· " 

Nils/} 77YJ - 77;370 

NASA TECHNICAL MEMORANDUM NASA TM-77370 

NASA-TM-77370 19840007051 

HELICOPTER BLADE TIPS 

R. Lyothier 

Translation of "Les Extremites de Pal~s d'Helicopteres", 

Association Aeronautics et Astronautique de France, 

Colloque d'Aerodynamique Appliquee, 18th, Poitiers, 

France, Nov. 18-20, 1981, Paper NT 81-19, 23 pp. 

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NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, D. C. 20546 OCTOBER 1983 

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NF00316 
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Translation of "Les Extremites de Pales d'Helicopteres", 

Association A~ronautics et Astronautique de France, 

Colloque d'Aerodynamique Appliquee ,. IB.th, Poi ti e:t'. 5 , France, 

Nov. 18-20, 1981, Paper NT 81-19, 23 pp. 

16. Ahrn.ct· 

Methods of improving heTicopter performance and vibration 

level by proper shaping of helicopter blade tips are considered. The principle ·involved consists of redllcing the 

extent of the supersonic zone above the advancing tip and 

of the tfirbulent ~nteraction. For stationary and advancing 

flight, the influen·ce of the rotor and the problems posed • 

by blade tips are reviewed, The theoretical methods of 

dealing with the two types of flight are briefly stated, and 

the experimental apparatus is described, including model 

triple and quadruple rotors. Different blade tip shapes 'are 

shown and briefly discussed. The theoretical ·results includ~ 

an advancing speed of 309 kni/h_and a blade tip rotational 

speed of 215 m/s. The ·exper·imental values are advancing 

speed of 302 km/h and blade tip Mach number 0.86 for both 

types of rotor. 

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SUMMARY 

HELICOPTER BLADE TIPS * 

R.· Lyothier 

SNIAS, National Industrial Space Society 

Marignane, France 

A substantial gain in the vibrational level performance of a 

helicopter can be brought about by using blade tips of particular 

shape. 

**/1 

The principle consists of reducing the extent of the supersonic area which appears over the advancing blade as well as reducing the effects of vortex interaction. After first selecting the 

shapes by calculation, performance and force measurements were performed in a wind tunnel on a rotor model. 

1. INTRODUCTION 

Research has been performed over several years in order to 

improve the performance and comfort of helicopters. Two important 

• results have been achieved (Puma 330J and Dauphin 365C): 

- increase in the liftable mass for stationary flight 

- increase in the maximum speed. 

These gains were obtained by using new blade profiles (OA 

family, reSUlting from a SNIAS-ONERA collaboration. 

New phenomena appeared which were related to the increase in 

velocity, both for flight and altitude and for the payload factor. 

* 18th Applied Aerodynamics Col1oqium, November 18-20, 1981, Poitiers, 

France. ISSN 0243-0177, French Aeronautical and Astronautical Assn. 

** Numbers in margin indicate pagination of foreign text. 

1 

· .... ,.1 0" 
c __ ...--'- .. ____ ~_-_':o!..,.. .... ol~ 

Depending on the flight conditions, this affects the power consumed 

by the rotor, the loads and the vibration level. This is the result 

of two main causes: 

- the appearance of a super-critical regime at the extremity 

of the advancing blade, whose Mach number can reach 0.9. 

This results in a torsion excitation and an increase in 

drag. 

- the vortex interaction which comes from an encounter of the 

blade extremity region with the wake of the preceding blade, 

which can lead to flow separation. 

An additional refinement of the extremity profiles would be 

desirable, but is not feasible technologically (as an example, the 

relative thickness can already reach 6% at the extremities of the 

blades of Dauphin N). In addition, this does not solve the problem 

of vortex interaction. Therefore, it seems necessary to find a 

solution in the modification of the plan form of the blade extremities. 

2. GENERAL REMARKS ABOUT THE ROTOR 

Problems associated with the blade. extremities appear in two 

distinct operational ranges of the rotor (references [1] and [3]): 

- stationary flight, for which the aerodynamic conditions 

remain stationary and the Mach numbers are moderate (on the 

order of 0.6). 

- forward flight for which the flow is unsteady and can reach 

the critical speed. 

2.1 Stationary flight 

The loads on the blades are independent of their azimuth, 

except for a small cyclical command which provides for equilibrium 

of the craft. The rotor is symbolized by a lifting disk (Figure 1) 

and the loads have the distribution indicated in Figure 2. 

