N4SA CONTRACTOR REPORT 177340

I_.ql

STUDY FOR PREDICTION OF ROTOR/

WAKE/FUSELAGE INTERFERENCE

PART I: TECHNICAL REPORT

(ZASA-CR-177340-VoI-I) S2UD! EOR PREDICTION

OF ROTOa/WAKE/¥USEL&GE _NTZ_¥RR_NCE, P£_T

Final Re_ort, I dun. 198u - I Nov. IS83

(Analytical _etbods, Inc., F_dnond, Wash.)

92 p HC A05/_F A01 CSCL 01& G3/u|

D. R. Clark

B. Maskew

N85-223_6

,Q

"qi

i

CONTRACT NAS2-10620

March 1985
NASA CONTRACTOR REPORT 177340

STUDY FOR PREDICTION OF ROTOR/WAKE/

FUSELAGE INTERFERENCE

PART I: TECHNICAL REPORT

D. R. Clark

B. Maskew

Analytical Hethods, Inc.

2047 - 152nd Avenue, N.E.

Redmond, WA 98052

Prepared for

Ames Research Center

under Contract NAS2-10620

National Aeronautics and

Space Administration

Ames R_rch Center

Moffett Field, Cahforn_a 94035
ABSTRACT

A method has been developed which allow_ the fully coupled

calculation of fuselage and rotor airloads for typical helicopter

configurations in forward flight. To do this, &n iterative

solution is carried out based on a conventional panel representation of the fuselage and a blade element represent&tion of the

rotor where fuselage and rotor singularity strengths are

determined simultaneously at each step and the rotor wake is

allowed to relax (deform) in response to changes in rotor wake

loading and fuselage presence. On completion of the iteration,

rotor loading and inflow, fuselage singularity strength (and,

hence, pressure and velocity distributions) and rotor wake are

all consistent.

The results of a fully coupled calculation of the flow

around representativ_ helicopter configurations are presented.

The effect of fuselage components on the rotor flow field and the

overall wake structure is detailed and the aerodynamic

interference between the different parts of the aircraft is

discussed. In particular, the flow field developed by the rotor

head is followed and the effect of a rotor head cap and pylon

modifications in redirecting the rotor head flow are illustr_ted.

Good correlation between measured and calculated fuselage

airloads in low-speed flight is achieved and correspondence with

observed flow field behavior is demonstrated.

•

i
TABLE OF CONTENTS

section Eaa_ Sn.

ABSTRACT ........................ i

LIST OF FIGURES .................... iii

LIST OF SYMBOLS .................... vi

1.0 INTRODUCTION ................... 1

2.0 ANALYSIS

: Background to the Analysis ......... . 6

2_.3WakeCuttingProcedure . .........

13

2.4 Rotor Blade Aerodynamics ........... 18

2.5 Rotor Disc Panel Model ............. 21

3.0 THE COMPUTER PROGRAM ............... 24

4.0 DISCUSSION OF RESULTS

4.1 Application of the Analysis to a Simple Body/

Rotor Example . • 25 42 BasicBody/Rotor _e_uit__or_ _e_r_s_n_a_i_e

Configuration - • 25

Rotor/Fuselage Flow Field ...... 39

4.4 The Potential to Study Full Configurations . . 48

5.0 CONCLUSIONS AND RECOMMENDATIONS .......... 50

6 •0 REFERENCES .................... 51

APPENDIX A: Wake Shape and Model Surface Pressure

Correlation Results

Ii

I
LIST OF FIGURES

E_No.

1

2

2iLtl

Typical Output from Vortex Filament Wake Model

Typical panel Representation for Helicopter

Analytical Model ...............

10

11

12

13(a)

13(b)

14(a)

14(b)

14(c)

15

16

3 BodyRotorWake Influence Block Matrix ....

4 Panel Model with Vortex-Lattice Singularities--

Implicit Velocity Component .........

5(a) Comparison of Off-Body Streamlines Calculated

using VSAERO with Exact Solution Result . . .

(b) With Pressure Distribution ..........

6 General Program Schematic ..........

7 Schematic of Wake Cutting Model .......

8(a) Wake Cutting Study--Side View ........

8(b) Wake Cutting Study--Top View ........

9 Wake Cutting Study--Spanwise Doublet Distribution (0.25 Chord) ..............

Wake Cutting Study--Spanwise Velocity (0.25

Chord) ....................

Body/Rotor Interference Calculation Schematic

Blade Model/Panel Model/Wake Schematic ....

Body Rotor Interference Study--Starting

Condition ......... • .........

Body Rotor Interference Study--Starting

Condition ..................

Body Interference Study, Relaxed Wake ....

Body Interference Study, Relaxed Wake ....

Body Interference Study, Relaxed Wake ....

Isolated Rotor, Relaxed Wake .........

Outline of Freeman-Mineck Model used in Analysis (Scaled to 20 ft. Rotor Radius) .....

iii

7

7

8

i0

I0

ii

14

15

15

16

17

22

22

26

26

27

27

28

28

30

I

Y'_B
LIST OF FIGURES (Continued)

KQ. 2LLLt E gl No.

17 Basic Panel Model for Body/Rotor Studies . . 30

18(a) Aft Fuselage--Typical Correlation with Test Data;

0.05 Advance Ratio, Run/Point 22/139, Station

17.6 (0.88) ................. 31

18(b) 0.15 Advance Ratio, Run/Point 25/148, Station

17.6 (0.88) ................. 31

19(a) Front Fuselage--Typical Correlation with Test

Data--0.05 Advance Ratio, Run/Point 22/139,

Station 1.80 (0.09) ............. 32

19(b) 0.15 Advance Ratio, Run/Point 25/148, Station

1.00 (0.05) ................. 32

20 Calculated Distorted Wake--Low Advance Ratio . 33

21 Cross Section through Typical Low Advance

Ratio Wake (Station 30.0) .......... 33

22 Typical Effect of Rotor Wake/Pylon Intersection--Body and Rotor Alone

(a) Calculated Pylon Doublet Distribution--Low

Advance Ratio (Waterline 3.00) ........ 35

23

(b) Calculated Pylon Horizontal Velocity Ratio,

Vx/V ® --Low Advance Ratio (Waterline 3.00_..

