Improvement of two blade sections for helicopter rotors

PAPER Nr.: l

IMPROVEMENT OF TWO BLADE SECIIONS FOR HELICOPTER ROTORS

BY

K.H. HORSTMANN, H. KOSTER

DEUTSCHE FORSCHUNGS- UND VERSUCHSANSTALT FDR LUFT- UND RAUMFAHRT (DFVLR)

BRAUNSCHWEIG, GERMANY

Ai'JD

G. POLZ

MESSERSCHMITT-BOLKOW-BLOHM GMBH

MDNCHEN, GERMANY

TENTH EUROPEAN ROTORCRAFT FORUM

AUGUST 28-31, 1984 -

THE HAGUE, THE NETHERLANDS

IMPROVEMENT OF TWO BLADE SECTIONS FOR HELICOPTER ROTORS K.H. HORSTMANN, H. KOSTER

DEUTSCHE FORSCHUNGS- UNO VERSUCHSANSTALT FUR LUFT- UNO RAUMFAHRT e.V. (DFVLR) BRAUNSCHWEIG, GERMANY

Abstract

G. POLZ

MESSERSCHMITT-B0LKOW-BLOHM GMBH MUNCHEN, GERMANY

In cooperation between OFVLR and MBB two new advanced airfoils for heli-copter rotor blades have been developed. In a second step these airfoils are now improved by further theoretical and experimental investigations. For the inner airfoil higher maximum lift coefficients at low Mach numbers could be obtained and for the tip airfoil the drag rise Mach numbers could be increased.

Some design aspects are discussed and the main experimental results are

pre-sented. Some rotor performance calculations show the improvement of rotor effi-ciency by these new airfoils.

1. Introduction

The nessecity of improving the aerodynamics of helicopters in order to increase cruise speed and payloads and to decrease the fuel consumption has clearly been seen by helicopter manufacturers and they made big efforts in this area for at least fifteen years. One of the most essential task was the improvement of the rotor performance first by using composite materials which allows the design of rotor blades with a spanwise evolution of the blade•s sections and second by the application of new airfoil shapes.

Since the end of the sixties and the beginning of the seventies research work on transonic airfoils for wings has been raised to get a better under-standing of the flow phenomena and to develop more reliable and simpler calcu-lation methods. As helicopter rotor blade sections pass flow regimes in which regions of supersonic flow on the airfoil contour occur, this advanced know-ledge has been used for developing improved rotor blade sections. In ref.[1] to [11] the problems of rotor airfoils are discussed in detail and some results of these efforts are given.

In 1981 a cooperation has been started between the Institute for Design Aerodynamics of DFVLR and the Helicopter Division of MBB for developing ad-vanced airfoils for new helicopter rotors. Two airfoils have been designed which are designated by DM-H1 Tb and OM-H2 Tb and have been investigated in the Transonic Wind Tunnel Braunschweig of DFVLR [12]. These airfoils fulfilled almost completely the stated requirements in regard to aerodynamic performances and moment behaviour.

The experience which has been gained during the design process of the new airfoils and the comparison of theoretical and experimental results offered the possibility to improve or to change some characteristics of these airfoils in view of a higher degree of adaption to rotor requirements. In order to realize the possible improvements the airfoils have been modified in a second development step.

2. Features of airfoil improvement

Depending on the mission profiles of a rotor aircraft a lot of parameters have to be specified determining the operational conditions of a blade section:

rotor tip speed

radial position of a blade section

blade twist distribution

rotor disc area loading

rotor solidity

inclination of the blade tip plane flight speed

The influence of some of these parameters on airfoil design has been

dis-cussed in ref. [12]. The main airfoil requirements for different flight condi-tions of a transport rotorcraft at the radial blade posicondi-tions r/R = 0.8 and 0.95

are summarized as design objectives in following table:

design objective inner airfoil tip airfoil

thickness ratio tic 12% 9%

drag divergence M > DDo 0.79 M DDo > 0.84 drag at M = 0.6, CL=D.7 _CD ~ D.Ol _CD ~ 0.01 maximum 1 i ft at M = 0.3 _{cl max} = 1.5 M = 0.4 1. 45 _clmax = 1.3 M = 0.5 1.3 1.2

pitching moment below _{lc I :;;}

0.01 lc I ;;; 0.01

stall inception m m

In the figs.l and 2 the design objectives and the main operational con-ditions of the two blade sections are characterized by the shaded areas.

