TWENTYFIFTH EUROPEAN ROTORCRAFT FORUM

Paper n° P3

ANALYTICAL AND EXPERIMENTAL STUDIES OF

TWO TIP-SWEPT BLADE VERSIONS

BY

Dr. V. IVTCHIN
MIL MOSCOW HELICOPTER PLANT, RUSSIA
Dr. A. LISS
KAZAN SCIENTIFIC INDUSTRIAL ENTERPRISE, RUSSIA

SEPTEMBER 14-16, 1999

ROME
ITALY

ASSOCIAZIONE INDUSTRIE PER L'AEROSPAZIO, I SISTEMI E LA DIFESA

ASSOCIAZIONE ITALIANA DI AERONAUTICA ED ASTRONAUTICA
 ANALYTICAL AND EXPERIMENTAL STUDIES OF TWO TIP-SWEPT BLADE
VERSIONS

Dr. V .A. Ivtchin

Team Leader, Aerodynamics Dept
Mil Moscow Helicopter Plant
Russia

Dr. A. Y. Liss
Head, Aerodynamics Dept
Kazan Scientific Industrial Enterprise
Russia

The paper presents some basic results of analytical studies and flight tests for two swept-tip blade
versions: for a swept tip having no twist without anhedral, and an anhedral swept tip. The target
objective of the anhedral swept-tip blade development is to fmd a way to reduce pitch link load and
control loads and, at the same time, to improve aerodynamic characteristics of helicopter main rotors in
hover and in level flight. The contribution to the reduction in the pitch link load was theoretically
substantiated by Yu. A Liss (Kazan) in 1984. The tapered tip shape tested on full-scale the main rotor
blades for the Mi-28 attack helicopter and its possible alteration to have an anhedral blade tip were
offered by VA Ivtchin (Moscow) in 1985 and 1993 respectively.
Within the framework of a research progranune the Mil Moscow Helicopter Plant (MMHP) developed
all-composite main rotor blades for the Mi-28 attack helicopters and conducted comparative flight tests
of two versions of the swept-tip main rotor blades.

I. NOTATION blade chord plane, m

R~ swashplate radius, m
R main rotor radius, m lJ blade pitch arm radius, m
0) rotor angular velocity, s-1 C1max airfoil maximum lift coefficient
b, blade chord at 0. 7R, m c,, profile drag coefficient
r design cross section radius, m Cy" angle of attack airfoil lift coefficient
ZI number of blades derivative
(j rotor solidity ratio along 0.7R a. angle of attack airfoil, deg
a, coning angle, deg K, blade section lift-drag ratio
D1,D2 swashplate kinematics coefficient mw longitudinal moment coefficient at airfoil
Xi blade rotation axis in feathering hinge zero lift
relative to blade nose coinciding with axis xr airfoil aerodynamic centre relative to blade
of spar stiffuess, m. nose,m
equivalent rotor angle of attack in blade tip p air density, kg/m3
plane, deg ooR blade tip speed, m/s
'I' rotor blade azimuth, deg M Mach nlllllber
Itip tip length along blade longitudinal axis, m v airspeed, kmlh
ftip radius of tip center of pressure relatively to v, radial component of flow velocity along
rotor axis of rotation, m blade longitudinal axis, kmlh
bti, tip inboard chord, m P,1 constant portion of collective pitch sleeve
botip tip outboard chord, m loads, n
Xtip tip sweep angle along leading edge, deg; + M"w total swashplate constant moment, Nm
sign mean backward M~ constant portion of swashplate longitudinal
~;,, tip anhedral, deg, + sign means downward moment, Nm
a. tip tip angle of attack, deg Mxsw constant portion of swashplate lateral
a.avtip average anhedral tip angle of attack, deg moment, Nm
a.vtip tip angle of attack versus blade azimuth and Pm amplitude of main rotor blade pitch link
airspeed, deg load, N
tip area, m2 Pmo amplitude of constant portion of main rotor
location of tip pressure centre along blade blade pitch link load, N
longitudinal axis relative to beginning of Prole coefficient of cosine component of main
tip, m rotor blade pitch link load 1/rev, N
X tip location of tip pressure centre relative to Prot;; coefficient of cosine component of mam
blade axis about feathering hinge, m rotor blade pitch link load !/rev, N
Ytip location of tip pressure centre relative to Prot amplitude of main rotor blade pitch link
P3-1
 load llrev, N model rotor conducted by M.G. Rozhdestvensky [2].
Vn component of flow velocity normal to blade
The tapered tip shape testes on the full-scale
chord plane, kmlh
main rotor blades of the Mi-28 combat as well as the
&m amplitude of variable portion of main rotor
possibility to replace this tip shape by an anhedral one
blade pitch link load, N
were proposed by V.A. lvchin in 1985 and 1993
respectively.
2. INTRODUCTION
Helicopter high aerodynamic performance is The results of the analytical studies presented
achieved by improving the aerodynamic configuration of in the paper show the main principles lying in the effect
the main rotor helicopter blades. The most effective produced by the helicopter rotor blade anhedral tip.
trend here is to develop new blade airfoils possessing
high lift coefficient and LID ratio. However, the The results of the flight tests conducted by the
introduction of new airfoils in the conventional Mil Moscow Helicopter Plant on the Mi-24 flying test
rectangular blade planfonn results in quite a significant bed fitted with the Mi-28 rotor system in which both tip
increase in control loads. Ref. 1 presents criteria to be shapes, i.e. a tip having no twist and an anhedral tip, are
used in estimating helicopter control loads versus the presented. They substantiated the concept of the
main rotor figure of merit for Mil helicopter rectangular anhedral blade tip in real operating conditions of the
blades. At the same time it shows that better main rotor.
aerodynamic performance of rotors having rectangular The reduction in the pitch link loads caused by
blades results in a considerable growth of control loads. the application of the anhedral rotor blade tip leads to an
This problem becomes more acute when critical loads increase of the helicopter control system service life. It is
are found out in the process of full-scale flight tests well known that the service life of any component is
(when the blades themselves and the required tooling proportional to the value of the component loads to the
have already been manufactured). Therefore a proper sixth power. Therefore, a reduction in the control system
selection of the blade geometric and aerodynamic loads by 10 % in cruise can lead to an increase in the
configuration on the design stage is of paramount service lives of the rotating parts of the swashplate by 3-
importance. 4 times and those for bearings, by 6-8 times.
Ref 1 presents the results obtained from
studies devoted to a new component in the blade 3. MI-28 EXPERIMENTAL MAIN
configuration, i.e. the "glove" on the Mi-28 full-scale ROOTOR BLADE AERODYNAMITC
rotor made of composite materials. The «glove" is a CONFIGURATION AND CHARACTERISTICS OF
change in the blade root configuration due to which the PITCH LINK LOADS
centre of the blade root section has bee moved forward
in relation to the feathering hinge axis. This feature has While designing the main rotor for the Mi-28
allowed the pitch link load and control loads to be attack helicopter, the Mil Moscow helicopter plant in
reduced by 25-40%. collaboration with the TsAGI developed a new
aerodynamic configuration for that main rotor. It was
The main objective in the development of an conceived at the very beginning of the work that the
anhedral main rotor blade tip is also an attempt to fmd main rotor having that new aerodynamic configuration
new ways of reducing the pitch link load and control could be used to upgrade the Mi-24 helicopter, that was
loads with improved helicopter main rotor performance why the diameter, tip speed and the nwnber of the
in hover and level flight. This blade tip was developed blades were taken as those of the prototype helicopter.
and tested by the Mil Moscow Helicopter Plant in The general arrangement of the blade is shown in Fig. 1,
several steps to be used in the Mi-28 attack helicopter while its geometric twist, in Fig. 2.
all-composite blades.
These blades featured a high twist with a
A favourable effect of the anhedral tip on the gradient t-<p =-9.7', new airfoils TsAGI SB(-6*6) and
rotor relative figure of merit in hover was obtained from KS( -4.5*6), composite materials. As the new airfoils
the Mi-26 eight-blade main rotor whirl tower tests have trailing edge plates, the blade chord became wider
conducted at M.N. Tischenko's initiative in 1978 [12]. by 0.04 m (the width of the plates) as compared to that
The aerodynamic effects caused by the tip of this type, of the Mi-24 rotor blade. To improve the rotor
from the point of view of reducing the cosine component aerodynamic performance in hover, the blade tip starting
of the pitch link load, were theoretically proved by at 0.9R was twisted by 0.9°, as shown in Fig. 2. The
A Yu. Liss in 1984. These were followed by parametric same figure shows the Mi-24 production main rotor
studies of the anhedral tip angle by using a four-blade geometric twist for comparison.