Calculated Vertical Velocity Ratio Vz/V® ;

Waterline 0.0_ Body and Rotor Alone

35

(a) Advance Ratio 0.05 .............. 36

(b) Advance Ratio 0.15 .............. 36

24 Calculated Effect of Body on Rotor Angle of

Attack--0.05 Advance Ratio .......... 37

25 Calculated Effect of Body on Rotor Angle of

Attack--0.30 Advance Ratio .......... 37

26 Calculated Contours of Upwash Velocity in Rotor

Plane; Body and Rotor Alone_ 0.30 Advance Ratio 38

27 Body-Rotor Panel Model with Tail Assembly . . 38

_V
LIST OF FIGURES (Concluded)

28

29

30

31

32

33

34

35(a)

35(b)

36(a)

36(b)

37(a)

37(b)

38

39

Pressure Distribution on Lifting-Surface Model

Effect of Tail Surfaces on Calculated Blade

Angle of Attack; 0.15 Advance Ratio .....

Basic Body Model with Nacelles and Tail Surfaces Added .................

Effect of Nacelles on Calculated Blade Angle

of Attack--0.15 Advance Ratio ........

Calculated Rotor Head Model Wake Development--

0.15 Advance Ratio ..............

Rotor Head Wake Cross Sections--0.15 Advance

Ratio .............. ......

Effect of Rotor Head and Wake on Calculated

Blade Angle of Attack--0.15 Advance Ratio . .

Calculated Rotor Head Cap Wake Development--

0.15 Advance Ratio ..............

Calculated Rotor Head Wake Development with

Modified Pylon--0.15 Advance Ratio ......

Calculated Wake Cross Section with Rotor

Head Cap Wake--0.15 Advance Ratio ......

Calculated Wake Cross Section with Modified

Pylon--0.15 Advance Ratio ..........

Effect of Rotor Head Cap on Calculated Blade

Angle of Attack--0.15 Advance Ratio .....

Effect of Pylon Modification on Calculated

Blade Angle of Attack--0.15 Advance Ratio . .

Mock-up of Full Configuration with Rotor Head

Cap (Actual Run Exceeds Present Program Panel

Capacity) ..................

Effect of Tail Rotor on Vertical Stabiliser

Loads at WL 5.75 (Main Rotor Present--0.15

Advance Ratio) ................

40

4O

41

41

42

42

44

44

45

45

46

47

47

49

49

v
,

LIST OF SYMBOLS

Note: In this report two different analyses which have

(separately) well defined sets of conventional symbols are

brought together. The conventions have been retained and, as a

result, some symbols have double definiticn. Their meaning,

however, remains clear when they are considered in the context of

their applications. The symbols involved are _ and u. In the

panel model convention these represent the doublet and source

singularity strengths, while in the rotor model convention, they

represent rotor solidity and advance ratio. The uses are sufficiently separate that confusion should not occur.

dA

b

C

"C

Z_Cp

CT

e

L

_L

_L

M

NC

9

r

R

_L

T

U

V

V !

X,Y,Z

Rotor panel model panel area

Number of blades

Blade chord

Mean blade chord

Rotor panel model differential pressure coefficient

CpUPPE R CpLOWE R ) _p'_=z AA

T

Rotor thrust coefficient =

Blade flapping hinge offset

Rotor lift in wind axis system

Blade model segment load

L b

Panel model time-averaged load = _ ._r. N--C

Mach number

Number of azimuthal increments into which disc is broken

for panel model

Rotor propulsive force in wind axis system

Radial location on blade

Blade tip radius

Blade segment length

Rotor thrust

Velocity components in blade section axis system

Velocity components in wind (global) axis system

Rotor local resultant velocity

Global coordinate system

f

8

¥

Roto_ control axis orientation

Body pitch attitude

Blade flapping angle

Blade flapping rate (_81_t)

Blade local yaw angle

vi

i
LIST OF SYMBOLS (Concluded)

o

O

D

U

U

Blade azimuth angle

Blade rotation rate (_/_t)

Rotor panel model, doubled strength

" _ r 2(V'/V®) _Cp dr

o b_

Rotor solidity _ _--_

Air density

Panel model source strength

Rotor model advance ratio, V /_R

4!

x,y,z

c

I,P,R

Components of local velocity in global coordinate system

Climb

Components of local velocity in blade section system;

I - inplane, P ffiperpendicular, R = radial

Free stream conditions. V® is defined along global

system x-axis

vii
i

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I

t

i !

E

m_

--!

.I

1.0 INTRODUCTION

AS helicopter designers work towards the development of a

vehicle which can compete with fixed-wing aircraft, if not in

terms of speed, at least in terms of passenger acceptance in the

areas of ride quality, vibration, and noise, they are being

forced more and more to acknowledge the complex interaction that

takes place between the rotor and the airframe. Body/rotor

interference manifests itself throughout the operational range of

the helicopter; it is as significant at low speed in the form of

rotor induced fuselage downloads as it is at high speed where the

most important effects are the irregularities in rotor loads

induced by the passage of the blades through the fuselage flow

field. Added to these effects is the controlling role of the

main rotor wake in the handling qualities, passing as it does

over and around the horizontal and vertical tail surfaces and the

tail rotor as speed and flight conditions change. The ter_

"interactional aerodynamics", coined by Sheriden and Smith, _

aptl_ describes the very involved process which controls helicopter loads, dynamics, handling qualities and performance.

_hrou%hout the first three decades of the helicopters

existence as a practical machine, the profound effect that the

pre_ence of the fuselage can have on rotor behavior was hardly

acknowledged. This was largely due to the fact that there was

not a strong, driving requirement to understand the interaction

and rotors were designed, analysed and tested in isolation. The

fact that when installed they behaved differently, trimmed at

different cyclic pitch settings and had considerably different

aeroelastic response and dynamic characteristics was correctly

attrlbuted, in most cases, to the presence of the fuselage. However, no serious attempt was made to understand the phenomenon

and since the modest performance and dynamics goals of the period

were being met, there was no incentive to refine the design

methods.

The situation changed dramatically in the early seventies as

a result of the competition to provide the U.S. Army with new

utility transport and attack h_licopters. A prime requirement in

both programs was that the designs must all be ai_transportable

within certain very clearly defined limits and this resulted in

designs in which the rotor was placed, initially, very clo_e to

the fuselage in an attempt to reduce the overall height of the

vehicle. All of the vehicles tested in this configu=ation exhibiued undesirable dynamic characteristics where were attributed

to fuselage induced rotor inflow variations.