Addi-tionally the measured performance boundaries of the airfoils OM-Hl Tb and

DM-H2 Tb are included. The stated requirements are completely fulfilled except in the case of maximum lift at the Mach number M = 0.4 of the airfoil DM-H2 Tb

for inner blade positions, fig. 1. The dent to lower values in the maximum lift coefficient curve in this important Mach number range reduces the airfoil per-formances noticeably.

In the range of zero or small negative lift coefficients at high Mach

numbers for the airfoil DM-Hl Tb, fig. 2, the design requirements are fulfilled too. But due to the high dynamic pressure at the tip of the advancing blade and the high drag caused thereby it is desirable to shift the drag rise to somewhat higher Mach numbers. This can be realized by slightly reducing the maximum lift coefficients of the airfoil DM-Hl Tb which are higher than stated by the design

objectives.

The advantage to change the DM-Hl Tb and DM-H2 Tb airfoil performances

in the manner discussed above can also be seen in fig. 3, showing the maximum lift coefficient of several airfoils at the Mach number M = 0.4 versus drag divergence Mach number at zero lift coefficient. The values of the airfoils

1.00

0.75

0.50

0.25

0

DFVLR- TWB- Tests

Re=8·10

6

·M

Azimuth Angle ljl=0°/180° lji:9QO

c

DM-H2Tb r/R:0.80

-0.25

L _ _ _ j _ _ . . . . J _ _ J _ _ _ _ j _ _ _L_--='--j---"

0.2 0.3

0.4 0.5 0.6

0.9

Fig. 1 Main operational conditions and measured performance boundaries for the inner airfoil

1.6

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

1.2

1.0

NACA \ \ DM-H2 Tb(12%)

A,/~

DM-H1 Tb(9%) 23012 ...

7'

...

I

'a..._

OA-Airfoils

~b

=

0

41

1.00

0.75

0.5

DFVLR TWB- Tests

Re=8·10

6·

M

0.25

Azimuth Angle ljl=0° /180° lji:9QO

0

-0.25

DM-H1Tb r/R:0.95

-0.5 0

L____L _ _ J _ _ L _ _ _ _ j _ _ _ _ _ L _ _

0.3 0.4 0.5 0.6 0.7

M

0.9

Fig. 2 Main operational conditions and measured performance

boundaries for the tip airfoil

DM-Hl Tb and DM-H2 Tb do not differ very much. To apply both

airfoils useful in a rotor blade

it is desirable to shift the

values in the directions marked by the arrows.

Modifying the airfoils

0.8

L _ _ _L_ _ _ i . _ _ J _ _ . . . . J _ _ _ j _ _ _ J

0.65 0.70 0.75 0.80 0.85 M

0.95

DDo

Fig. 3 Maximum lift coefficient versus drag divergence Mach number of several

rotor blade airfoils at M

=

0.4

the same methods are used as

des-cribed in reference [12]. For

subsonic calculations a modified

code from R. Eppler and D.M.

Somers extended to subsonic flow

[13], [14] has been used. This

code bases on a conformal map-ping procedure in its design

part and on a higher order panel/

boundary layer interaction method in its analysis part. For

tran-sonic flow the wellknown Bauer/Garabedian/Korn/Jameson method (BGK III) was used [15], [16] which also is coupled with a boundary layer method by adding

the displacement thickness to the airfoil contour.

First the desired improvement of the maximum lift coefficient of the

maxi--6~--~--,----.---,

1.4

M

1.2

c-P

-4

-2

Fig. 4

-1.0

-0.5

0

0.5

1.0

Fig. 6 DM-H2 Airfoil

IM

=

0.1.1

~ o TWB- Tests

L - - - -

_\ o \ Calculations: \~ Subsonic code ~ ~~ Transonic code (BGK nil

0.1

0.2

_0.3

_x/c

-4

1.0

-3

0.8

-2

-1

M

=

O.l.