G~====B~65=0==========~1]~2~15
Fig.1

P3-2
 M=0.6, especially the SB lifting airfoiL A little higher
value of C,, for Ct=O and Ct=O for the SB airfoil can be
' attributed by its 1.7% higher thickness ratio. However,

( 6
~
-- Mi-28 this airfoil is a lifting one and it runs up to R=O. 9 of the

4
.... ...........
_________
II --- Mi-24 blade, and even at maximwn helicopter speeds Mach
number of such values \vill never occur. The application
of these airfoils as well as a higher geometric t\vist of
2 ---------- ...... the blade have contributed to better figure of merit and

0
~k LID. It should be noted here that a favourable effect of
the KS tip airfoil starting at 0.9 R is a significant
reduction in the &nzo18M derivative at a critical Mach
-2 """"' ~ number of 0.85. As can be seen from Table I, the m,

'
value for the NACA-230M airfoil reduces by -0.035 with
-4 the increase of Mach number from 0.6 to 0.85, while
0 .2 .4 .6 .8 r that for the new airfoil, by -0.014 .

Fig. 2 To investigate the new main rotor, a flying tested

based on the Mi-24 was made. It was fitted with an

Table l

Airfoil K, I l<..NACA Chna:/ Cx, am,!aM m, m, ac,Jacx xr

ClmaxNACA
c,=o.s c,=o c,=o C,=O
M=0.6
M=OA M=0.9 M=0.85
M=0.6 M=0.6 M=O. M=0.7
6
NACA-230M I I 0.042 0 0 -0.035 6.1 0.21
SB(-6*6) 1.21 1.05 0.055 -0.03 -0.002 -0.049 7.0 0.25
KS(-4.5*6) 1.26 0.94 0.038 -0.01 +0.014 0 6.8 0.23
..
The new blade was fitted With a tlp famng expenmental set of compos1te rnam rotor blades from
featuring the leading edge \\lith a 30° sweep and a the Mi-28 and its wing consoles were removed.
relative width equal to 2.5% of the radius to reduce the
The first test results of the initial version of the
noise level and wave drag produced by the blade tip. The
experimental blade set showed that the main rotor
rotor blade on the prototype was fitted with a rounded
control loads had substantially increased as compared to
tip of conventional shape, but of lesser size (by about
1%). those of the Mi-24 production blades. The level of
constant forces of control system was so high that an
New airfoils developed by the TsAGI in installation of more powerful hydraulic actuators
particular were used in the new rotor design. Their /boosters/ had to be made and new design features had
aerodynamic performance was obtained in the T sAGI T- to be found to reduce pitch link loads.
104 wind tunnel. The main data are presented in Table I
in comparison with the NACA-230M airfoil with a l 0% Fig. 3 shows comparison made for blade pitch
thickness ratio, which was used in the Mi-24 production link loads versus airspeed for two sets of blade [3].
Pitch link loads for the Mi-28 composite blades are
main rotor blade. The new airfoils feature quite a wide
presented for one of experimental sets that later on was
plate numing along the airfoil trailing edge to adjust the
value of longitudinal moment. The plate parameters are used with non-anhedral and unheard swept tips. The
curves demonstrate that the gradient of the constant
shown in the airfoil designation in the brackets (the first
derivative of the new blade pitch link moment versus
digit refers the plate bend angle in degrees, the minus
sign means upward bending), and the second digit, the airspeed has a negative sign. The value of the sine
component of the pitch link load per revolution for the
plate size in the main chord percentage. For the
experimental set of the composite blades, the metallic new rotor gas a riluch higher dive gradient with airspeed,
leading edge plate was used to change the bed angle in that that for the Mi-24 production rotor. These changes
the process of flight tests. The plate is made of in the components of the pitch link load fore the new
composite materials for the production blades, and it blades can be attributed to a more i..rtboard and stable
does not change its the bend angle in operation. location of the SB and KS airfoil aerodynamic center
relative to the NACA-230M airfoil, as well as to a more
As can be seen from Table 1, the new airfoils substantial dependence of the m20 value upon airflow
with the reference bend angles of the plate have a 21- Mach number. For the Mi-24 production main rotor
26% better LID ratio. The value of the longitudinal blade, the positive gradient of the sine component of the
moment coefficient for airfoil KS is higher by pitch link: load per revolution with airspeed is
( Ll.mw=+O.Ol4, while that for lifting airfoil SB is lower by determined by the aerodynamic center of the NACA-
Ll.rn,=-0.005 than for airfoil NACA-230M. The new 230M airfoil shifting forward with Mach number
airfoils have a much greater derivative &nu/8M at increasing on the advancing blade.