The phenomenon was first explored analytically by Landgzebe

et al.,2 in a paper which examined the mathematical tools available for the design of the new generation of rotor craft. In

this work the flow field induced by the fuselage in the region of

the rotor was calculated by an early potential flow configuration

modelling program. For this early study there was no direct

coupling of rotor and fuselage e£fects, and the velocities

1
calculated in the rotor plane by the fuselage analysl e

simply fed as inflow into the rotor analysis. Because .e

size of the individual programs and the limitations of the computing facilities available at the time, no coupling of the rotor

on fuselage effects was attempted beyond very simple sourc_ plane

or vortex tube rotor models. Despite these limitations the

analysis was used with some success to explore alternative rotor

locations. The study showed how, when operating close to the

fuselage, the rotor is exposed to an azimuthally varying inflow,

predominantly up over the nose and down aft of the shaft but

containing higher harmonics, which significantly degrade the

aircraft vibration environment. Also, it was realized that the

upwash over the front fuselage was severe enough, causing very

large increases in angle of attack as the blade passed through

the forward portion, to precipitate stall as far out as mid-span.

Inclusion of the fuselage induced flow field in the dynamic

analysis dramatically improved correlation with measured data.

Fuselage/rotor interference has been explored from both

experimental and analytical sides. Noteworthy from the experimental point of view has been the work of the group at NASA

Langley. Following the early work of Wilson and Mineck,3 direrted mainly at handling qualities and low speed fuselage loads,

the work of Freeman with Mineck,4 and later with Wilson5 explored

systematically the influence of body shape and relative

rotor/body position on fuselage and rotor airloads. They showed

how with increased fuselage width and reduced body/rotor spacing

the performance of both systems is degraded. The work of

:_heriden and Smith,1 concentrating on a particular configuration

also explored the effects of body/rotor placement. More recently, Betzina and Shinoda6 working with a scale model of a wind

tunnel test module (from the NASA Ames 40 x 80 wind tunnel) have

examined coupled rotor/body integrated performance_ Formerly a

rarity, test of rotor/fuselage combinations are now standard

rocedure as designers try to define more closely the differences

etween analysis and actuality and between model and full-scale

test results. Reference 6 presents a fairly typical outline of

the gaps that still exist in the understanding of the coupled

flow field.

In parallel with the expansion of the experimental data

base, work has continued on the development of analytical tools

to explore rotor/body phenomena. Several different approaches to

modelling the flow field have been employed ranging from involved

combinations of vortex filament wake models and full fuselage

panel models to simple Stacked vortex ring arrays. In all cases,

however, inclusion of the effect of the presence of the fuselage

in the description of the rotor inflow improves the prediction of

unsteady effects.

At a rotor wake workshop held by the U.S. Army Research

Office (AROD), Smith8 presented a method where the fuselage was

represented by a single source element in a uniform flow and the

rotor by a series of constant strength vortex rings, displaced
L _ •

upwards as they passed through the sphere of influence of the

body source. Despite the relative crudity of the model, the

predicted rotor loading shape was good. The same basic model was

used by Young. 9 Wi_h a more refined fuselage model and considering both vertical and horizontal displacement of vortex

rings, he was able to snow some improvements in correlation with

test data. Incorporation of cyclically varying circulation

around the rings further enhanced the correlation.

A more detailed analysis which more accurately represents

the fuselage with a complex panel model and the rotor with a wake

filament model was used by Landgrebe et al.2 The method was used

in the work discussed in References 7 and i0. Here, the flow

around the panel model is first calculated and the velocity field

in the plane of the rotor determined. This is used as input to a

rotor performance calculation which can include as much detail

as desired, up to and including a full filament wake distortion

calculation. It is not clear from the published work whether any

higher-order coupling is involved. Certainly, in the schematics

presented in Reference 7, the arrows connecting fuselage and

rotor aerodynamics modules only go one way, from fuselage to

rotor, implying no higher-order coupling than a simple rotor

onset flow modification. However, comparison between results

predicted using the method and test data show good agreement.

A similar approach is taken by Huber and Polz. II Using a

detailed panel model they calculate the flow in the plane of the

rotor for input to a rotor analysis. Again, no higher-order

coupling is used. This is reflected in the fuselage induced

upwash profiles presented in Reference ii, which are symmetric

about the center plane. If coupling had been present, the upwash

contours would have been asymmetric, reflecting the differences

in loading between the advancing and retreating inboard blade

sections. _spite this, they show very graphically the large

impact of the fuselage on the rotor loads, especially in the

region of the forth and fifth harmonic.

Huber and Polzll also present results from earlier studies

where, for the first time, attempts have been made to calculate

the effect of regions of separated flow on downstream components. The earlier work 12 presents the method in detail.

Following the by now conventional technique (See References 13

and 14), the authors, using a panel method and streamline procedures, calculate the extent oF. regions of separated flow; then,

with a novel volumetric vorticity singularity model, are able to

determine the velocity field inside the separation zone

downstream. They do not, however, present any calculation on the

effect of these flows and, in fact, their sample cases are unnaturally truncated. This is done since the type of singularity

model used in the analysis cannot handle a direct vortex/surface

intersection.
Another theoretical approach to the body/rotor problem is

that taken by Freeman.15 Again, using a basic panel model for

the fuselage and a vortex tube model of the rotor wake, following

Heyson,16 he is able to show quite impressive correlation between

the results of his analysis and body/rotor test data. However,

there was no coupling present between rotor and fuselage flows,

the rotor model being simply used to perturb the fuselage model

onset flow, and no fuselage induced wake distortion was introduced into the rotor calculation.

p!

None of the rotor/fuselage analyses examined couples fully

the flow fields of the two components to the extent that terms or

higher than first order; that is, fuselage on rotor2,7,9 or rotor

on f_selagel5 are included, and none have been able to calculate

the ue_ormation o_ the _otor wake in the presence of the fuselage

or handle the direct vortex wake/fuselage cutting situations.

The reasons for this were outlined by the present author at the

U.S. Army Conference on Wake Modelling in 1979,17 and in detail

in Reference 18. The lack of an adequate coupling analysis

results partly _rom the limited capacity of the computing

machines then in use (the fuselage panel codes and vortex wake

codes on their own consuming most of the machine capacity--

precluding direculy coupled calculation), but mostly from the

inability of the potentlal flow models to handle the close approach of strong singularities in the external flow and direct

vortex/surface encounters.

The goal of the present study was a full description of the

highly interactive helicopter flow field including powerplant

exhaust, horizontal and vertical stabilisers, tail or other

auxiliary rotors and separated wakes from upstream components

such as bluff rotor heads. This was made possible by the coupling of a rotor module to an advanced potential flow modelling

code. The program, designated VSAERO (Vortex Separation

AEROdynamics),19,20 uses combinations of doublet and source

singularities, t)gethec with changes in the way in which the

boundary" conditions are applied, to solve for uhe local, scalar

doublet strength. This is differentiated to define the local

velocity field. Techniques have been developed to handle not

only close vortex approach but also direct wake cutting. The

rotor wake (wakes if more than one rotor is present) is represented by time-averaged vortex sheaths which are allowed to deform

in the presence of the £uselage flow field. The rotor/fuselage

coupling is made throu@h a blade element model for the rotor

supplied with inflows zrom the fuselage/wake calculation and

feeding back circulation strengths to the wake.