0.1

_{0.2 x/c}

1.2

M

1.0

0.8

0.4

Upper surface pressure distri- Fig. 5

butions near the leading edge calculated with different

meth-Upper surface pressure distribu-tions of several airfoils

calcu-lated with the BGK III-method

[16]

ods and measured in the TWB

1.4

M

1.2

1.0

0.2 0.4 0.6 x/c 1.0

Upper surface pressure

distri-butions calculated by the

BGK III-method [16]

mum lift coefficients in general is

difficult. In the case of complete attached flow the usual potential

flow/boundary layer interaction methods give satisfactory results. Near maximum lift, however, the flow in general is separated and additional empirical criterions are useful to estimate the maximum lift coefficient, e.g. to _compare the calculated suction peak with experi-ments at similar airfoils,

especial-ly in the case of an additional

tur-bulent separation bubble due to a shock at the end of a local

super-sonic zone. A further difficulty is illustrated in fig. 4, showing parts

of pressure distributions of the airfoil DM-H2 calculated by the

sub-sonic and the transub-sonic code and measured in a windtunnel at M = 0.4

near maximum lift. The transonic calculation itself and also compar-ison of transonic and subsonic cal-culation give no indication of a shock, in contrast to the experi-ment, which s haws a distinct shock. This lack of the transonic calcula-tion may be caused by a non

suffi-cient number of grid points (161 on the airfoil contour are used in this case). It leads to a wrong boundary layer calculation, shifting the predicted maxi-mum lift coefficient to a higher value and causing the dent in the c -curve

L max

in fig. l .

To improve the maximum lift coefficient a wellknown procedure, described

for the Mach number M = 0.5 by several authors, [2], [3], has to be carried

out. The small local supersonic flow field has to be extented as shown in

fig. 5. This is attainable by a higher curvature near the end of the

super-sonic region and corresponding changes of curvature up- and downstream. This modification influences, of course, the aerodynamic characteristics at other operational points. Especially at higher Mach numbers an additional expansion of the supersonic flow in the region of the higher curvature is caused,

fol-lowed by a stronger shock.

For the improved airfoil a moderate lift increase is chosen, marked by the solid curve in fig. 5. This is a good compromise for a smooth maximum lift coefficient curve for all the considered free stream Mach numbers and an only small reduction in the drag divergence Mach number at lower lift coefficients.

Just the opposite step has been done, modifying the thinner airfoil DM-Hl Tb. Fig. 6 shows the reduced local Mach numbers in the supersonic zone

caused thereby and the resultant reduced wave drag calculated with the BGK III code.

The two modified airfoils provided with a tab and designated by DM-H3 Tb and DM-H4 Tb with a thickness to chord ratio t/c

=

0.09 and 0.12 respectively are shown in fig. 7. The tab lengths amount to 5% of chord length for the OM-H3 Tb and to 4.5% for the

DM-H4 Tb airfoil. The essential

contour differences to the

DM-Hl and OM-H2 airfoils

are exhibited in fig. 8.

.0

DM-H3Tb

t/c

₌

0.09

_-.1 .0 • 1 .2 .3 .1

DM-H4Tb

t/c

₌

0.12

c

_---=-

DM-H2

.0

fig. 7 Helicopter rotor blade

airfoils 0M-H3 Tb and

DM-H4 Tb -.1 _{,0 .I .2 .3}

Fig. 8 The essential contour

diffe-rences between the airfoils

3. Results 3.1 Windtunnel

The experimental investigations have been carried out in the Transonic

Windtunnel Braunschweig (TWB) of the OFVLR [17]. The wind tunnel is of the blow-down type with a rectangular test section of 34 em by 60 em with slotted walls at the top and the bottom and is especially suited for airfoil tests at subsonic

and transonic flaws in the Mach number range of M = 0.3 to 0.9. All wind tunnel tests described in this paper have been carried out with natural transition at a

Reynolds number of Re = 8-lo 6·M.