P3-3
 1500
8000
1250 ·········~·- ···:············ ···-~·····
7000 ...... ... ...•. .. . .. ·-~-- ·········: ....
1000 ...... .:. ........... ,:.
6000

750 ···········•.· ............:···· ········~·-··

5000
c

.
0
E 500 .
··········-~- .. .:.............; ...
.,"-E 40oo
............ ; ........... ; ........... ;
·;;9·····
250
··············~4-
,h....._,- '
3000 ·····:

0 2000 ··········-~---- --~

-250 1000 ···········+··········+··········f ·--~ ·······~·-·········

-500
50 100 150 200 250 300 350 50 100 150 200 250 300 350
Speed, krnlh Speed, km/h

2000 4000

1500 ···········-#·-~~. 3500 . . ,. . . . . l.. . . . . L. . . . . L. . . . ..l...

3000
1000
E c 2500 ...........;..... !;.;;;;;:...:;:.a . -~--
c
0

..'E
vi 500
.E 2ooo
1500 ··:············
o-i--..,f----i---+-+__:;.;---1
\
1000 ···if#~~
-5001············'············'············0············0············0·\~
500 ··~· ..........•...... ························
-1000 ...___ _.___ __.__ _.__ ___.;_ _._,'-J
o.J-----i-----i---i---i---i----1
50 100 150 200 250 300 350 50 100 150 200 250 300 350
Speed, km/h Speed, kmlh

Fig. 3
As can be clearly seen from the curves in Fig. 3. inspection methods, etc. Therefore the results obtained
the sine component of the !/rev and the variable portion from the flight tests raised an issue of fmding relatively
of the pitch link load differ quite substantially. For the cheap and fast means to alter the experimental blades
experimental blades, the values of these parameters have (without manufacture of expensive and sophisticated
increased by 2-2.5 and 1.7 times respectively at a the tools and fixtures) to reduce the flight control system
maxinuun speed of 300 kmlh. The increase in the cosine loads.
derivative of the pitch link load !/rev is attributed by a
To study the possibility of a reduction in pitch
higher geometric twist of the e>.:perimental blades
link loads, a programme of introducing design changes
increasing this component of the blade pitch link load
in the experimental set was drawn up m1der Professor
when the blade sections are in the slipstream.
M.N. Tischenko, General Designer. It resulted in solving
The geometric twist of the new blade is 1.67
the above problems at the lowest costs and with
times higher than that of the Mi-24 production blade.
maximum reliability, as the test results were obtained
Another reasons for increasing the experimental blade
from tests on full-scale blades. A series of design,
pitch link loads is as follows: they have a wider chord
analytical and experimental studies on the instruction of
and lower torsional elasticity. This is due to the
the Mil Moscow Helicopter Plant was carried out in the
application of composite materials as well as the
TsAGI Helicopter Division within the frame of the above
requirement to develop blades of minimum weight. The
progranune. The were oriented to study possibilities
application of composite materials leads to an increase
leading to reduce pitch link loads in the blades with a
in the thickness ratio of the blade root end sections by 7-
new aerodynamic configuration provided the main rotor
8%, which, in its turn, contributes to a growth of the
aerodynamic performance remains the same or even
blade pitch link loads.
improves. In those studies, both conventional methods of
Blade manufacture, and that of composite blade correcting blade pitch link load characteristics (control
in particular, is a sophisticated and expensive process. It of the airfoil mass characteristics by selecting a rational
requires a long-term pre-production stage, manufacture spamvise distribution of the tail trinuning plate bending
of special tools, fixtures and equipment, development of angle, varying the blade cg position, etc), and an

P3-4
 absolutely new method developed by the TsAGI and
related to the application of a new root end fitting shale
were considered. The results of those studies were
presented at the 1996 European Forum [l]. Further ~~ =Pmo +Pmls *sin¢+ P,llc *CO&/J +L1~hg'
studies conducted in the direction of reducing pitch link
loads within the framework of the above programme where k Ra,,Z, is a constant depending upon
=

involved changes in the blade tip of the experimental 2

the control system geometry and number of blades.
composite blades.
The Pmls component can be easily affected by
changing the value of the blade airfoil lateral moment
coefficient Cmo· It can be achieved, for instance, by
4. THEORETICAL ANALYSIS OF THE
bending the trim tabs running along the blade trailing
EFFECT PRODUCED BY THE NUUN ROTOR
edge. Similar simple means of affecting the P mlc
BLADE SWEPT TIP ON ITS PITCH LINK LOADS
coefficient at llrev COS'-V in the Pm. Furrier-series
The analysis of the flight test results has shown expansion are unknown yet. TI1erefore in some cases the
that the blade tip portion occupying about l 0% of the attempts at reducing the constant of the moment acting
radius produces about 40% of the cosine component of on the swashplate failed and the blade design already
the pitch link load in flight at cruise and maximum developed and manufactured had to be rejected
speeds. This is attributed to the design features of the (composite blades for the Mi-28 developed in the 80s).
blade tip portion, airfoil variable geometry, as well as
The blade tip shown below is swept and bent
the features of the flow around them at high speeds. The
downward relative to the chord plane (of a wing
simplest and most widely used teclmical means to
anhedral type), it is an effective maenad to affect the
Mw1c value, thus the constant portion of
R the swashplate moment. Fig. 4 presents
a version of this blade tip. At frrst, the
other author of this paper, V.A. !vchin,
--...: Yti

tip center
of pressure /
v ~ ----..)_
l;ti pp
suggested another flat tip planfonn. It
was selected proceeding from the
analysis of the experimental results
obtained in studies for the main rotor

~p
blade tip shape that had been conducted
both in Russia and abroad. The main
( . ' tip center
of pre~