The model has been used with some success to look at basic

body/rotor performance over an advance ratio range from 0.05 to

0.3. Correlation of fuselage loads inside the wake interference

, zone is generally good. The test data used for the correl_ticn

i was that of Freeman and Mineck.4 The same basic configuration

J

was used as a starting point for a study of the effects of adding

_i configuration components and a full buildup was carried out. The

4
effects of adding horizontal and vertical stabiliser, tail rotor,

engine nacelles and exhaust plumes, a rotor head representation,

and finally, a rotor head fairing were explored. The role of the

rotor head "beanie" and pylon modifications in deflecting the

center portion of the rotor wake downwards was demonstrated.
• .-. ° • • i

2.0 ANALYSIS

2.1 Backuround to the Analysis

Earlier approaches to the analysis of body on rotor or rotor

on body interference has generally been limited to first-order

effects. That is, the presence of the fuselage has been included

in the rotor calculation by means of some perturbation of the

rotor inflow field or the presence of the rotor is considered in

a calculation of the fuselage aerodynamics through the inclusion

of an actuator disc or vortex tube model which alters the fuselage onset flow. No attempt is made, however, to introduce any

coupling by, say, including the presence of the fuselage in the

calculation of the rotor downwash field passing over the fuselage. One reason for this is the complexity of the models

required to adequately represent (alone) the rotor and fuselage

flow fields.

Figure 1 represents a fairly typical vortex filament model

of a helicopter rotor wake in forward flight. This illustration,

taken from the work of Landgrebe, 21 presents the rotor wake as a

series of straight line vortex elements. Every one of these

vortex segements induces a velocity on every other and on the

rotor blade; determination of the equilibrium position of the

wake, and, hence, the rotor inflow, loading and performance

involves the solution of the wake circulation matrix at each

instant in time, with the wake being progressively generated

until some equilibrium shape is reached. This is a procedure

which demands most of the resources of even today's advanced

computing machines. Similarly, in the calculation of the fuselage aerodynamics0 th_ airframe is discretized and represented as

a collection of flat 9a_els, Figure 2, each modelled by singularities whose scr_:7_:5_ ar_ determined by position of the panel

on the body and the o_ fiow. The strengths of the singularities, the unknowns in _:::e_olution, are corventionally determined

by the inverzion of _ :_rix equation involving the influence of

each panel on every _the, panel and the boundary conditions. In

the more advanced a;;al Jes of this type, Ref. 20 is typical,

viscous and se[azat_:_ _ _:low regions can be modelled. Again, as

with the vort6:_ _ila'_:,: c models of the rotor wake, an analysis of

this type _ema_:_s _= ¢f the capability of today's computing

machines, _i_:_ing. _hese two already involved analytic tools

immediately _:eref,:_ _, pre_:ents a problem of computing machine

capacity. F_ure _ pictures an influence block diagram where the

block_ o_ the di_gnal _epre_ent the influence of the fuselage on

the f_se_age _n,l the rotor and its wake on themselves_ _he offdiagonal _loc_ are the ,:oupling terms, rotor on fuselage and

f:_sela_e 6n _:oto_, which must be included for a second-or

higher-o_ _er _olut_on to be achieved and, as was noted above,

sl_ce the solu_ ion :>f each of the major blocks on their own, that

is_ fus:_lag_ on fuselag,_ or rotor/wake on rotor/wake, absorb most

o:[ the machine capu_ity, some way must be found to simplify the

mo_iel i_ e full solution is to be achieved. This is made even

mo_e complex by the fact that although the body/body block is
Figure I. Typical Output from Vortex Filament Wake Model.

i

Figure 2. Typical Panel Representation for Helicopter

Analytical Model.
% " LI

ROTOR

WAKE

FUSELAGE

ROTOR

/ ,/ ,/ /

/ , / /

, " / i

Ry' // //

WR

ORIGINAL PAGE ,_

OE POOR QUALITY

WAKE FUSELAGE

R_ RF

/ •," / /

/////

F R Fll

_// / .,.-"

/// ,///,./

I//,/, ".,, (/}

Figure 3. Body/Rotor/Wake Influence Block Matrix.

TANGENTIAL NO F__AAL

COMPONENT COMPONENT

ERROR IN V, % . ERROR IN W, % .

_ ? _ _ -o._ •

z,

0

.2 ,o

.225 "J .if5

Y

Figure 4. Panel Model with Vortex-Lattice Singularities--

Implicit Velocity Component Errors.
b

J

steady, all the other blocks are time dependent when considered

in the body axis system.

In the present analysis this is overcome by working with a

time averaged vortex sheath rather than the time varying vortex

filament model that is more conventionally used. The vortex

sheath is attached around the edge of the rotor disc and represents the envelope within which the rotor wake will be transported downstream. This type of wake model may be distorted

(relaxed) in the same way as the filament models and will provide

all the inflow distributional effects in the rotor disc plane in

that a rotor blade moving through the inflow field will experience the same low harmonics that would be felt in a filament

model. This is felt to be adequate for performance prediction.

The higher harmonics of inflow needed for rotor loads work may be

simulated, but this is beyond the scope of the present study.

Another technical problem which must be overcome before

coupled fuselage and rotor wake calculations could be achieved

arises from the nature of the fuselage model itself. In the

conventional panel model representation, the flow solution is

strictly only correct at the panel control point. Away from this

location substantial errors can exist and Figure 4, taken from

Reference 18, shows how these can very in a typical case. For

the body/rotor flow field calculation, streamlines close to the

body have to be defined if fuselage induced rotor wake deformation is to be calculated and some way must be found to avoid the

errors pictured in Figure 4 if this is to be achieved.

A solution to this problem is provided by a new method

developed by Maskew and described in detail in Reference 20. In

this approach the surface is modelled using doublet singularities. The use of doublets together with an appropriate choice of

boundary conditions and an interpolation technique which determines local doublet gradients, and through them surface and offbody velocities, gives a continuous definintion of the local flow

field. With the earlier source singularities models this would

have been impossible, especially when the strong vortex elements

from the leading edge of the rotor passed close to the fuselage

panels. Figure 5(a), again taken from Reference 20, illustrates

the ability of the doublet code to handle this type of close

passage problem. Here, the streamlines around a stzo,g vortex

positioned above an airfoil leading ed%e are shown. As can be

seen in Figure 5(b), the calculated pressure distribution is

smooth and the derived off-body streamlines are well defined.