3.2 Experimental Results

The aim of the modification of the OM-H2 airfoil to raise the maximum

lift coefficient at the Mach number of M = 0.4 has been attained as shown in

fig. 9 presenting the maximum lift coefficient vers. Mach number of the airfoils

DM-H2 Tb, DM-H4 Tb and OA 212.The maximum lift coeffient at M = 0.4 could be increased from c L max=1.36 for the DM-H2 Tb airfoil to 1.53 for the DM-H4 Tb

airfoil. In addition the cl max-values for M = 0.3 and 0.5 could be raised

con-siderably to cl max=l.64 and 1.33 respectively. On the other hand the zero lift

drag shown in fig. 10 increased for a small amount but the drag divergence Mach

number defined by dc₀/dM=0.1 at c = 0 was not changed and is considerably higher than that of the OA 212 airfoil. LThe total performances of the DM-H4 Tb airfoil

measured in the TWB are summarized in fig. 11 and compared with those of the

DM-H2 Tb airfoil. Experiments TWB/S3MA

1.5

'a..,

DM-H4 Tb - · ' Re=8·M·106

;-·~-.

OA212 ..._~ S3MA, Re=7·M·106

1.0

0.2 0.3 0.4 0.5

_M

0.7

Fig. 9 Maximum lift coefficients of

several airfoils with thick-ness to chord ratio t/c=0.12

0. 0 2 0

,---.---,---.--;r----,

0.015

0.010

Experiments TWB/S3MA

I

OA212

I

S3MA, Re=7· M·106· /

. /

i

0.005

... . - . f - - - _ . - - - ~ DM-H4Tb DM-H2Tb TWB, Re=8 · M ·106

0

0.5

0.6

0.7

0.8 M 0.9

Fig. 10 Zero lift drag coefficients of

several airfoils with thickness

to chord ratio t/c = 0.12

The desired improvement of the high speed behaviour, especially the in-crease of the drag divergence Mach number for the DM-H3 Tb airfoil has been obtained and is shown in fig. 12. The drag divergence Mach number at zero lift

could be raised from Mooo=D.82 for the DM-Hl Tb airfoil to M₀₀₀.D.846 for the

OM-H3 Tb airfoil. The decrease of the maximum lift coefficients shown in fig. 13

does not exceed the expected amount of 6cl=0.05 and the c values of the DM-H3 Tb airfoil are larger than those of the OA 209 airf5i"la.'

1.5 CL 1.0 0.5 Experiments TWB Re=8·106·M t/c = 0.12 - - DM- H4Tb - - - DM- H2Tb

0

-0.5L__J__...l.__L...___.L__..L___j _ _ j

0.2

0.3 0.4 0.5 0.6 0.7

M

0.9

Fig. 11 Measured performance boundaries

of the airfoils DM-H2 Tb and DM-H4 Tb

1.5

r , , , ,

-1.0

0.2

Experiments TWB/S3MA

...

- • ._

~

... DM-H1Tb

/---.

.,_

OA209 ... -...,_ S3MA. Re=7·

M·10~

DM-H3Tb TWB. Re=B·M·106

0.3 0.4

0.5

_M

0.7

Fig. 13 Maximum lift coefficients of several airfoils with thick-ness to chord ratio t/c=0.09

0.020

Coo Experiments TWB/S3MA

I

0.015

CL: 0

~

I

OA209

fj

0.010

S3MA. Re=7·M·10~

>.·/

/

.

\

0.005

_DM-H1Tb DM-H3Tb TWB. Re=B·M·106

0

0.5

0.6

0.7

0.8 M 0.9

Fig. 12 Zero lift drag

coeffi-cients of several airfoils with thickness to chord

ratio t/c = 0.09

1.0

,----.,----,---,---0.6

0.4

0.2

-0.2

Experiments TWB M

=

0.75

Re

=

6 ·10

6 OM-H3Tb ' -

.-

,..;::r---:;....o

__

,

-<....

"' OM-H1Tb

0.005 0.010

co

0.020

Fig. 14 Drag polars of the air-foils DM-Hl Tb and DM-H3 Tb

The main reason for the drag decrease at higher Mach numbers is the

diminishing of the shock strength on the upper side of the DM-H3 Tb airfoil

which first causes a smaller wave drag and second prevents the formation of a

turbulent separation bubble behind the shock or at least reduces the length of the bubble. This is confirmed by the comparison of the two polar curves of the airfoils DM-Hl Tb and DM-H3 Tb at the Mach number of M = 0.75 on fig. 14.

