moments and pitch link loads showed excellent agreement with the measured loads !

Analysis of the C81 predicted blade elastic twist revealed that the modest camber of the Model 412 blade airfoils was causing a substantial elastic twisting of the blade. A steady elastic twist of 2.0° and a llrev twist of 1 o was predicted at VH as shown in Figure 4. Flight test data showed even larger elastic twisting

(based on measured torsional moments). Subsequent C81 analysis suggested that the steady and 1/rev blade torsional

"'

"'

"

1 8 UJ ₀ H 3 8 -1 u H 8 _-2 UJ

::i

"'

-3

,..

"

.;: -4

"'

8 UJ ~ 1

"'"

z 0 0 0 0 FLIGHT TEST THEORY H ' 0~---U 8

---<L---ZUJ

g

-1 ~8 0 0 0.. ~ -2 .-<::J ~ -~----~----~~---7~--~~--~ _{0 20 40 60 80 100} PERCENT RADIUS

Figure 4. Spanwise distribution of baseline blade steady and lP advancing blade twist, tc

=

0.17, ~

=

0.282

moments could be reduced to an acceptable level by modifying the blade to have negative camber over an outboard portion of the blade. Figure 5 shows how the blade torsional moment varies as a func-tion of trailing edge tab angles in pro-gram C81. Figure 6 shows the effect of the change in airfoil pitching moment on elastic twist. Based on these predictions and the predicted reduction in blade loads, it was decided to modify an experi-mental set of blades to a negative camber configuration.

The blades were modified by bonding a 1.25-inch chord aluminum tab to the trail-ing edge of the blade from 80- to 87-percent radius (see Figure 2). The tab angle was set to 12 degrees (trailing edge up) to achieve the desired change in aerodynamic pitching moments.

Flight test results with the modified blades verified the benefits of the change in aerodynamic pitching moments predicted by C81. The effect of the tabs on loads, performance, vibration, and handling qualities are presented in the following sections.

STEADY, LBS

0~~~---100

-200

-300

-50

0 FLIGHT TEST --THEORY lP SINE, LBS

1001~

-20:r-t~===================:~~----~~===

lP COSINE, LBS

20

QL_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ ___

=r""·

₀

o

2

Figure 5. LBS I

10

4

6

8

TAB ANGLE, DEG

Variation of blade torsional moment at 48% radius with tab angle, tc = 0.17,

~ = 0.282

4. Effect of Negative camber on Rotor Loads

4.1 Pitch Link Loads

~

12

Variation of steady and oscillatory pitch link loads with airspeed for the baseline and the tabbed blades is depicted

in Figure 7. The reduction in the steady component had been predicted by the analysis and was verified by flight test. Both the theory and measured data indi-cated that the reduction in the oscil-latory pitch link loads due to the negative camber was quite small. The reason for this was that the benefit of the tab on the one-per-rev reduction was partially offset by an increase in the three-per-rev component. A harmonic decomposition of pitch link loads of the baseline blade is compared with that of the tabbed blade in Figure 8.

Measured and calculated pitch link load waveforms for the baseline blade and the tabbed blade at the same flight con-dition are shown in Figures 9 and 10, respectively. Correlation between theory

C) 1

"'

a

'"'

0 Ul H

"'

'"'

-1 u H

'"'

Ul _-2 j

"'

>< -3 a .:

"'

'"'

"'

-4

1

<.!)

"'

"'"

z

0

H

.

UE-'

zw

-<:H >:;<

-1

"'"'

.:.,

"'"

""'j

-2

"'

-3

---

---

---0 0 0

BASELINE 0 FLIGHT TEST

BLADE - - - THEORY TABBED BLADE ---THEORY NOSE UP

---

_---

--0---

_{0 0}

0.2

0.25

ADVANCE RATIO

0.3

Figure 6. Variation in elastic twist at

75% radius with advance ratio, tc = 0.17

and the measured data is good for the baseline blade. The measured increase in three-per-rev component with the tabbed blade was not predicted by the analysis even with the inclusion of a free wake in the analysis.