Together the vortex sheath model of the rotor wake and the

doublet analysis provide the tools with which to represent the

flow around helicopter fuselage/rotor combinations. What is

needed now is some means of coupling the wake and fuselage

models. This is provided by a rotor blade element model embedded

within the potential flow solution. The solution proceeds iteratively and is represented by the block diagrams given in Figure

6.

9
*.5 0 .3"

X

Figure 5 (a) . Comparison of Off-Body Streamlines Calculated

using VSAERO with Exact Solution Result.

Cp

-I0

4

Figure 5(b). Comparison of Pressure Distribution Calculated with Program VSAERO and Exact Solution.

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//' , ,

INTEGRAL ME'I'_OS PRESSURES, FORCES

AND "C"E _"rS

IoFF-._c:Y ',E_.:Cl :Y::,_.Y I

NEW wA_ SHAPE]

\

FORCE-FREE w_wE

(SECOND-COOER CURVE FIT

TH;ObGM T_nEE VALUES

• T_O DIRECTIONS

• uNEI.AL INTERVALS)

Figure 6. General Program Schematic.

11
The approach, first outlined in Reference 17, is built

around a panel model of the fuselage and a vortex sheath model of

the rotor wake attached around the edge of a panel model of the

rotor disc. The whole assembly is driven by a blade element

model of the rotor. Having panelled the fuselage and the rotor

disc and set up the rotor wake in some initial position, the

rotor blade element calculation is used to define the variation

of blade loading around the azimuth. Assuming some initial

inflow, the blade element calculation provides the loading distribution which becomes the boundary conditions that are enforced

on the rotor disc panel model. This loading is fed into the

rotor wake sheet and is included in the solution for the strength

of singularities used in the fuselage model. With fuselage

singularities then known and the rotor loading defined it is

possible to calculate the flow field velocitiles around the rotor

wake and to relax (deform) the wake as dictated by the loca_

flow. Following the wake relaxation the rotor inflow field is

re-evaluated based on the new wake position and the blade element

calculation re-run to update the now radially and azimuthally

varying time-averaged, disc loading distribution. The whole procedure may then be repeated as often as is required to arrive at

a converged rotor loading, wake position and fuselage airloads

solution. When this is complete, the loadings represent fully

the effect of rotor on fuselage and fuselage on to+or, including

the second and higher-order effects.

The blade element model used in the present calculation is

the conventional, rigid blade, flapping analysis. Airfoil section data is included in the standard manner. Since the procedure is modularized, it would be very easy at a later date to

replace the performance routine with a more elegant model.

2.2 Body AerodynamicR

The aerodynamics of the body and wake components are calculated using program VSAERO. Program VSAERO (Vortex Separation

AEROdynamics analysis) is a refined surface singularity analysis

which removes the limitations of the earlier generations of codes

(Reference 22 is typical) and provides a much more rigorous

aerodynamic model without sacrificing the simple, flat panel

model of the aircraft shape. The program development was funded

by NASA and the U.S. Navy and has been documented most fully in

Reference 23. Using a combination of source and doublet singularities and modifylng the way in which the boundary conditions

are applied, the program solves for the local doublet strength.

This is then differentiated to obtain the local velocities. The

method of solution has been extended to handle strong external

vortex/surface interactions and is no longer constrained as were

the earlier codes to align external flow vortices along panel

edges. This permits relaxation of the wake (iteration to a

force-free location) without the repanelling between iterations

that was implicit in the earlier programs.

!

12
The program used in the present study was the potential flow

model for general configurations with multiple components. The

program capacity is for 1,000 panels on each side of the plane of

symmetry with an additional 1,000 panels of separated wake. The

wakes may be shed by all the components along any edge (say the

wing tip edge) and any or all wake(s) may be relaxed. Engine

inlet and exhaust flow may be modelled and high energy jets

simulated. The program can also be used to survey the velocity

field off the body, and to carry out iterative calculations of

the viscous/potential flow on the oody surface. Also available

in other versions of the program dr? on- and off-body streamline

capability, coupled viscous/potential flow iterations with extensive separation modelling and time-stepping and harmonic wake

analysis, respectively, for large and small amplitude unsteady

body motions.

The principal problem working against the application of the

old panel codes to the rotor and other highly interactive flows

is their inability to handle wake/surface cutting. In a conventional source or vortex-lattice method, impact of a vortex

element onthe surface anywhere other than along a panel joint

will cause a divergent solution. In VSAERO, because of the

nature of the solution, it simply causes a jump in the doublet

distribution along the line of the cut. Provided that this jump

is accommodated in the surface differentiation used to determine

the velocity field, the resulting solution can be continuous

through the cut. To demonstrate the procedure on a configuration

somewhat simpler than the typical helicopter, a simple test case

was set up.

2.3 The Wake Cuttina Procedure

In order to explore whether the potential jump associated

with wake cutting would violate the basic formulation or at least

cause a numerical problem in the solution, a tes_ case was set up

with a vertical surface ahead of a wing, Figure 7. The vertical

surface is modelled with a 3 x 12 panel array on a zero thickness

lifting surface. The tips of this surface have a 5 degrees

backward rake. The vertical surface is set at 20 degrees to the

x-axis; i.e., t_ carry a side force directed to the plane of

symmetry. The horizontal wing has a rectangular planform with an

aspect ratio 4 and a NACA 0012 section: it is modelled with a 24

x 20 array of panels on the main surface and a 3 x 12 array on

the half-round tip. The spanwise panelling is arranged with

concentrations near the tip and just outboard of the spanwise

station of the vertical surface trailing edge (i.e., where the

maximum vortex wake interaction is expected to occur). The onset

flow was set _t I0 degrees.

Figure 8 shows two views of the calculated wake geometry

after three wake shape iterations. The side view, Figure 8(a),

demonstrates a very nice behavior of the basic vortex roll-up

calculation on the vertical wake. The top view of the wake

13
.ORIGINAL_ PAQE iS

OE [_)OR QUALITY

-1.5"

1.0

1.0

NACA 0012

2,(

e

+

X

Figure 7.

TOP VIEW SIDE VIEW

Schematic of Wake Cutting Model.

+

I

I

i

14
!!

i

ORIGINAL PAGE I_

oE POOR QUALn_

TIP VORTEX BUNDLE

X SHEET

TIP VORTEX BUNDLE

Figure 8(a). Wake Cut+%ng Study--Side View.

P

i

Ii

Figure 8(b) .

\

1 /

Wake Cutting Study--Tot View.

15
ORIGINAK PAGE I11

,,, OF+POOR qUALI'r_

row.

t.4 -- 1.5 -

O._

i.O

I4,J

-O._'

-0.4

Figure 9.