1.5 .----r--.----,---,----,---, 1.0 c₀=0.01 0.5 Experiments

0

Re=8·106· M tic = 0.09 - - - DM-H1Tb - - DM-H3Tb

-0.5'---'---'--__j_-_J_____i _

_.J

0.3 0.4

0.5

0.6 0.7

M

0.9

Fig. 15 Measured performance boundaries of the

air-foils DM-Hl Tb and DM-H3 Tb

0.01

0

Experiments TWB Re = 8·M·106

-0.02

--o-- DM-H1Tb OM-H3Tb

In the Mach number range of M

=

0.6 to 0.85 the drag limits of co

=

0.01 and 0.02 could

be shifted to larger Mach numbers as can be seen in fig. 15 in which the total

perform-ances of the two airfoils DM-Hl Tb and DM-H3 Tb are presented.

The pitching moment coefficients at zero lift cmo in dependance of Mach number of the considered airfoils are given in

fig. 16. While the airfoils DM-Hl Tb and

DM-H3 Tb show about the same behaviour, the

DM-H4 Tb airfoil yields more negative values compared with the DM-H2 Tb which

has been desired during the design process.

The cm evolution with lift coefficient at M = 0.4 on fig. 17 shows that the measured

values essentially do not exceed ~ 0.01

and that the tabs of the modified airfoils are more effective than those of DM-Hl Tb and DM-H2 Tb airfoils, that means the slope

dcl/dcm is more negative. 1.6 1.2 1.0

r-Experiments TWB

M=0.40 Re=3 2·10

6

-:f')J

?'oM-H1Tb I

f

DM-H2Tb

-0.0 3

L _ _ - - ' - - - ' - - - ' - - - ' - - ' - - - ' - - - r t - J

0.8

I

~

0.01

0

-0.02

0.2 0.3 0.4 0.5

0.6 0.7

M

Experiments TWB Re = 8·M·106 --<>-- OM-H2Tb DM-H4Tb

0.6

0.4

0.2

-0.03

L---'---'--~----'--'--~__J

-0.2

0.2 0.3 0.4

0.5

0.6

0. 7

M

0.9

-0 01

DM-H3Tb

~

t\

I I

~

I I <} I I

~

I

0.01 c

-0.01

m

0.01

Fig. 16 Zero lift pitching moment

for several airfoils

Fig. 17 Pitching moment evolution of

1.6

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

1.2

1.0

t/c:13% \ DM-H4 Tb 2, > 1 /0

O....""o...

DM-H2Tb 12%7-_{... ""a...} 9% ' OA-Airfoils

7~\

6~

I

M = 0.31

0.8L__J___j___..L__...L_.. _

_ J _ _ _ _ j

0.65 0.70 0.75 0.80 0.85 M

0.95

DDo

1.6

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

1.2

tlc~13% '\ e - DM-H4 Tb ~ ""12% NACA 23012 '"~ ~ DM-H2Tb

IS% Tab) 12% ... o--DM-H1Tb >g%

7-~

DM-H3Tb 9% ... OA-AIRFOILS

7

~

1.0

_IM

₌

_0.41

_6%b

0.8L__J___j___J___J___.J____J

0.65

0.70 0.75 0.80 0.85

M

0.95

DDo

1.6

,---,.---,.---.---....--..,--~ DM-H4Tb 0 / > 12% tlc~13 Vo ~

y

DM-H2Tb 12_%

l '...:

_OM-H1Tb ... ~ >9% NACA 23012

J':i:

DM-H3Tb (5% Tab) / 9% ' OA-Airfoils

'~

I

M = 0 51 7_{% 'o}

1.2

1.0

0.8 L

~-==~-~-L--~L---~--~6~%~·__j

0.65 0.70 0.75 0.80 0.85 M

0.95

DDo

Fig. 18 Measured maximum lift

coeffi-0.020

Coo

0.010

1-0.005

1-0

0.5

cients and drag divergence Mach number at zero lift for several airfoils

'