~
objective of the development of this type
btip - Xtipp
---...... of the blade tip was improvement of the
1/4 chord Xtipp
/ I main rotor aerodynamic and noise
perfonnance and some lowering of blade
/ pitch moments. The blades with these
tips were tested at the Mil Moscow
rtipp ~ Helicopter plant [6]. Let us consider the
principle of operation of a swept
anhedral tip. Let the helicopter fly at a V
Fig. 4 speed the main rotor of attack a.HB
improve blade pitch link load characteristics in the main relative to the blade tip plane. The radial
rotor control system is upward bending of the trim tabs component of the V speed along the blade axis equals
nnming along the trailing edge. However, this has not Vr=Vcosa.cos(~.V-X), and the airflow component nonnal
produced the desired effect as this action provides to the tip axis is Vn=Vcosacos(ljf-y:)sin(l;ti,-!!o). The
correction of the constant portion and sine component of above statements are illustrated by Fig. 5.
the blade pitch link load !/rev only. Besides, it leads to
quite a substantial deterioration of the blade airfoil
aerodynamic perfonnance and, thus, to a lower figure of
merit and efficiency of the main rotor. <"110'
A search for new design features aimed at
reducing the blade tip hinge moments is an important (., v
task whose solution can help to lower loads in the main +--
rotor control system, increase its service life and widen
the helicopter flight envelope.
v, I a
One of tl1e new designs was conceived and
theoretically proven by one of the authors of this paper,
\'
~~J.,
A. Yu Liss, in 1984. The main concept was to use a v,
( swept tip bent downward. As has been stated in [Ref. l],
main rotor blade pitch link load l /rev de lines the
swashplate moment constant: Fig. 5

P3-5
 The airflow velocity in the main rotor tip plane 0.06
and normal to the blade tip is defmed by the value 1-····C········'·······'········'·······c....... ;.•.•....i ........ i;:Q.:
Vr=ror+Vcoscx*sin('t'-X)- Then the angles of attack in the o.os j ...... ;.....;....... ;........ ;........; ....... ;........;/·-/··;········>·
/ .. . ''
design cross-section of the blade tip CXtip for level flight at
the V speed could be roughly defined by the following 0.04 j j ....... i j !...... i '
i/-··········<·······C··'· '
equation: 't" ...... c........ c....... c......... c..... ·'······-// .... .L.: ..
! 0.03 j ....... j........ ;........• ·····C·········i···-/1·······-:- ······1 ....... [.
:
+·
!

v 1-··'·······'·····'·········'·······'// ······"······<·······•·+·
a tip
v, =
= aotip + afj·ip = aotip + -" o.oz
1 ; ..... : ......;.../·_· v~_7 ...~ ::::::1 ·······f --~ ·······r·
__-,__ f'
_ V *cos a* sill(~ -a,)* cos(l{f- X) 0.01 •••.••••• ····- •••••••••••• -1-·

- aotip + /:
1-·····'····---·····,'/·-··<· I
......,......, ......, .......,..:··
«1r + V *cos a"' * si11( If/- X) 0.00

Taking into account only the constant portion and 1/rev

0 ~ u u u u u u u u ,., ...
of the aerodynamic forces, the blade tip lift could be Fig. 6
roughly defined by the following equation:

-
Y,;p - c:(aotip + ll\-;;p 2p~~ stip =
P V cosasi11'l'-a
1 )cosli"-x)J [ . ]2
=C'-S. a.+ " • l'i" xor.. +Vcosasm(l{f-X)J =
Y 2 up[ ''P artip + V cosasill(l{f- X) up

=C: p S,;p(v cosasill(~-a.)cos(l{f- x)x[artip +Vcosasi11py- xJl+a.,;px[artip +V cosasill(lf/- xJJ'}=

2

=c; P Stip(v cosaortipsill(~-a.)cospy- x)+a.,;px[artip +Vcosasill(lf/- xJJ'}

2
As can be seen from the equations obtained, the
while drag could be roughly defmed by the following application of a swept tip without any anhedral with the
equation: design sweep angel equal to Xtip changes the constant
pv; P r
X~~p=Cz.--' Sq=Cz"-S~~pxlar~~p+Vcosasin(I!'-X)J =
~ portion and the tip sine component of the blade pitch
2 2 link load 1/rev to negative pitching, while the cosine
component to positive pitching. If the blade tip is bent
=Cz., 1Su, x[rarqf +0.5(V cosaf +2arlfp * sin(w- xJ}
downward by an angle !;tip, only the constant portion and
sine component of the blade pitch link load !/rev
As can be seen from the above equation, lift for
change. The cosine component of the blade pitch link
an anhedral swept tip with the blade cross-section
load 1/rev could be changed to negative pitching only if
constant setting changes mainly illlder the cosine law, the blade tip is swept and bent downward at the same
while drag, under the sine law with the main rotor speed
time.
(i.e. under 1/rev). In addition, the blade tip drag has a
constant portion. These components of aerodynamic The equations obtained contain the centre of tip
forces applied to the blade tip will result in an pressure. As it has been mentioned above, the blade tip
appropriate torsional moment made simple conversions: operates in a 3-d airflow and therefore it is necessary to
a theoretical basis to determine correctly the centre of tip
!:J.M"P = -L!.Y,,pxlip - MtipYttp pressure in compliance with the suggested method of the
Simple conversions can give us equations tip pitch link load estimation. Taking into account the 3-
detennining the value of the constant portion, sine and d airflow effects is quite a challenge, and it requires
cosine components of the blade tip appropriate torsional calculations to be made by using special analytical
moment methods. To make a preliminary estimation of the blade

M1,;po =- ~ s,p[ (&,p) + ~ (V cosa)

2 2
] X (CxoY!ip + c; a.,px,,)

l:llv!,;ph =- ~ v cos a X wr,,pS!ip 2x (C,oy!ip cos X+ c;xo,;pao,;p cos X )