I+0--

0.5 --

-O.S

A A

IR SURFACE

[R SURFACE

V V V V V

! l i l u v 1 i

T

A

V

V

l I ! I

O.O O.S 1 .O

v

+777v

%

v u I i l ! I' !

.5 a.o a.5

Wake Cutting Study--Spanwise Doublet Distribution

(0.25 Chord) .

16
Y.. .

ORIGINAL PAGE I,_

OF POOR.QU_,II_

!

-!

-2

UPPI

LOWI

T

I1,!

:R SURFACE

:R SURFACE

v

I I ! I

I

@.,

@

@

e

l I I T

e.e

¢1

' 'T' ! I !

%

A

7

I 1 I I

e.1 1.0 L

I

S a.e a.s

Figure i0. Wake Cutting Study--Spanwise Velocity

(0.25 Chord).

t

17
geometry (Figure 8(b)) shows reasonably good behavior except that

beyond the wing trailing edge the effect of widening the spacing

of the grid planes used in the wake calculation, for reasons of

economy, can be seen.

A spandise cut through the doublet distribution, Figure 9,

shows the expected jump in doublet value (i.e., jump in surface

perturbation potential) due to the vertical wake intersection.

The jump occurs at slightly different spanwise locations on the

upper and lower surfaces due to the local tilt of the wake panels

at this station (x/c - .25). It is easily verified that the

spanwise doublet 9_adient (i.e., Vy), plotted in Figure 10, is

essentially continuous as we pass through the wake--only the

doublet value is discontinuous. During the analysis phase of the

calculation, where the doublet distribution is differentiated to

obtain the velocities, the program senses the wake cut induced

jump and selects the appropriate differencing scheme. Similar

good behavior was also noted in the other velocity components and

in the surface pressure distributions. The vortex pazr from the

vertical surface clearly dominates the spanwise flow on the wing.

The otherwise inboard flow (negative Vy) on the wing upper surface has been totally reversed by the vortices except in the very

tip region where there is clearly still some flow moving around

_rom the lower to the upper side.

._.4 Rotor Blade Aerodynamics

Rotor blade aerodynamics _s calculated using a simple blade

element theory model. Inflow velocity, after the first iteration

during which the inflow is assumed constant at the momentum

value, is calculated in the body/wake portions of the program and

passed over to the rotor internally. This includes all three

components of velocity and contains, in addition to the contribution of the fuselage, the wake induced terms. These velocity

components coming over from program VSAERO are in the body coordlnate system. As a result they must be transformed into the

rotor control axis system to be considered as elements of the

inflow velocity.

In the conventional rotor calculation the velocity components at the rotor disc control plane are commonly represented

as below:

Voo

__/U (Inflow)

_Control Plane

: Axis

' Contiol Axis Angle

18
In the present case the situation has been generalized by

introducing three components of induced velocity which contain

the influence of the fuselage and the rotor self induced (through

the wake) effects. The more general case appears as below:

_Iv v_ + Vc = Climb Velocity

V + V x

_IW_I Plane

/ Control Axis Angle

and includes an out of plane lateral component, Vy.

Following convention these are resolved into and

perpendicular to the control axis system and become

(V + vx) cos_ -

(Vu + Vz ) cos_ +

(V¢ + Vz ) sin

(V_ + Vx) sin

vy

Still following normal rotor conventions these are then

resolved into the blade axis system where they take on the form:

Perpendicular to the Blade

Up =- (r - el_°+ [(V + V )sin_ + (Vc + V z) cosa 1cosg

-Ill(V= + Vx)COSe- (Vc + Vz)sins#_ cos_ + VY sin_] sins

U I = _r + ,!(V® + Vx>COSe - (Vc + Vz)sine !) sin_ - Vy cos$

[_I(V_ + Vx) COS_ - (Vc + Vz)sin_ I cos_ + Vy sin_J] cos_

+ [cv® + v >si_ + Ivc + v }cos_] sinB

19
The induced velocity components are updated as the rotor

body iteration proceeds. As an example, on the first pass

through the rotor analysis, the vertical induced velocity, V_

will contain only the prescribed momentum inflow velocity. On

the second pass, it will contain the first-order body induced

upwash (the isolated body effect) and the induced velocity from

the skewed vortex sheath model of the rotor wake. The strength

of the rotor wake is set by the rotor loading calculated in the

first iteration. On the third pass, the V. term contains the

body effect, this tlme Includlng second-order rotor and wakeinduced components, and the inflow from a now distorted sheath

wake. Subsequent iterations add higher-order corrections. Experience has shown that for low-speed flight, 0.05 advance ratio,

two wake relaxations are required (3 passes through the rotor

code). At high speed, advance ratios of greater than 0.15, two

have been found to be adequate.

Convergence was based on several parameters, chief among

them being the behavior of the _otor wake. This was followed

from one iteration to the next and was observed to stabilize very

quickly as the interactions were included in the wake relaxation.

Other criteria included rotor total and blade section parameters,

and perhaps most sensitive, the local velocity at rotor panel

centers passed back to the blade element calculation by the

fuselage/rotor wake segments of the program.

No small angle assumptions have been made to permit

extension of the analysis to propeller, tilt rotor cases at some

future date.

With the blade section onset flow determined, the local

angle of attack is known and the section and blade loads can be

calculated in the conventional manner with radial and then azimuthal integrations. Only rigid blade flapping is permitted. No

aeroelastic effects are considered. The rotor module has deliberately been isolated from the rest of the calculation so

that, if required at a later time, a more involved rotor algorithm may be substituted. The present module may be operated at

prescribed collective and cyclic pitch settings or may be allowed

to iterate to requested gross weight and rolling and pitching

moment targets.

20
2.5 Rotor Disc Panel Model

Coupling of the body and rotor aerodynamic phases of the

calculation are carried out through the panel model of the rotor

disc, Figure Ii. This happens in two ways. _he first is the

straightforward model of the incremental velocity added to the

flow by an actuator disc (or more correctly, mosaic) where the

velocity on individual "tile" or panel is the impulse produced by

the time averaged loading on a disc segment bounded by preidentified radial and azimuthal boundaries. The second is the

way in which the radial distribution of loading is integrated to

provide a local doublet strength and ultimately provide the

strength of the vorticity passed over into the wake attached

along the disc inner and outer edges. This is illustrated in the

schematic in Figure 12.

One of the features available in program VSAERO, 23 is the

"type 4" patch. This type of patch allows actuator disc panel

modelsof propellers and rotors to be constructed and requires

that both the source and doublet parts of the singularity be

specified.