'

DM-H3Tb

'I

CL = 0 Re=8·M·106 Experiments TWB

/

--7

Calculation IBGK ID)

-Transition xlc=0.07

Re=6 ·106

'

0.6

0.7

M

0.9

Fig. 19 Comparison of measured and

calculated zero lift drag

coefficients of the airfoil

A comparison of the OM-airfoils with the OA-series and

the NACA 23012 (5% Tab) is

presen-ted in fig. 18 in maximum lift coefficient vers. drag divergence Mach number diagrams for the Mach

numbers M

=

0.3, 0.4 and 0.5. The

desired gain in CL max for the

DM-H4 Tb airfoil has been obtained especially for M

=

0.4 and also

for the other two Mach numbers without decreasing the drag diver-gence Mach number at zero lift.

The shift of the drag divergence Mach number for the DM-H3 Tb

air-foil has also been realized and the loss in maximum lift has been

limited to ~cL = 0.05.

The agreement between the calculated values of maximum lift and zero drag coefficients and the experimental results shown in

fig. 19 and 20 is acceptable but

does not replace accurate wind tunnel measurements.

1.8

1.4

1.2

1.0

DM-H4 Tb Airfoil -..._ /Experiments TWB.

/"'-'~Re=8·

M·106 Predicted

~

values

0.8'---L--'--'----l...-...1

0.2 0.3 0.4 0.5

M

0.7

Fig. 20 Comparison of measured and predicted maximum lift coef-ficients of the airfoil

I

-

~ Q) :::E

-

0 Q) ~

4. Influence on helicopter performance and load factor capability

The useful effects of the modified airfoils DM-H3 Tb and DM-H4 Tb have been evaluated based on the BO 105 main rotor by theoretical comparison with respect to performance and load factor capability with the foregoing airfoils OM-Hl Tb and DM-H2 Tb as well as with the NACA 23012 airfoil, which is used on the BO 105 production rotor blades.

For the comparison of the new profile types, the 12% airfoils OM-H4 Tb, OM-H2 Tb are applied up to 80% radius, with a linear transition to the 9% air-foils DM-H3 Tb, OM-H1 Tb at the blade tip. The twist angle of -8 deg ang the tip speed of 218 m/s is kept for the three rotor versions.

With the improved airfoils the figure of merit (fig. 21) is increased

by about 1% compared to OM-H1 Tb/-H2 Tb and by 10% as against the NACA 23012

airfoil, thereby raising the F.M. to a maximum of 0.77.

0.80,---.---...---.---,

DM-H3Tb/H4Tb

-"1'---·---._

...

-

\ '-...;-. DM-H1Tb/ H2 Tb

0.70

_....

_···"(·

_{•···•···}

_{... ..}

....•• / NACA 23 012

6,---,,---.---.---,

H

[kml

4

2 DM-H3Tb/H4Tb

....

···

_··....

...

_...

_...

...

...

/

···..•

_·..

...,.

_'

NACA

23012:·...

' ,

··..• '· DM-H1 Tb

I

··· ... \ H2Tb

....

\ •.

.

g,

0.60

B 0 105

_(\SAl

f

LL 8 0

105

Main Rotor

0.50

0.06

0.08

0.10

0.12

Blade load coefficient cT/6 Fig. 21 Influence of blade airfoils

on the figure of merit

OJ

2.0.---.---.---,

-"' DM-H3Tb/H4Tb

...

~

1. S D M- H 1

l!:_/

H :.,.2

T~b:,.-,...;.,._,.,_,_,___ -1

... ••••• •• ;•••••u·••• ••• .. ,,., Q) OJ c:

c

1.0 0:: u

-

u

0.5

Q) a. (f) ... NACA 23 012 .