2 

. . .. . 
· .• ,-... -.-.. -~ 

--, '-- ,-, --~--------,-----'-~---"----~----

vortex 

2 V, 

Fig,ure 1. Stationary flight 

I 

I 

I 

0,7 R 

R R 

Figure 2. Stationary flight loads 

3 
4 

J 

mast 

Figure 3. Inclination of the rotor disc during forward 

flight. 

\iii, ~-----\.'-- ----p-oy-oFigure 4. Dynamic equilibrium of the blades. 

FA aerodynamic force 

Fe centrifugal force 

F. inertial force ~ 
The lift is a maximum in the extremity area and the resultant FN 

is applied at about 0.7R. This induces a vortex sheet and an edge 

vortex. which interferes with 'the other blades. t 

2.2 Forward flight 

The rotor provides a resulting force whose components FZ and 

FX respectively provide the lift and the propulsion of the craft 

(Figure 3). The component FX is created by inclining the rotor disk 

in the desired direction. This is the result of a cyclical variation of the step and the composition of the translation and velocity 

compositions. The blades have a flapping degree of freedom and this 

leads to a mechanical system which responds to this excitation at 

the rotation frequency. This is equivalent geometrically to a 

balancing of the disk. 

The forces involved (Figure 4) are the following: 

- aerodynamic forces, modulated by the cyclical nature and 

composition of the speeds, 

- the inertia forces, 

- the centrifugal force. 

Figure 5 shows the variation of the -load distributions as a 

function of blade azimu~h. 

The response of a blade to cyclit::al excitation occurs \d th a 

phase delay of 90° (close to resonance). In order to obtain the 

inclination of the rotor disk in the forward direction required for 

propulsion, the incidence has to be a minimum over the advancing 

blade and a maximum over the regressing blade. The aerodynamic 

excitation resulting from the composition of the speeds acts in a 

similar manner. 

3. PROBLEMS CAUSED BY THE EXTREMITIES 

5 
a~v~n~ing position G 

0 

~ ~= 900 

ROTOR 

pos 1. t1.on " advaz:c~ng 0 

v= 1eoo 0 

receding eJl d " a e paSJ_ t1.on 

v= 2700 

,. ~ - ---..-.~ 

b.lade '. speed 

o R 

r 

o R 

~, lin1Il, ORO R 

~,-, o . R 0 R 

o R o R 

Figure 5. Behavior of the rotor discs in translation 

flight. 

The problems caused by the blade extremities are essentially 

the result of the following: 

6 

- supercritical phenomena (the appearance of shockwaves) on 

the advancing blade, 

- from the vortex interaction which by increasing the incidence of the extremity can lead to flow separation when the 

Mach number is sufficiently large. 

3.1 Stationary flight 

The problems encountered over the extremities are caused 
.... __ 25% . 

, ---

Figure 6 

essentially by the attached vortex. It starts at about 25% of 

chord and passes above it along the top side (Figure 6) and therefore creates a depression or an additional lift (reference [4]). 

However, the passage of the vortex below the following blade 

increases the incidence angles over the extremity and decreases 

them over the interior part of the blade (references [5] and [6] )" 

This is translated into a reduction in lift of the blade for a 

given step and an increase in the drag with a given lift. 

Finally, since the extremity is subjected to large dynamic 

pressures, the last 10% of the blade provide 25-30% of the lift 

(ref erence [4]). 

This is also the zone where the Cz values are maximum. This 

leads to two important consequences: 

- the extremity has the tendency to separate before the rest 

of the blade, 

- the incident Mach number is on the order of 0.6, and shockwaves appear because of the increased incidence angle. 

7 
A 

sect,ion A-A 

~:=~ V~ 

interaction 

vortex 

. ~-V~ 

advancing speed 

of the apparatus 

Figure 7 

8 
.,~. __ ....... ·~ _____ T ____ ._· _~ __ :..._.: ______ _ 

3.2 ~orward flight 

The functioning of the extremities has been studied by many 

authors (references [7], [8], [9], [10]). 

These studies have allowed one to distinguish various problems 

related to the blade extremities. 

compressibility (shocks) over the advancing blade, 

- vortex interaction of the rear blade. 