4.2 Torsional Moments

The outboard negative camber was very effective in reducing the midspan tor-sional moment. A nearly fifty percent reduction in the steady and one-per-rev components was realized as illustrated in Figure 11. Magnitudes of the higher harmonics are small in comparisons with the one per rev. A more than forty percent reduction in the oscillatory torsional moment was directly attributed

to the reduction in the first harmonic

moment.

4.3 Beamwise and Chordwise Moments

Tab effectiveness on blade midspan beamwise and chordwise moments is shown in Figures

12

and

13,

respectively.

A

twelve percent reduction in bearnwise oscillatory loads was measured at design gross weight

ADVANCE RATIO

O'r-~·;1~5---~·~2r0---~·~2r5---~·~30 BASELINE BLADE 0

FLIGHT THEORY ~ -200 >-< >< 0 ..:

"'

E-< Ul Ul

"'

>-< 0 (!j >-<

"'

z H

"'

_:r: u E-< H

'"

Ul

"'

-4 00 -600 500 400 300 200 100 0 0 -100 "' -200 0 (§ -300

"'

"'-400 z H "' -500 :r: u ::: -600

'"

-700 z 0 H Ul Ul

"'

"'

"'

>: 0 2. z -:

"'

>: 40 TABBED

"'

FLIGHT 400 BLADE ----THEORY Ul

"'--

---

"'

>-<

-

6.-- -.._ _><

"'

"'

0 E-< ₂₀₀ ..: 0 >-< >-<

~

H u Ul 0 0 .15 .20 .25 ADVANCE RATIO Figure 7. Variation of steady and oscillatory pitch link loads

with advance ratio, tc = 0.17

BASELINE BLADE ₅₀₀ TABBED BLADE

400 300 _z 0 H ~ 0 E-< j >-< H u

I

I Ul 0 _I 2P 3P 4P 1P Ul ~ ZUl <1!0> ₀

"'"'

;;:

>:o. >:

"'

0

"'

u H u Ul 0 200 100 5P 0 1P 2P 3P 4P

Figure 8. Harmonic decomposition loads, tc = 0.17, ~

of measured pitch link 0.282.

AZIMUTH, DEG

120 200 280 360 THEORY; UNSTEADY + FREEWARE THEORY, UNSTEADY FLIGHT 17C/1372, MEASURED AZIMUTH, DEG 0 120 200 280 0 (§ -300

\

"'

~ -400 H

"'

-500 :r:

~

-600

'"

_-700 TEST TEST . 30 5P 360 ./

Figure 9. Pitch link loads waveform, baseline blade, t = 0.17, ~ = 0.282. c

Figure 10. Pitch link loads waveform, tabbed blade, tc = 0.17,

2400

"'

';' 2000

"

H .1600

,..

"

"'

1§

1200 >: >-< ~ 800 0 H Ul

"'

0

,..

400 0 5000

"'

>-< I

z

4000 H ~ 3000

"'

1§

>: 2000 (!) z H ~ 1000

"'

"'

0

-

z 0 H Ul Ul

"'

"'

"'

>: 0 u z ..:

"'

>: ><

"'

0

,..

..: >-< >-< H u Ul 0

BASELINE BLADE 'rABBED BLADE

2400 1---2000 1600 G! 0

,..

..: >-< >-< H u Ul 0

I

I

I

1P 2P 3P 4P 5P 1200

-

>< z

"'

0 0 H

,..

z

Ul _j Ul ..:

"'

>-<

"'

_"'

H >:

"'

u >: Ul 0 0

n

I

u 800 400 0 2P 4P 1P 3P

Figure 11. Harmonic decomposition of measured torsional moment

a.t 48% radius, tc = 0.17, Jl = 0.282.

BASELINE BLADE ₅₀₀₀ TABBED BLADE

4000 3000 G! 0

,..

..: 2000 >-< H H u Ul 1 000 0

h

I 1P 2P 3P 4P 5P 0 1P 2P 3P 4P SP

Figure 12. Harmonic decomposition of measured beamwise

12000

"'

10000 >-< I z H 8000

,..

z

"'

6000 >: 0 >: (!) 4000 z H Cl z

"'

2000

"'

0

bending moments at 48% radius, tc = 0.17, Jl = 0.282.