P3-6
 tip efficiency, we shall use calculations of the blade (about 10%).
spanwise distribution of the linear aerodynamic force
Thus, the swept tip with a 7° anhedral does not
with the help of the main rotor blade vortex theory
actually change the constant portion and the torsional
method developed by M.N. Tischenko. Fig. 6 shows
( dtofdr for the Mi-28 main rotor blade in cruise at a speed
moment J/rev, as the presented estimate show. But
anhedral does affect the cosine component of the
of 260 km/h. The blade tip under consideration starts
torsional moment 1/rev.
from 0.93r, as shown in the diagram. Having integrated
thrust distribution over the tip area, an approximate In general, it should be noted that the swept
value of the tip pressure centre is obtained. The point of anhedral blade tip while producing a favourable effect
application of the tip lift Xtip will be located at an on the cosine component of the blade pitch link load
outboard distance 0.41,,. Let us take 0.7 of the tip span 1/rev results in a significant nose down in the constant
as the point to which drag is applied, as the blade tip portion and sine component of the blade pitch link load
drag is mainly affected by friction and velocity of the !/rev. But this unfavourable effect can be eliminated by
free stream. bending the trim tabs running along the trailing edge and
consisting 7% of the chord.
Taking into account the above assumptions, the
location of the centres of aerodynamic forces and area The equations presented in this paper are quite
for a swept anhedral tip can be found from: simple and take into account the basic physical essence
of the swept anhedral tip, but they do not account for
x"' = 0.41"' xtgx"' some effects related to 3-d airflow around the blade tip
ydp =0.7l"'xtgt;dp that produce a favourable influence on the blade pitch
link load of the swept anhedral blade tip. These effects
include the influence of the variable induced wake at the
blade tip for azimuths IJI=O' and IJI= ISO•, airflow around
To make a quantitative estimate of the effect of the end surface of the blade tip at an azimuth ljl=l80,
the swept anhedral tip on blade pitch link loads, let us flexible blade twist, etc. Their consideration requires
consider, as an example, the Mi-28 blade tip tested on numerical methods to be used and labour-consuming
the Mi-28 test bed y the Mil Moscow Helicopter Plant. calculations, which are beyond the scope of this paper.
The geometry of the blade tip is a s follows: R = 8.6 m;
C/1• =6.2; C/·ao =.21; Cxo =0.02; ltip=0.6 rn; Stip=0.32 rn2;
Xtip=O.ll m; Ytip=0.05 m; Xtip=23.8°; ~tip=7°. Assuming 5. FLIGHT TEST RESULTS SHOWING
that p=0.125 kgf/m4; COSC<mn =I; Olrtip =210 m/s, the THE EFFECT PRODUCED BY THE SWEPT TIP
values of the constant portion and components of the OF THE MAIN ROTOR BLADE ON ITS PITCH
blade pitch link load !/rev for a swept, swept anhedral LINK LOADS
and non-anhedral tip can be obtained for speeds of 270
To verifY in practice the results obtained from
kmlh (75 m/s) and 320 kmlh (89 rnls):
the analysis, as well to continue the search for further
Table 2 shows quantitative estimates of the reduction if the blade pitch link loads of the main rotor
effect produced by the blade tip shape on the constant composite blades for the Mi-28 attack helicopter, the
portion and components of the torsional moment 1/rev. flight test programme went on for an experimental blade
For the swept tip without anhedral, the changes in the set that had been and initiated under M.N. Tischenko
components of the torsional moment 1/rev are given in [I]. At first the blades were modified to have untwisted
relation to a rectangular tip without anhedral, while for tip without anhedral. The geometry and size of the tip
the swept anhedral tip, in relation to the swept tip were offered by one of the authors of this paper, Mr. V.
without anhedral. The estimates presented can lead to a Ivchin, proceeding from a theoretical analyse, as well as
conclusion that the swept tip without anhedral for the the analysis of research into different blade tip shapes
main rotor blades affects mainly the constant portion and done by the TsAG!, Mil Moscow Helicopter Plant and
sine component of the torsional moment 1/rev, all other foreign researchers. A swept tapered blade tip ensuring
things being the same. For comparison, the Mi-28 main the best aerodynamic performance in hovering and cruise
rotor blades at speeds of 320 kmlh obtain additional in level flight was chosen. The view of the blade
negative pitching Prno = -1070 N in the constant portion planfonn thus modified is shown in Fig. 7, and its
and Pmls = -780 m. In addition the sine component of the geometric twist, in Fig. 8. This was done to reduce the
torsional moment 1/rev increases by Pmlc = +340 N interaction of the airflow slipping along the blade and

Table 2

Speed 270kmlh 320km/h

Blade tip I'Mmo LlMmlc d.Mmls I'Mmo .6.Mmlc .6.Mmls

+6.9
( Swept, without anhedral -21.6 +5.8 -22.6 -15.8 -13.3

Swept, anhedral -0.9 -24.3 -0.9 -28.+8 -0.8 -0.7

P3-7
 SWEPT TIP CONFIGURATION

-·t·-~'-=====·~~
~~-· -
WITHOUT ANHEDRAL TIP ... ·:~:~... ,

··-----e::::::<:::·==============================3:·~--==-::·33.
WITH ANHEDRAL TIP

·---<==E::=====================~·-~
Fig. 7

the twisted blade producing in the end an increase in the blade pitch link loads by -970 to -730 N depending on
cosine component of the blade pitch link load. the airspeed. These results are in disagreement -with the
All the flight tests of the three blade tip
configurations were conducted on the Mi-24 test bed
equipped with the Mi-28 all composite main rotor ---~---
........,..
blades. As soon as the flight tests of one configuration
were completed, the blades were modified to have the ....• , ....... , ........ ~----- ·! ........ , ....... ; --------~ ---~----···-·1··

nex1 blade tip configuration. The untwisted swept tip

was attached to the main blade by means of a special ... ·····················
adapter having an untwisted shape. The author of this ........ ~- ·······: -------~- ·······: ........ , ·······j··
presentation suggested that new angular adapters should
be manufactured which made it possible to manufacture
an anhedral tip with low labour consumption. In addition _,
to an improved cosine component of the blade pitch link
--<>-Swept, with & wittloutanhedral
loads, the anhedral tip improves the aerodynamic ... ~--

performance of the main rotor which was shown by the -...,.-Rectangular

experimental research done on the Mil Moscow -------~---·····:····

Helicopter Plant whirl tower.

u ~ u u u u u u u u ,.
Blade pitch link loads were measured during the
flight tests of the swept blade tips, and later an analysis
of their harmonic content was made. In this case the Fig. 8
constant portion, 1/rev sine and cosine components as quantitative estimate of the constant portion of the blade
well as the amplitudes of the blade pitch link loads were pitch link loads made above. The Pmo versus V in Fig. 9
compared. The results obtained are given in Fig. 9. shows that the difference in the constant portion of the
As can be seen from the diagrams, the flight tests blade pitch link loads for the rectangular blade tip and
have shown that the swept blade tip without anhedral swept blade tip without anhedral decreases with
resulted in nose down in the constant portion of the airspeed while according to the calculations it should
virtually be the same. This can be attributed to the fact
400 400
350 ················+········· 350

300 ·········-······--~---······ 300

E 250 E 250
c
c
~ 200
~200 ··········-a-swept without
~
::; 150 anhedral
~Rectangular tip ::;E 1so
100 100
~Swept without
50 ···························-~--- anhedral 50

0~--------~------~------~ 0
200 250 300 350 200 250 300 350
Speed, km/h Speed, kmth

Fig. 9

P3-8
 10000
4000
9000 -<>-Rectangular tip ·"?-
( 3000 ···-:- ·······:··················1"···