The vortex interaction which is produced on the rear blade of 

a four-blade rotor under certain conditions of flight can lead to 

flow separation near the extremity (reference [7]) which is the 

case for steady flight. 

Transsonic phenomena which appear on the advancing blade 

extremity, whose Mach number can reach 0.9, are aggrevated by an 

increase in the incidence due to the vortex interaction, as shown 

in Figure 7 (see reference [10]). 

The appearance of shockwaves is translated into an increase in 

drag and noise and induces large variations in the torsion moments. 

4. STUDY MEANS 

These are both experimental and theoretical. 

4.1 Theoretical means 

Complete methods for studying the case of forward flight or 

stationary flight are not available, in spite of SUbstantial efforts. 

Only partial aspects of problems can be studied. 

4.1.1 Stationary flight 

9 
advance speed 

1/1 = 1800 

1/1 = 900 

) . I 

i' " 

1/1=0 flow zone which is 

treated at a time t 

Figure 8 

The load distribution over a rotor blade for stationary flight 

is obtained by a vortex method (work of M. Hirsch reference [11]). 

The blade is represented by a lifting line with a circulation 

distribution similar to that of Prandtl for a finite span wing. 

Nonlinear effects and compressibility effects are induced by 

corrections made from the polars of the profiles used. 

10 
The balancing of the wake is done using an iteration process 

whose convergence is acquired. when the velocity induced is tangential with the vortex wake. 

4.1.2 Forward flight 

The flow over the extremity of a helicopter blade in forward 

flight is three-dimensional, unsteady and supercritical, a problem 

which has not been solved up to the present: 

The available calculation methods only take into account partial aspects of the problem. This includes three aspects. 

- Three dimensional calculations for a supercritical type wing. 

The flow is steady and the reference speed is constant. This 

is a large perturbation method which predicts shockwaves. This 

allows one to estimate the geometric shape (thickness law and 

extremi ty) • 

- Finite elements 

At the present time, it is applied for the same conditions 

as above. The representation of the shape of the extremity is more 

refined. It is possible to introduce a reference speed which varies 

linearly with span in order to closely approach the conditions of 

the helicopter blade. 

- Three dimensional unsteady calculation with no lift for 

a helicopter type situation. 

This method considers three dimensional effects and introduces 

unsteady effects (Figure 8). It uses the hypothesis of small perturbations. In spite of its limitation to the non-lifting case, 

it is useful to research shapes and this eliminates supersonic 

areas caused by the sweep back effect. 

4.2 Experimental means 

11 

_.J 
emitting blade 

plane 

Figure 9. Wing-vortex interaction 

4.2.1 Wing-vortex interaction 

---;-----.-----

Above we discussed the importance of the influence of the 

extremity vortices on a blade. 

In order to better understand these phenomena, tests were performed by ONERA in the S3 wind tunnel at Chalais Meudon (reference 

[15]). The principle is summarized in Figure 9. 

We wish to observe the influence on pressure distribution over 

a blade, of bhe variation intensity and position of an edge vortex 

coming from a preceding blade, using two half wings at the wall. 

The vortex intensity value is modified by varying the lift of the 

12 
0, 

", "-.' ,.-~ .• 'f - --.----.-. .-------~. -_-______ -•• 

twist axis ---- - -------

o 

I 

R ·1 

Figure 10. Rectangular extrimity 

--. ..... _---

u 

I 

R 

.. I 

Figure 11. Parabolic extremity with sweep back 

13 
emitting blade. The blade is equipped with pressure transducers. 

4.2.2 Rotors-model, 

These are test devices developed by the S2 of the ONERA at 

Cha1ais-Meudon. A fixed component balance allows one to measure 

performance. The velocity range extends from steady flight up to 

a forward speed parameter of 0.5. The geometrical parameters are 

the following: 

- a collective step, inclination of the rotor axis. 

After being used to study the influence of profiles and the 

laws of twist, this installation is used in order to analyze the 

operation of blade extremities. Two types of rotors are tested. 

- Three blade ONERA rotor 

The purpose is to perform fundamental studies on the extremities. They can be detached at 0.8 R. The chord dimension allows 

one to install pressure transducers. The rotor diameter is on the 

order of 1.700 m and the chord is 0.123 m. For the test performed, 

only the influence of the plan form is examined, the twist and the 

profiles are retained. The shapes studied are the following: 

rectangular extremity, parabolic extremity with sweep back (Figures 

10 and 11). The latter has the purpose of producing shockwaves 

which appear on the advancing blade, using the sweep back effect. 