For a type 4 patch, the doublet strength is given by:

r

bC a =_ P

2 (V'/V)

O

dr

bCp is the panel differential pressure coefficient (upperlower), V' is the resultant local velocity at the panel and V = is

the onset flow velocity. The incremental loading for the panel

model may be easily determined from the blade element calculation.

bE

bC =

P _oV®2_A

where bL is the time averaged segment load and _A th_ seglaent

area and aL is given simply for a particular azimuth and radial

location by

l

p4

"'i

bE- n_LL __b

_r " NC

In this expression, bL/_r is the blade lift per unit span, _R is

the segment spanwise extent, b is the number of blades and NC is

the number of azimuth increments into which the disc circumference is broken.

21
BLADE

ELE.MENT

MODEL

ORIGINJsJ_PAGE m

oE pOOR ,QU_

ROTOR PARAMETERS

ROTOR CALCULAT! ON

BLAI_E ELF..FJENTMOOEL

I

INFLOW

DISC

la)ADING

BODY SHAPE

INITIAL WAKE SHAPE

i BODY/WAKE CALCULATION PANEL MODEl.

,

DISTRIBUTION

PANEL

MODEL

I

L

ROTOR LOADS

PERFORMANCE

AIRFRAME AIRLOADS

INTEGRATED LOADS

WAKE SHAPES

OFF-BODY VELOCITIES

Figure ii. Body/Rotor Interference Calculation Schematic.

i i

i , i = I

LOAD|I_G INTE_iRATF...DI RADI!ALL_ TO _-._.

FORM D_OUBLETISTREI__TH _ o_-_"_'_

| .- .... ' ' TIP lOADING'

I _-,_ _,-.-_ _', SHED, [0 WAKE] '

,

"_'_ _¢f"" CARRIES SPREAD ([(ME-AVERAGED)

° -"- ' _ BLADE LOAD

BLADE ELEMENT MODEL

SPANWISE LOADING

-. LOCAL PANEL LOADING SETS

_PANEL SOURCE S[RENOTH

{NNER WAKE "- .

.'-_,,. V I N

VOUT

WAKE MODELLED WITH PIECEWISE

CONSTANT ELEMENTS BETWEEN

: GRID PLANES;-.-.._

DOUBLET GRADIENT GIVES

WAltZ STreNGTH

(VouT - VIN)

QLI_.ERWAKE

VIN- _

Figure 12. Blade Model/Panel Model/Wake Schematic.

22
Definition of the source term in the singularity expression

requires the normal velocity on the disc panels and this also can

be determined from the blade element calculation output and

simple momentum theory. The doublet strength and the normal

velocity for each panel are passed over from the rotor code to

the body solution £nternally and require no user input.

The rotor wake is automatically fed by the doublet strength

determined for those panel fringing the disc model using the

basic VSAERO capability, with the wake sheath being fed by the

tip doublet strength, the sum of the spanwise integration of

loading, the inflow calculation has the same order of fidelity as

a filament model of the wake in which only the tip vortex is

modelled.

23
3.0 THE CO:_PU_ _ROGRAM

The Body-Rotor analysis program (BodRot) is an extension of

the basic program VSAERO.23 The rotor program is a self contained subroutine called by the main program whenever a type 4

patch, a rotor or propeller disc, is loaded. The code is available as an update deck to the basic program. The analysis has

been loaded and executed on Control Data Cyber 176 and 7600

machines and has recently been demonstrated on CRAY. The only

additional data required beyond the general configuration description is the rotor details and the blade section airfoil data.

Part 2 of this technical report contains the program user's

guide, sample input and output listing and a listinq of the rotor

subroutine source code.

24
I,

t

0

" I

.t

4.0 DISCUSSION OF RESULTS

4.1 Application of the Analysis to a Simple Body/Rotor

To evaluate the combined body/rotor analysis program, a

simple test case was set up with the fuselage sized to exaggerate

the body/rotor interference effects. Figure 13(a) and (b) shows,

respectively, an oblique view and a side view of this test configuration. The panelling on both body and rotor is deliberately

less dense than in a practical case in an attempt to preserve

some clarity in the illustration. The rotor panelling is set to

coincide with the azimuthal stations at which the blade element

calculation is performed, while the panel centroids are coincident with the blade element control point locations. Also shown

in this figure is the initially prescribed wake location with

only the streamwise legs of the wake panels shown for clarity.

Note that the wake has been "draped" over the fuselage. At this

stage in the development of the analysis, the calculation is

limited to cases where the wake passes either completely above

the body or totally immerses the body.

In Figure 14(a) and (b) the rotor wake is shown from the

same view points as above after two iteration cycles. For this

case, which represents a rotor at a low advance ratio and a

thrust coefficient solidity ratio (CT/_) of roughly 0.08, the

very pronounced wake roll-up that wo61d be expected is clearl

seen. The rotor wake behaving in its time-averaged form muc_

like the wake of a low aspect ratio wing begins to roll up at the

outer edges as it leaves the trailing edge of the disc. The

roll-up process begins well ahead of the rotor shaft axis as can

be seen by the pronounced divergence of the streamwise wake

lines. In the side view, Figure 14(b), the presence of the body

is evident in the path followed by the lines from the disc

leading edge, much higher than would be expected from the rotor

in isolation. Body influence is most marked, however, in the

plan view, Figure 14(c), where the wake roll-up is taking place

at a spanwise (lateral) location further outboard than in the

isolated rotor case. The plan view for the equivalent isolated

rotor case is shown in Figure 15. With the capability to calculate the wake trajectories and coupled behavior demonstrated with

the simple model, the study was broadened to examine more representative shapes.

4.2 Basic Body/Rotor Results for a Representative

Confiuuration

The model chosen for the study was that tested by Freeman

and Mineck4 scaled to a rotor radius of 20 feet. This size was

chosen to give a full scale machine typical of helicopters in the

small to medium size range. A target vehicle weight of 7,000 lb.

gave a rotor thrust coefficient of 0.00554. Calculations were

made at three advance ratios: these were 0.05, 0.15 and 0.3.

25

- Q,

i

t ̧
F

Figure 13(a). Body Rotor Interference Study--

Starting Condition.

PRESCRIBED

CENTERLINE

_._. I..._ FILAMENT

m,

BD.,

Figure 13(b) . Body Rotor Interference Study--

Starting Condition.

26
Figure 14(a).

Figure 14(b).

Body Interference Study, Relaxed Wake.

CALCULArEDCENrERLINE

F[LARENT

Body Interference Study, Relaxed Wake.

27
°

\

Figure 14(c). Body Interference Study--Relaxed

Wake

Figure 15. Isolated Rotor,

Re I axed Wake.