BO 105

(1500m

!SA) 0

_{1oLo~---1~s~o---v-[~k-m_/_h_J~2SO}

Fig. 23 Specific range

o~,s~o---~z7_{oo~----~·~·~---73o·o}

V

_{1A5 [}

km/h]

Fig. 22 Maximum cruise speed

.~ 0.2

-

'<:) Q) a. (f)

50

80

105

(1500m

\SA)

100

_{150 v [ km/h]250}

The maximum cruise speed (fig. 22) at low altitude is increased by

about 3 km/h due to the improved airfoils and exceeds the figure of the actual NACA 23012 equipped rotor by 16 km/h. It should be noted that with increasing speed, not only the parasitic drag of the helicopter rises, but also the Mach

number of the advancing blade is increased, both effects normally leading to higher power required. The advantages due to the improved airfoils become more

significant at higher flight altitudes since the stall onset on the retreating blade, which affects also the power consumption of the rotor, is delayed by the

higher maximum lift capability of the new airfoils.

Similar improvements were achieved in the specific range, which is

in-creased by 2% over the DM-Hl Tb/-H2 Tb and 11% compared to the NACA 23012 in its optimum (fig. 23). The advantage of the modern airfoils is more evident with increasing flight speed, so that also the flight speed for optimum range

is increased. However, it must be said, that from a profile design point of view the optimum range state is characterized by only moderate conditions on

both the Mach numbers on the advancing blade and the high lift coefficients on

the retreating blade. Nevertheless for the operator reductions in specific fuel consumption can be turned into useful payload, hence reducing the overall oper-ating costs.

Similar advantages are to be seen at the optimum endurance flight

condi-tion (fig. 24), whereby improvements of 0.5% compared to the DM-H1 Tb/-H2 Tb profiles and 10% over the NACA 23012 are obtained.

At the flight speed of optimum rate of climb which is close to the

opti-mum endurance speed, a nearly constant increase in the optiopti-mum rate of climb

of 0.2 m/sec respectively 1.2 m/sec is achieved over the whole altitude range (fig. 25).

The maximum load factor envelope (fig. 26) of the BO 105 rotor is also

improved by the new airfoils. The advantage, compared with the former airfoils DM-Hl Tb/-H2 Tb, mainly comes from the increase in maximum lift capability of

the DM-H4 Tb airfoil at the inner blade section. Due to the improved airfoils the maximum load factor is increased by .2g over the DM-Hl Tb/-H2 Tb and .4g over the NACA 23012. Correspondingly the maximum achievable speed for a specific load

factor is increased. 6~~k-~----~----~---,

_{··· ..}

H

[km]

4

2

0

-4

'< DM-H3Tb/H4Tb

·· ...

...:~ ···."~

··•· ..

~

.

...:

·.

_·..

"'

_"

DM-H1Tb

I

H2Tb

/

·

..

_•·

"

_...

"

NACA 23012"· .. "

BO

105 ( ISA)

4

·.

_·.

"

_'

•··.

_·

_..

~ ~

1

8 12

Vz[m/s]

Fig. 25 Maximum rate of climb

'

3 0> c 0 2 u 0 lL ' , DM-H3 Tb/H4 Tb

··. ',, I

·· ...

/'

/

···

...

'',,,,

·.

_·.

' _' DM-H1 Tb /HZ Tb / · · ...

~~

..

"'

0 NACA 23 012 0 _J 8 0 105 (Om IS A) 0 0.10 0.20 0.30 0.40 Advance Ratio _~

Summary

The rotor airfoils DM-Hl Tb and DM-H2 Tb resulting of a cooperation

bet-ween MBB and DFVLR are improved in a second development step. Aims and features of desired modifications as well as contour changes are discussed.

At the inner airfoil the maximum lift coefficient at M = 0.4 should in-crease and the drag rise of the blade tip airfoil should be shifted to higher Mach numbers. The new airfoils designated by DM-H3 Tb and DM-H4 Tb are investi-gated in the windtunnel. The presented results show that the desired

improve-ments are obtained. The maximum lift coefficients of the airfoil DM-H4 Tb raise

up to values of CL max=l. 64, l. 53 and l. 33 at Mach numbers M = 0. 3; D. 4 and 0. 5

respectively. The reduction of drag rise Mach numbers are very small.

The drag divergence Mach number at zero lift of the DM-H3 Tb airfoil is shifted to a value of MDDo=D.846 while the maximum lift coefficients are slightly reduced to values of cL max=l.28 and 1.18 at the Mach numbers M = D.4 and 0.5 respectively.