- Four-blade SNIAS rotor 

The purpose is to study performance. The hub used is articulated for flapping and drag, and has a damper. The blades are 

made of composite material (reference l18]). The dimensions of the 

chord (0.05 m) do not allow the installation of the pressure 

transducers, but load measurements are performed on the rotor. The 

parameters studied are the following: 

14 

(' 
- - -.--~-~-'------.-~.~' '.- _. ------~~ 

- ---- .- -- - -- -~ -- _. --

ROTOR T1VIST PLAN FOR~ PROFILE 

NUMBER t: .-

1 - 8° I BV 23010 -1,58 

2 - 14° ) 8V 23010 - 1,58 

:} - 8°.3 - - --

~ . linear and OA209 

nonlinear 

4 - 8°.3 ~ ......... OA209 

5 - 8°.3 S ...... OA209 .......-

I 

6 - 8°.3 ? I" OA209 

0.95 RT I 

- J R 

- the fineness, the twist and the plan form of the extremity. 

The configurations which were tested are given in the table 

above. 

5. CLASSIFICATION OF EXTREMITY SHAPES 

The first tests performed showed the influence of the following 

parameters: 

- relative profile thickness 

-twist 

- plan form. 

Numerous studies were performed on ingots with various plan 

forms. 

15 
arrow 

sweep back of 

leading edge' 

I simple shapes 

file 

trailing edge 

sweep 

Figure 12. Ingot shaped examples 

- Simple shapes (Figure 12) 

parabolic 

OGEE·TIP 

There are two types, those where the leading edge is modified 

with respect to a straight extremity (sweep back, file, parabola, 

sweep back of the leading edge) and those which have a modification 

to the trailing edge (sweep back of the trailing edge, ogee tip) • 

. The former tend to modify the flow by reducing the over-velocity 

caused by a sweep back effect. The second ones are of interest 

because of vortex interaction. This allows one to calibrate the 

extremity vortex and to accelerate dilution. 

16 
file sweep back 

(type S76) 

M.B.B. ingot 

«SUB·WING» 

composite shapes 

trunc~!:e<!_ 

~ ~weep 

. back 'a t _ the- traiii; 

sweep back ~~ __ . 

parabola 

vlestland ingot 

evolution sweep -back 

sweep back + 

truncated- sweep 

«SPLIT-TIP» SPOILER 

Figure 13. Ingot shaped examples 

- Composite shapes (Figure 13) • 

They tried to combine the advantages of various simple shapes. 

This is a case for example of a swept back extremity--file shape 

which adds the effect of reducing the over-velocity caused by 

sweep back, with an effect of lowering the vortex intensity at the 

extremity, caused by a reduction in the chord. 

In addition to these complex shapes, we have those which have 

external elements (SUBWING, SPLIT-TIP). These tend to accelerate the 

dilution of the extremity vortex by a velocity composition effect. 

17 

".o .......... 
6. RESULTS 

6.1 Theoretical results' 

Only the unsteady three-dimensional program at the present 

time gave results which cou~d be exploited. We give an example 

in Figure 14. This is a comparison of the isomach line distribution 

over two extremities with different shapes. The calculation conditions are the following: 

.~. '. "''t:. ' . .L~ 

- quasi-stationary (the terms which depend on time are made /14 

zero) 

- azimuth of 90° 

- forward speed 309 Km/h 

- rotation rate at the blade extremity is 215 m/s. 

Over the rectangular shape the supersonic zone is bordered by 

very close lines which allows one to conclude that there is a shockwave there. 

This phenomenon disappears when the leading edge is swept back. 

6.2 Experimental results 

6.2.1 Three-blade rotor (ONERA) 

The most interesting results were obtained using a swept back 

extremity whose leading edge is partially in the shape of a parabola 

(see Figure 11). 

The results shown in Figure 15 were obtained for the following 

conditions: 

- forward speed 302 Km/h 

- Mach number at the blade extremity: 0.86. 

18 
Figure 14. Velocity map 

p rectangular extremity 

2.0 

,~/ 

.~ 

1.5 

~ 

~ ~~ " ~ ~. back extremity 

Cz rn 

1.0 I 

.3 .4 .5 

Figure 15. Three blade rotor 

D 

19 
FZ 

· - --,--- ~ ..... -..... ------~-.-... , .. ~~--~~.-.-.- --- ._ .. - - ~.- .... ~ .. 