BASELINE BLADE TABBED BLADE

12 000 10 000 G! 8 0 I

-,..

j 6 >-< H u Ul 4 0 000 G! 0

,..

j >-< H u Ul 000 000 0 2 000

I

1P 2P 3P 4P 5P 0 1P 2P 3P 4P 5P

Figure 13. Harmonic decomposition of measured chordwise bending moments at 48% radius, tc = -.17, Jl = 0.282.

J l

5P

and v

8. The chordwise oscillatory bending

moment was reduced by twenty percent. Load reduction was realized in most of the harmonics. The lower loads are the result of decreased elastic twist as discussed earlier.

The effect of the tab on beamwise and chordwise oscillatory loads in the hub and blade is shown in Figure 14. Also shown are the predicted loads for the baseline blade and the tabbed blade. Both the measured and analytical data demonstrate some reduction in the rotor beam and chord loads using the negative camber with the most beneficial reduction in the oscil-latory yoke beam component.

BASELINE BLADE _{0 FLT TEST}

--THEORY 20

TABBED SLADE _{61 FLT TEST}

---THEORY o+---~~--~--~~ _{0 20 40 60 so 100} JO

"

'"'-o,+---~---_:~--~~ 0 20 60 8 0 100 PERCENT RADIUS

Figure 14. Spanwise distribution of blade loads,

tc

=

0.17, ~

=

0.282

5. Effect of Negative Camber on Vibr~~ion

Measured blade yoke beam three-per-rev and five-per-three-per-rev amplitudes from the tabbed blades are comparable with those from the baseline blades. However, some differences in phase were noticed. As a result, the characteristics of hub pitch-ing and rollpitch-ing moments are different between the untabbed and the tabbed blades. Measured four-per-rev vertical vibrations in the Model 412 cabin as

influenced by the tabbed blades are given in Figure 15. With the baseline blades, the pilot seat vibration was quite low, but the copilot seat vibration was high at VH" The negative camber lowered the copilot seat vibration to a comfortable level but increased the pilot seat vibra-tion. However, the pilot seat vibration with the tabbed blades did not exceed the design goal. - - - - BASELINE BLADE - - - TABBED BLADE .lS _{I PILOT SEAT!} DESIGN GOAL

~

. OS 0 4/REV VERTICAL .2S VIBRATION, .20 !COPILOT SEAT! G .1S .10 .OS

~

--

...

___

....

0 60 80 100 120

TRUE AIRSPEED, KNOTS Figure 15. Effect of negative camber

on cabin vibrations. design gross weight, aft cg

6. Effect of Negative Camber on Per-formance

While simplified theory might suggest that the nose-up tabs would degrade hover performance while improving forward flight performance (the addition of nose-up pitching moment effectively reduced the geometric blade twist), analysis indicated only a small effect. This was confirmed by flight test data. The effect of blade tab on hover performance is presented in Figure 16; while that on forward flight performance is presented in Figure 17. Within the tolerance of data scatter, the tabs' effect on perfo.rmance is, in general in agreement with the C81 analysis.

7. Effect of Negative Camber on Handling Qual1t1es

Both C81 and flight test data indi-cated that less collective input was required with the tabbed blades. This was due to the fact that the tabs reduced the

nose-down steady elastic twist by as much as 0. 5 degree at the blade root. for the

design gross weight and VH (Figure 18).