2000 ·········7·····
8000 -swept tip without
"·:····

1000 -------------·:-------------------:-- 7000 anhedral ··~

-swept tip with

0 6000 anhedral
1:
c_ -1000 ....................
a.E
0 5000
&. -2000 .................t..................). <l
4000
-3000 -<>-Rectangular tip
3000 "'·!··
-4000 -swept tip without ,.
anhedral 2000 .......~ ..
-5000
-swept tip with
1000 ..................... .......................
-6000 anhedral
-7000 0
0 90 180 270 360 0 90 180 270 360
Speed, km/h Speed, km/h

4000 2000

1000
3000 ........... ~·-·· ........... ·---~

-1000
c 2000 c
u
~ ! ~2000 """""t ....................
E E
a. 1000 a. -<>-Rectangular tip
-3000
-swept tip without
-swept tip without -4000 anhedral
0 anhedral -swept tip with
-swept tip with -5000 anhedral
anhedral
-1000 -6000
0 90 180 270 360 0 90 180 270 360
Speed, km/h Speed, kmlh

Fig. 10
the above equations do not take into account the blade above and attributed to low stiffness of composite
torsional flexure. According to the analytical and blades, which had been ignored in the munerical
experimental data, composite blades have half the estimate. The second one lies in the fact that the blade
torsional flexure as compared to that of the Mi-24 cross-section airfoils have trailing edge tabs whose
duralumininm blade. At the same time, a higher value of consist equals to 7% of the blade chord numing along
the negative pitching moment produced by the swept the whole span of the blade. These tabs are intended for
blade tip will lead to a substantial change in the blade adjusting the level of the blade pitch link load by their
geometric twist, to the redistribution of the aerodynamic bending that is why they are made of metal. These tabs
loads applied along the blade and, thus, to a change in were recent by using special templates in the process of
the blade pitch link load. This is substantiated by the modification of the blade to accommodate different tips
comparison of the constant portion of the blade pitch and during the flight tests. However, due to insufficient
link load versus airspeed for the swept tip with and supervision and poor accuracy of the templates, no
without anhedral. The diagrams of Pmo versus V in Fig. 9 reliable infonnation on actual table setting is available.
for these blade tips are ahnost equidistant in airspeed, as Yet another reason lies in the fact the swept blade tip
the distribution of the twisting moment applied along the without anhedral was also fitted with a tab but no fixture
blade, and, thus, elastic twist is the same for them. to use for adjusting the tab and checking its setting was
made. Therefore, it may happen that a change in the tip
The value of the sine component of the blade
tab setting can lead to a change in the sine component of
pitch link load 1/rev after fitting the blade with a swept
the blade pitch link load 1/rev.
( tip without anhedral has change to nose up 6Pmls =+970
N at a speed of 320 kmlh. The numeral results obtained These conclusions are substantiated by the flight test
above give a value equal to -770 N. This difference can results. The twisting moment measured at the cross-
be caused by a few reasons. The first one was given section located at 0.9R of the reference blade with a

P3-9
rectangular tip was 280 Nm at a speed of 320 km/h. The swept tip without anhedral was 300 Nm [6], so that the
same moment for the blade with a swept tip without difference was - 70 Nm.
anhedral was 30 Nm [6]. It means that the swept tip
Thus, the effect of a reduced Mm!c obtained in flight
without anhedral resulted, according to the analytical
tests is determined not by a swept tip, but by the fact
predictions, in a change in the sine component of the
that the tip has no geometric twist.
blade pitch load !/rev by- 250 Nm.
Fig. 10 shows changes in 1/rev components of the
As for the cosine component of the blade pitch load
torsional moment obtained by measwing it at 0. 9R in
1/rev, the flight test results were opposite to those
flight tests for a rectangular tip and a swept tip \vithout
obtained analytically. The flight test results showed a
anhedral. The relations presented substantiate well the
decreases in the cosine component of the blade pitch link
effects produced by the swept tip on the blade pitch link
load !/rev by -290 N, while the analysis had shown its
loads obtained analytically.
increase by +340 N. The reason for this phenomenon is
the effect of the interaction between the slipstream Let us consider the effect produced by a swept
flowing around the blade tip and a change in the blade anhedral tip. As can be seen from the diagram in Fig. 9,
geometric twist. M.N. Tischenko predicted the presence the flight tests have shown an occurrence of an
of this effect in the process of flight tests of the Mi-28 additional negative pitching in the constant portion of
new rotor system on the flying test bed. the blade pitch link load by -1220 Nm. Qualitative
estimates show that this should not take place. An
Reference [7] shows that, for a twisted wing, the
analysis of potential causes has shown that when the
equivalence hypothesis is not valid, and the airfoil
blade was modified, the trailing edge tab settings were
performance obtained in 2-D airflow are not acceptable
adjusted in the shop by using protractors. Those settings
in this case. Therefore in calculation of the helicopter
should have been in compliance with the settings
rotor twisted blade it in necessary to consider the effect
established for the swept tip without anhedral, but our
produced by the slipstream on the airfoil aerodynamic
attempts to obtain reliable information on actual settings
performance in case the blade has a geometric twist.
on the modified blade failed. The following has actually
Equations required to obtain corrections for the
been found out. The tab setting on the blades with swept
aerodynamic performance of airfoils obtained in 2-D
anhedral tips that were intended for adjusting the rotor
airflow when they operate on a twisted blade under the
blade track and implemented on three sections on each
slipstream effect were derived in Reference [8]
blade, were reduced by 1° in average. The result was an
proceeding from the lifting line theory. The following
additional negative pitching moment in the constant
equation is suggested for calculating the airfoil value of
portion of the blade pitch link load by -195 N. Bending
illl>w in Reference [8]:
the tip downward leads to an additional negative
t=~ ~ 2~\V [co.7- .~:,)m: -0.0625c;]bR, v cos a." tgx ~ pitching moment in P mo produced by resistance on the
ann Ytip by another -50 N. An increase in the tip moment
or 6r,;p
of inertia produces an approximately the same value of
2°\V
8r
[co.7 -.~: )m~ -0.0625c"] b, v cos a."
l- YR
the negative pitching moment due to a longer distance
6."lt;;p between the blade rotation axis and the tip centre of
mass. Thus, these data can explain an occurrence of a
Taking into account that for the considered blade 25% additional negative pitching moment in the constant
m,'~-1.6, x,lb~0.21, b!R~0.07, d<p/dr=-0.26, the portion of the blade pitch link load. The comparison of
following equation for the cosine component of the tip the relations Pmo and Pm!s presented in Fig. 9 for tips
torsional moment can be obtained: with and without anhedral shows that this nature of the
curves is inherent in the change of the value mzo of the
blade sections. Therefore it is most probable that the
blades with and without anhedral tips had different tab
settings due to which an additional negative pitching
For a speed of 320 kmlh, the value of the cosine moment was produced for the blades having anhedral
component of the torsional moment 1/rev obtained by tips.
using the equation will be + 110 Nm. For a swept tip
without anhedral and geometric twist the total value of The tip anhedral has affected most effectively the
the cosine component change is determined by the cosine component of the blade pitch link load !/rev at
following value: airspeeds higher than 270 km/h. As can be seen from the
diagrams in Fig. 9, at a speed of 320 krnlh the Mm!c has
M.,, ~M.,., +M.,,,. ~+69-136~--07[Nm] reduced by 20%, and at a speed of 340 kmlh, by 35%,
although these values are lower that those obtained from
That is, the value of the cosine component of the the qualitative estimates. Now let us consider the effect
torsional moment 1/rev for a swept tip without anhedral of the blade tip shape on the amplitude of the variable
when compared to that of the initial blade \Vith a portion of the blade pitch link load. As can be seen from
rectangular tip is close to the total value obtained from the diagrams in Fig. 9, at airspeed up to 280 km/h the
flight test results (Fig. 9) which equals -70 Nm. rectangular tip and the swept tip without anhedral
This is also substantiated by flight test results. The virtually have the same dP m, but at airspeeds exceeding
measured torsional moment Mm!c at the blade cross- 320 kmlh this value for the swept tip without anhedral is
section at 0.9R of the initial blade with the rectangular 970 Nm lower. At the same time the tip anhedral
tip was 370 Nm at a speed of 320 Ian/h. The value of the reduces the amplitude of the amplitude of the variable