(influence of twist I 

lifting performance FM (S2 Chalais-Heudon tests) 

FM = Fz 312 

'liPS p 

twist 

nonlinear 

::::J twist 

1..!... 0,65 

0,5 R 

I twist law 

O---~ __________ ~_ 

,-

0,6 

f .50 

0,55 Cz m = 0,03 Z 

·loo 0,3 0,4 0.5 0,6 0,7 nonlinear t , , I .. twist 10 15 20 - 100 FZ z=---

112 pS a U2 

Figure 16. Stationary flight performances 

The gain over a classical extremity (rectangular) is on the 

order of 7% for the reduced power, for an average lift of 0.48. 

By considering the case of the Dauphin 365 N, the possible increase 

in speed would be on the order of 8 Km/h and the increase in takeoff we~ght would be 60 Kg (references [19] and [20]). 

6.2.2 Four-blade SNIAS rotor 

In order to show the influence on performance of the twist and 

the plan fprm for stationary flight, Figures 16 and 17 show the 

variation of the figure of merit (ratio of the power supplied by the 

rotor to the FROUDE power derived from the theory of the induced 

velocity) as a function of average Cz. The nonlinear twist (Figure 

16), by reducing the influence of the vortex interaction, allows 

one to increase the figure of merit. Also, the four plan forms are 

shown in Figure 17 but only the parabolic extremity gives a substantial gain. 

20 
S2 Cha1ais-Meudon wind tunnel 

4-b1ade SNIAS rotor a =.E.!.. = 0,085 0 =. 1 ,474 m Ow =. - 8,30 

1l'R 

FM 

0,65 

" --- • 

, 

t:? ··'·r~ • 

0,60 , /~ \::::J t?' 

e9 t . :::::J right extremity 

0,55 (REFERENCE) 

0,3 0,4 0,5 0,6 0,7 Czm 

I ! I , I I ~ 

10 15 20 Z 

FM 3-b1ade ONERA rotor a = 0,"37 o = 1,714m 8vr = -12 

0,70 ~ es~~"" 

~ ~ 

0,65 ~ right extremity 

(REFERENCE) 

0,60 

0,3 0,4 0,5 0,6 0,7 ezm 

I I 

I I i , ~ 

10 15 ~O Z 

Figure 17. Influence of the shape of the extremity of the blade 

on the lifting performance for stationary flight. 

21 
.. 

., h :: 0,4 U = 210 m/s 

(V 0::.302 km/h) 8 2 Chalais-Meudon wind tunnel 

CXS - 0,1 . 

So 4 blade 8NIA8 rotor . 

/ 

. 0=0,085 

o :: 1,414 m 

____ e vr :: - 8,30 

P 100 P 

-1/2 p Sa U3 

3~blade ONERA rotor 

0=0,137 

0= 1,734 m 

2 file 

O - - 120 

vr -

1,5 swept 

t 

0,3 0,4 0,5 Cz m 0,032 

-;,._ --------~,--~'~ __ --_II------~r--~~ 

10 12,5 15 17;5 Z _ 100 FZ 

- 1/2 p S a U2 

Figure 18. Cruise performance 

22 
· . 

... 

For for~ard flight, the tests were carried out for a reduced /17 

drag of the craft of 

\ CxS Sla = 0.125 

(S master couple,a ratio of blade area to area Sl of the rotor disk) • 

The forward speed was 302 Km/h and the Mach number at the blade 

extremity was 0.86. Figure 18 shows the various results obtained 

for the four extremities. We can see that for a large load, the 

straight extremity remains the best. However, for Czm less than 

0.4, the parabolic extremity gives a slight gain. 

These latter conclusions only apply to the four-blade rotor. 

Tests on the three-blade rotor show an advantage of the parabolic 

extremity with sweep back for the entire range considered. The 

same holds true for stationary flight for the two rotors. Therefore, 

it is this extremity which gives the best compromise. It is used 

for confirmation in f·light using the Dauphin N helicopter. 

7. CONCLUSION 

After making substantial improvements to helicopter rotors by 

defining specific profiles, especially the OA family, the adaptation 

of extremity shapes can be considered in order to exploit their characteristics in the best way. 