From the comparison, based on the BO 105 rotor, it has been demonstrated

that significant improvements are to be found at all flight conditions with the new DM-H3 Tb/-H4 Tb airfoils, when compared with the initial DM-Hl Tb/-H2 Tb advanced airfoil design and the classical NACA 23012 profile.

References

[l] G. Reichert and S.N. Wagner, Some Aspects of the Design of Rotor Airfoil Shapes, AGARD-CP-lll, 1973, Paper 14.

[2] F.X. Wortmann, J.M. Drees, Design of Airfoils for Rotors, Paper ~resented at the CAL/AVLABS 1969 Symposium on Aerodynamics of Rotary Wing and VTOL

Aircraft, Buffalo, N.Y ..

[3] J.rJ.Sloof, F.X. Wortmann, J.M. Duhon, The Development of Transonic

Air-foils for Helicopters, Paper presented at the 31st Annual National Forum

of the American Helicopter Society, Washington D.C., May 1975.

[4] R.W. Prouty, A State-of-the-Art Survey of Two-Dimensional Airfoil Data.

AHS Symposium on Helicopter Aerodynamic Efficiency, March 1975.

[5] J. Renaud and F. Nibelle, Effects of the Airfoil Choice on Rotor

Aerody-namic Behaviour in Forward Flight. Paper presented at the 2nd European

Rotorcraft and Powered Lift Aircraft Forum, BUckeburg, September 1976. [6] L. Dadone, Rotor Airfoil Optimization: An Understanding of the Physical

Limits, Paper presented at the 34th Annual National Forum of the American

Helicopter Society, May 1978, Washington D. C., Preprint 78-4.

[7] J.J. Thibert and J. Gallet, A New Airfoil Family for Rotor Blades, Paper

presented at the 3rd European Rotorcraft and Powered Lift Aircraft Forum,

Paper No. 41, Aix-en-Provence, September 1977, T.P. ONERA 1977-131.

[8] J.J. Thibert and J. Gallet, Advanced Research on Helicopter Blade Airfoils, Paper presented at the 6th European Rotorcraft and Powered Lift Aircraft Forum, Paper No. 49, Bristol, September 1980, T .P. ONERA 1980-93.

[9] J.J. Thibert and J.M. Pouradier, Design and Test of a Helicopter Rotor Blade with Evolutive Profile, 12th ICAS Congress,Munich, October 1980, T.P. ONERA 1980-125.

[10] L. Dadone, The Role of Analysis in the Aerodynamic Design of Advanced Rotors, AGARD-CPP-334, Paper l, May 1982.

[ll] J.J. Thibert and J.J. Philippe, Etudes de Profiles et d'Extremites de Pale d'Helicoptere, AGARD-CPP-334, Paper 3, May 1982.

[12] K.H. Horstmann, H. Koster, G. Polz, Development of Two Airfoil Sections for Helicopter Rotor Blades. Z. Flugwissen. Weltraumforsch. 7(1983), pp. 82-91; Paper presented at the Eight European Rotorcraft Forum,

Aix-en-Provence, France, August 31st to september 3rd, 1982.

[13] R. Eppler and D.M. Somers, A Computer Program for the Design and Analysis of Low-Speed Airfoils, NASA TM 80210, 1980.

[14] R. Radespiel, Erweiterung eines Profilberechnungsverfahrens im Hinblick

auf Entwurfs- und Nachrechnungen von Laminarprofilen bei

Verkehrsflug-zeugen, DFVLR IS 129-81115, 1981.

[15] F. Bauer, P. Garabedian, D. Kern, A. Jameson, Supercritical Wing Sections

II, Springer-Verlag, Berlin, Heidelberg, New York, 1975.

[16] F. Bauer, P. Garabedian, D. Korn, Supercritical Wing Sections III, Springer-Verlag, Berlin, Heidelberg, New York, 1977.

[17) E. Stanewsky, W. Puffert-MeiBner, R. MOller, H. Hoheisel, Der

Trans-sonische Windkanal Braunschweig der DFVLR, Z. Flugwissen. Weltraumforsch.