P3-10
portion of the blade pitch link load at airspeeds lower versions of the main rotor blades. The results obtained
than 280 km/h and increases it at higher speeds. For were reduced to standard conditions at sea level by using
instance, the dPm value reduced by 20% at a cruise speed the procedure developed by Mr. Ivchin, co-author of the
of 260 km/s and increased by I 0% at a speed of 340 paper. Fig. II presents maximum relative TOW in
km/h. As cruise makes up about 50% of the helicopter standard conditions at the engine takeoff power rating
life cycle, it is quite clear that a reduction in the variable versus hovering height above the ground. The data is
portion of the blade pitch link load can result in quite a given for the Mi-28 main rotor with the swept anhedral
substantial exiension of the service life of the helicopter tip (for three versions of the trailing edge tab setting, the
rotor and flight control system components. Mi-28 main rotor with the rectangular tip [I OJ and two
versions of the Mi-24 production main rotor [11]. The

Speed 250 krnlh Speed 320 km/h

10000 ,----~----~---~-----, 10000,-----,-----,------,-----,

5000 -······ .....................,.......... . .................... ,.

c c

..
E -sooo
............ .;- ... -<>-Rectangular tip
£ -5000

-10000 .....
-10000
~Swept tip without i -o-sweepttip
-15000 ·······:·· anhedral -15000 ..................... ~ ..................... , without anhedra\ ........ .
-:J- Swept tip with 1""'&-Sweept tip witfl
anhedral anhedral
-~OOOl__ _ _l__ _~~==========l__j -20000 L----'-----'~~==~~d.._.J
0 90 180 270 360 90 180 270 360
Azimuth, degree Azimuth, degree

Fig. 11

To illustrate the change in the blade pitch link relative TOW was determined in relation to the
load in azimuth, Fig. 11 shows dependencies Pm(\f!) for maximum TOW of the Mi-24 equipped with a
the tips being considered for speeds of 250 km/h and production main rotor obtained in the above conditions
320 kmlh obtained from the flight tests. As can be seen while hovering out of ground effect. The diagram shows
from the diagram, there exists quite a considerable that the thrust characteristics of the main rotor having a
negative pitching moment within the azimuth range from rectangular tip were virtually those of the Mi-24
90° to 130°. The Pm value becomes as high as 1270 Nm production main rotor. The swept anhedral blade tip in
for the blade with the swept anhedral tip at speeds of combination with the most favourable trailing edge tab
320 kmlh. At the same time, the amplitude of the blade setting leading to an increase of the helicopter TOW by
pitch link load for the swept anhedral blade tip does not 4% and 5% when hovering out and in ground effect
practically differ from that for blade s without the respectively. However, it necessary to take into accotmt
anhedral tip within the azimuth range from I 80° to 360°. the fact, that for this version of the main rotor the blade
It means that the mzo values of the cross-section airfoil of pitch links load components would be much higher as it
this blade are more negative than for the previous tip has been discussed above which is presented in Fig. 8.
version. In accordance with the analytical data [9], as Therefore, to reduce the blade pitch link load the tabs
well as the flight test results (3] a change in the mzo running along the tailing edge of the blade with the
value of the airfoil caused by bending do\Vllward trailing swept anhedral tips were bent downward. The result was
edge and trim tabs results in a change only in the value a reduction of the trust characteristics of this main rotor
of the constant and sine components of the blade pitch to 2.5o/o-4%, as shown in Fig. 12.
link load 1/rev. Therefore, using the expressions from
Ref. 9 to determine the mzo versus the trailing edge tab
setting, we shall obtain that to half the Mmls value at a
speed of320 kmlh, it is necessary to bend downward all
the tabs running along the blade trailing edge by I o. At
the same time this will lead to an increase in the
constant portion of the blade pitch link load (+ 1700 N). '
~ 1.1

The result would be some degradation of the main rotor l

;
aerodynamic performance, but it would be negligible. ~
p 1.0
6. EFFECT OF A SWEPT ANHEDRAL TIP
ON HELICOPTER PERFORMANCE
During flight tests of different blade tips, the 0.9
effect of the swept anhedral tip on the helicopter 0 5 10 15 20
performance was studied. To do that, comparative tests Wheel height, m
were conducted in hover and in level flight for several Fig. 12

P3-11
 1.10 to a change in the setting of the tabs running
--o-Mi-28 anhedral lip along the blade trailing edge.
--o---Mi-28 rectangular tip
1.05 5. Blades having a swept anhedral tip allow the
---ik-Mi-24 blade thrust characteristics be increased by 4-
5% in hover or the power required to be
reduced by 8-9%.
6. Blades having a swept anhedral tip allow the
powerplant power required be reduced by 10-
14% in cruise flight.