The sensitive parameters are the following: 

- the plan form, in order to limit the supersonic zone and to 

attenuate shockwaves, 

- the twist law and the tapering law, in order to reduce the 

vortex. interaction effects. 

The gains obtained primarily involve the increase in performance ; 

(takeoff weight and maximum speed) and a reduction in the vibration 

level. They were demonstrated by wind tunnel tests of rotor models 

23 
having small dimensions, as well as by flight tests. 

~ 

. . ~ 

The theoretical means for studying the propleni are still being 

developed and at the present time only allow treatment of partial 

problems. The importance of knowing the aerodynamic field and 

criteria of selection, however, can be used for the examples mentioned above. 

REFERENCES 

1. Giravions. F Legrand. 

2. Technological helicopter evolution perspectives. 

G. BEZIAC 

1st Inter Society Atlantic Aeronautical Conferences 

26-28 March, 1979, Williamsburg, USA. 

3. Helicopters. 

P. LEFORT and J. HAMMAN. 

4. Wind Tunnel Simulation of Full Scale vortices 

JAMES B. RORKE and ROBERT C. MOFFITT 

NASA C.R. 2180 (March 1973) 

5. Aerodynamic Problems of Helicopter Blade Tip 

B. MONNERIE, J. J. PHILIPPE 

3rd European Rotorcraft and Powered Lift Aircraft Forum 

(Sept. 1977) 

. 6. Influence of blade extremities on the performance of a rotor 

Bibliographic synthesis 

J. M. POURADIER 

Document SNIAS 

.7. Flights experimental on aerodynamics features affecting helicopter blade design 

P. BROTHERWOOD, M. J. RILEY 

3rd European Rotorcraft and Powered Aircraft Forum 

(Sept. 1977) 

8. Detailed aerodynamic measurements on a model rotor in the blade 

stall regime 

McCROSKEY and R. K. FISHER, JR. 

9. Behavior of a rotor outside of the axial flight range in the 

large Modane wind tunnel 

M. LECARME 

AGARD CP 111 

24 
·1 . 

10. ONERA 

Flow over the blade extremities of a non-lifting helicopter 

rotor in the tran~onic range 

J. C. BOHL, P. LAFON, J. J. PHILIPPE 

RTS 64/1369 AY432 (August 5, 1976) 

11. Aerodynamics of rotors in stationary flight, balancing of 

the wake 

POURADIER 

contract DRET No. 75/276, lot No. 2 

12. Extension of the finite element method for two-dimensional 

transonic conditions 

Y. MORCHOISNE 

12th Applied Aerodynamic Collogium 

13. Numerical calculations of unsteady transonic potential flow 

over helicopter rotor blades 

FRANK X. CARADONNA, MORRIS P. ISOM 

AIAA JOURNAL, vol. 14, no. 4, 1976. 

14. Calculation of three-dimensional unsteady transonic flow past 

helicopter blades 

J. J. CHATTOT 

NASA TP 1721 

15. Characteristics of a helicopter blade - influence of the plan 

form and the vortex of the preceding blade 

J. BIRON 

Lecture describing tests No. 63/1369 AY 442 

16. ON ERA 

Work performed in 1977 at the helicopter rotor test stand of 

wind tunnel S2 at Chalais-Meudon 

J. C. BOHL, P. LAFON, J. J. PHILIPPE 

Rt. No. 91/1369 AY 466 

17. ONERA 

Aerodynamics of helicopter rotors studied in 1979 

J. J. CHATTOT, A. DESOPPER, J. J. PHILLIPPE, H. WERLE 

RT of synthesis No. 106/1369 AY 

18. Design and wind tunnel testing of 1.5 m diameter model rotor 

A. BREMOND, A. CASSIER, J. M. POURADIER 

4th European Rotorcraft and Powered Lift Aircraft Forum 

September 13-15, 1978 Stresa, Italy. 

19. Aerospatial Survey Testing of Wind Tunnel--Testing of Small 

and Large Scale Rotor 

J. P. SILVANI, A. VUILLET 

7th European Rotorcraft and Powered Lift Aircraft Forum 

GARMISH (September 1981) 

25 
20. ONERA t 

Synthesis of studies on "the aerodynamics of helicopter rotors 

P. LAFON, A. DESOPPER, J. J. PHILIPPE 

RT No. 102/1369 AY 052 

26 

-... ~~ 
End of Document