1600 l~l)(} 1200 tOOO

'

0 'Bo 0 0 0 0 0~ 0 0 0

fl

0 0

Figure 16. Effect of negative camber

on hover performance 1800 0 ~1600

<:l

El4oo

C/l z

"'

-E:l200

"'

~ ~ 1000

"'

C/l

"'

§l

8 00

"'

0 E-< ~ 600 z H

ill

400 200 0 48A -GW/cr'

s

QR CG 0 55A -GW/cr' s QR CG 0 0 BASELINE BLADES 12000 LB 1117 FPS 756 FPS 138.8 IN. TABBED BLADES 12000 LB 1117 FPS 756 FPS 138.8 IN. 0 0 0 0 0 oo 0 oO 0 0 0 <:0 0 0 20 40 60 80 100 120 140 TRUE AIRSPEED, KNOTS

Figure 17. Effect of negative camber on main rotor forward flight

performance l 0 -1 t>

---"

0

-CJ

"'

"'"

z H ' UE-< ZC/l 0:H >S:

"'"

.:.,

"'"

...;<: -2 -3 -4 1 -2 0 BASELINE BLADE TABBED BLADE

"

_"

_"

0 0 0 0 FLIGHT TEST THEORY

"

FLIGHT TEST ---THEORY 0 0 H

"'

-3L---~~--~----~----~--~~ 0 20 40 60 80 lOU PERCENT RADIUS

Figure 18. Effect of negative camber on steady and 1P advancing blade twist, t = 0.17, ~ = 0. 282 c

Figure 19 shows the variation in the

longitudinal cyclic stick position with airspeed for the baseline and the tabbed blades. Data indicate that the tabbed

blades require about 8 percent more for-ward stick position, which is equivalent to nearly 2 degrees. As shown in Figure 18 a reduction of 2 degrees one-per-rev advancing blade twist was measured at the

0 BASELINE GW=7630 0 TABBED 100 80 F/A CYCLIC 60 STICK POSITION, % 40 20 0 GW=8374 20 40 lb, 1b, 60 CG=144.2 in. 80 100 120 CALIBRATED AIRSPEED, KNOTS Figure 19. Effect of negative camber

on longitudinal cyclic

outboard end of the blade using the tab. In order to achieve the same trim, the tabbed blades require more longitundinal

cyclic stick input. It was also observed

from the data in Figure 19 that the stick

gradient is positive and is increased as a

result of the tab.

Other flight test data (not shown) indicated that the blade nose-up pitching

moment did not affect the dynamic

stabi-lity of the flight modes.

B. Conclusions 1. 2. 3. 4. 5.

Mod1fying the Model 412 blades to incorporate outboard negative camber effectively reduced steady and one-per-rev elastic twisting of the blades. As a result, the steady

control load was reduced by 40

per-cent and substantial reduction in blade torsional moments was realized. The negative camber had a beneficial

effect on Model 412 cabin vibrations.

Considering the measured performance data scatter, the blade nose-up

pitching moment did not significantly affect hover or forward flight

per-formance.

Less collective and more forward

cyclic were required with the tabbed

blades in order to achieve the same trim condition.

The state-of-the-art flight

simu-lation analysis C81 accurately

predicted the effect of negative

camber on rotor loads, performance,

and handling qualities. As a result 1. 2. 3. 4. 5. 6.

of the analysis and flight test verification during the Model 412 development program, trailing edge

reflex with increased spanwise length and less camber was incorporated into

the Model 412 production blades. These reflexed trailing edge blades

were designed to provide the same

benefits as the experimental tabbed blades.

References

Blackwell, R. H., 11_{Investigation of}

the Compliant Rotor Concept, 11

USAAMRDL-TR-77-7, June 1977.

Doman, G. S., Tarzanin, F. J., Jr.,

and Shaw, J., Jr. I 11lnvestigation of

Aeroelastically Adaptive Rotors111

USAAMRDL-TR-77-3, May 1977.

Blackwell r R. H. I et al, 11Wind

Tunnel Evaluation of Aeroelastically

Conformable Rotors, 11 _American

Heli-copter Society 36th Annual Forum, May 1980.

Weller I

w.

H. r 11Experimental

In-vestigation of Effects of Blade Tip

Geometry on Loads and Performance for

an Articulated Rotor System," NASA TR 1303, January 1979.

Jepson, D., et al, 11_{Analysis and}

Correlation of Test Data from an

Advanced Technology Rotor System, 11

NASA CR-152366, July 1980.

McLarty, T. T., "Rotorcraft Flight Simulation with Coupled Rotor

Aero-elastic Stability Analysis r II