0 50 100 150 200 250 300 350

Speed, kmlh 8. REFERENCES

Fig.13
1. V. A Ivchin, M. N. Tischenko, V.A. Animitsa,
V.A. Golovkin, Analysis of Model and full-
To assess the effect of the swept anhedral tip on scale Investigation Results of Helicopter Blade
the power required for level flight, comparative tests of Aerodynamic Configuration Effect on Pitch
the Mi-24 main rotor were conducted. Fig. 13 presents Link Load, 22"' European Rotorcraft F arum,
the test results in the dimensionless form. The presented London, England, 1996.
relations were obtained by establishing the relation
between the Powerplant power required for the 2. M.r. Po)[():(eCTBeHcKHii, Km.mneKc pa6or,
helicopter equipped with different sets of main rotor HanpaBJieHHbiX Ha IIOBI:.III.IeHHe JienibiX
blades and that for the helicopter having an Mi-24 xapa:rcrepHCTifK neprorreros MH, Tpy.zu,r
production main rotor. As can be seen from the om.rmo- KOHCTp)IKTOpcKoro 6ropo :m.t: M.Jl
<liagrams, the power required for the Mi-28 helicopter Mnm! 1997 r.
equipped \Vi.th a main rotor with a rectangular tip has 3. AKT N• 56-86 rro pe3yJThTaTilM JaMepa
increased by 7.5% at higher speeds and decreased by Harp)'30K Ha OllblTI!bfX )'lllicjllllU!POBaHHhfX
2.5% at lower speeds when compared with the norracr!IX Hec)'lllero Bmrra qepT.280-2903-
helicopter equipped with an Mi-24 production main 00CE yCTaHoBneHHbrx IDI(. "245",
rotor. In hover, there is almost no gain, which is in a TeXHH'!eclGf!l Alcr 56-89, MocKoBclGf!l
good agreement with the diagrams in Fig. 11. The swept BepTonenn,Ill JaBal( HM. M.JI. Mnm!, 1987 r.
anhedral tip makes it possible for the power required in
hover and in cruise to be reduced by 9.5% and ahnost 4. Pe3yJibTaThi nenn:.rx HCilhrra:HlfH o6oeKTa
14% respectively. It is of interest to note, that the power 8Q:MJ' CO CTeKJIOIVIaCTHKOBI:.IMH JIOIIaCTIIMH:
required for all the blade sets under consideration does Hecymero nmrra. TeXHJ1l.IecKHii OTtleT 54-87,
not actually change at the best economy cruising power. KBIIO, 1987 r.

Proceeding form the data presented in the paper, 5. Pe3yJII:.ran.r nenn:.rx pa6oT Ha o6oeKTe 80MT
the following conclusions can be made. N 93158 !10 OIIpel(eJieHlilO Harpy30K B
ynpaanemm H Kp)'T51IUHX MOMeiiTOB OilhiTHbiX
creKJioiUiaCTIIKOBI:.IX nonacre:H: Hecymero
7. CONCLUSIONS Bmrra. TeXHH'!eclGf!l OT'ICT 32-88, KBIIO,
1988 r.
1. The swept anhedral tip allows the cosine 6. JieTHI:.IX
component of the blade pitch link load to be JKcnepHMeHTa.JlbHI:.IX nonacre:H: co
reduced due to the fact that it has no geometric Ha JieTa.IOIT.(eif
CTpeJIOBH,lJ;HI:.IMH 38KOHJ.(OBKaMH
twist. However, quite considerable negative na6oparopirn ID.U. 242, TeXHl£liecl<l1if orqer 5-
moments occur in the constant and sine 89. MocKosci<HH sepro11enn:.lli 3aao.n :m.t:. M.Jl
components of the blade pitch link load 1/rev. Mnm!, 1989 r.
2. The swept anhedral tip allows the cosine 7. B.A. HB'!IfH, B.B. IIplKe6em,clGlll,
component of the blade pitch link load to be Hccne.nosaHHe ropo,ll;KHa.MWieCKHX
reduced by 20-35% at speeds excee<ling 300 xapaKTepHCTiiK rrpocjJI!lli! I<p)"!eHoll norraCTII
km/h. npH Ha.Jll£liHH CKOJibiKeHIDJ., Tpy.zu,r Hayqm:.rx
3. The swept anhedral tip allows the amplitude of lf!'eHH.if nocnsm.J;eHHI:.rx rrawrrn aKa,neMHKa
the variable portion of the blade pitch link load B.H. :EOpi:.eBa, Teopel1iliecme OCHOB~I
to be reduced by 20% at cruise speed of 260 nepTOJieTOCT'poe:mrn:, MocKBa, HHET AH
CCCP, 1987 r.
km/h.

4. An additional negative pitching moment 8. B.A. Hnl.JHH, BJIIDIHHe CKOJibiKeHIDI Ha mz H Cy

occurring in the constant and sine components rrpo¢HJIH JaKpyqemmM nonacrn, TeXffifl-IeciG:Iii:
of the blade pitch link load !/rev for blades oT'!er MB3 N•1390, 1987 r.
having a swept anhedral tip can be attributed 9. B.A. HBlfifH, B.JllliiHMe 3aKpi:.UIKOB Ha

P3-12
 ropo.u;mtaMI{I.lecKMe xapaKTepH:C'I'HKH rrpo¢HJUI
orraCTH neproJiera, TeXHJ{qeci<Illi orqer MB3
N"212, 1987 r.

10 Onrer no pe3y.JlbTaTaM cpaBHJITeJlliHE:.IX

JieTHbiX H:CIJJ:>ITa.HJill: .H.B)IX KOMIIJieKTOB
JiorracreH HecYIIlero Bmrra m~. 280 Ha
JieraromeH Jia6opaTopmr, ON:er N!~ 104-86
MB3 HM. M.Jl. Mluu!, 1987 r.

11 JleTHbxe H:CIThiTa.HIDI no cprume:mno TIITOBbiX

xaparcrepl!CTI!K 1!3)(. 245 N"21l 08 Ha peOK!iMe
BHCeHIDI C cepiDfHhlMR H OITbllHbiMH
280 (KOMIUJeiiT N2 2901 ,
rronaCTl!MH ID)(.
npoljliDib NACA" CE c HaiU!biBOM), Qrqer N2
3-86 MB3 l!M. M.Jl. MIDrn, 1986 r.

12 A.3pO)J.HHaMlf1JeCKMe HCITblTaHIDI MO~eJlliHOrO

BHHTa C Blili3 KOillleBoif l.JaCTII
O'ITif6oM
JionacreH, T eXffi{qecKHI1 orqer MB3 RM. M.n.
Mluu!, 1978 r.

P3-13