/

28
I

The two low values were chosen to provide an overlap with the

Freeman data, the high value to give a more realistic case for

studying body rotor interference close to the cruise condition.

The body is shown in outline in Figure 16 comparing the original

model and calculation scales.

Figure 17 shows an oblique view of the basic panel model of

the fuselage and rotor. Correlation with measured surface pressure data Is generally acceptable. Certainly on those regions

most influenced by the rotor wake, the fuselage aft of the rotor

center, agreement is good. Figure 18 shows comparisons at sta T

tion 17.6 for two advance ratios. The comparison is presented at

the same data scales as in the original report,4 at advance

ratios of 0.05 and 0.15. Correlation over the front fuselage,

Figure 19, is less good. This probably results, to a certain

extent, from the relative crudity of the panel model in this

region, and from a mismatch in the calculated and actual angles

of attack. Wind tunnel corrections had been allowed for the test

data. In the analysis, the lower surface was made deliberately

sparse to allow for increased panel density on the upper and aft

sections. Sections within the rotor influence should be less

sensitive to model rigging angle/analysis angle mismatch than

would be those sections on the nose. For comparison purposes the

isolated body results are superimposed on the open scale (u -

0.15) plots. The full correlation set for all the stations at

which data were measured in the original test are shown in Appendix A of this report for the advance ratios at which analysis and

test overlapped.

An oblique view of the low advance ratio case, Figure 20,

shows the wake draped over the fuselage. This is explored further in Figure 21 where cross sections are taken through the wake

as it develops aft and downwards. It is interesting to note the

roll-up of the edges of the rotor wake. Since the wake is made

up of rectilinear vortex elements, some crossing is to be expected in the model. Inpractice, however, the wake simply coils

around itself. In the model, the wake leaving the leading edge

of the disc moves first upwards as it is convected aft. It then

passes downward, held up above the isolated rotor position by the

flow developing around the fuselage. The wake cuts the pylon,

recombining above the tail cone in the region aft of the rotor

head. In the absence of detailed experimental data on wake/body

cutting, it is difficult to judge how valid the present model is

in this region. It does, however, appear to behave as a membrane

made of discrete, traverse vortex elements would act on an

oblique approach to a surface, first deforming as the surface is

sensed and then, at some point, dividing with the torn edges

rolling up along the surface. The theoretical wake does not, of

course, tear. Here, the streamwise elements defining the edges

of the wake panels pass, following the external flow lines, over

the body, reacting to their reflections in the surface while the

29

I
Y/R • o.o

Y'O.O

ROTOR AT _ - 0.0"

ORH_NAE PAn

OF. POOR (_.IALIT_

0 L.O

I !

ROTOR

RAD lUS

Figure 16.

Outline of Freeman-Mineck Model used in Ama]ysis

(Scaled to 20 ft. Rotor Radius).

Figure 17. Basic Panel Model for Body/Rotor Studies

(After Freeman and Mineck, Ref. 4).

3O
Figure 18 (a) .

ORIGINAL PAGE Ilk

oE ,qUW 13 

20-

7

to .'_0

-LO

-20

LOW_ SURFACE l

UPPL_R SURFACE ,_

ANALY ;IS

,\ A

r

t

A

'I----------

]

I

f

! ! " !

-G -4 -2 0 ,._ 4 6

Y

Aft Fuselage--Typical Correlation with Test Data;

0.05 Advance Ratio, Run/Point 22/139, Station

17.6 (0.88).

L.e

0.

-e.s i

-1._ 1

-6

_ LOWER SURFACE )

V UPPER SURFAC_ ANALYS

• TEST DATA i/

i

s ..---,--ROT(

2

R OFF

FLAG ;ED SYMI;OLS.-I'E_TDATA .OWER SURFACE

l T ! m v I

-, -2 0 a 4 G

%,#

Figure i8(b). 0.15 Advance Ratio, Run/Point 25/148, Station

17.6 (0.B8).

31
0tRI_NAI/PAGE I$

OE POOR QUALi'I%:

20 -,"

L_ LOW

t _7 UPF

• TE_.

I0--_

0

-iO

-20 • , , ,

F.RSURFACE

_R SUR _F,

F,.DAI'_

t

il I "_

i v v

/' ANALYS S

i

A A A

- -ev _'0"_6 •

- e

v

w "_-

[111

2 3

Figure 19(a). Front Fuselage--Typical Correlation with Test

Data--0.05 Advance Ratio, Run/Point 22/139,

Station 1.80 (0.09).

_.5--

-0.5 -

LOWER

YS - ROTOR OFF

TEST DA[ /" _ :

d 'I

FLAGGED S'_'MBOLS--TES']

: 1 ; I I I it" l I

- e

,T--.-

DATA LOWEI;

1 ! I !

SURFACE

I ! I I

L a

Figure 19(b). 0.15 Advance Ratio, Run/Point 25/148,

Station 1.00 (0.05).

32
OR1QINA[ PAGE IS

OF'. POOR QUALIT'(

Figure 20. Calculated Distorted Wake--Low Advance Ratio.

Figure 21.

20

-2O

r

R0[0R

- DISC

, ! liT i r T f | T T T '

""T" , T ! [ T * ' _ ! T !

-]O -29 -tO 0 10 a0 30

Cross Section through Typical Low Advance Ratio

Wake (Station 30.0) .

33
cross elements pass through the surface connecting to the adjacent streamwise filament. These carry with them, into the body

solution, the jump in doublet strength which accounts for the

change in flow conditions from outside to inside the wake.

The effectiveness of the procedure for handling the wake

cutting problem is shown in Figure 22(a) where the doublet distribution along the pylon waterline cut is shown. The jump in

strength across the wake edge is clear. In Figure 22(b), the

velocity component in the streamwise direction shows no perturbationc other than the changes in velocity associated with the

changes in shape.

Based simply on the relative magnitude of the rotor inflow

velocities, the effect of the rotor on the fuselage would be

expected to decrease with increasing flight speed. This shows

very clearly in Figure 23 where the vertical velocity component

along a horizontal cut close to the model maximum is plotted for

values of advance ratio of 0.05 and 0.15. In both cases the

fuselage was set at the same angle of attack and the rotor was

trimmed for level flight at 7,000 lb. GW and 12 ft.2 of drag with

nominally zero pitching and rolling moment. At u - 0.15, the

rotor-induced vertical velocity is very small and differences

between right and left (advancing and retreating) sides are

slight. At u - 0.05, however, where the rotor downwash is a

substantial fraction of the forward flight speed, the side to

side differences are quite marked with the more highly loaded,

inboard, advancing side showing the stronger effect.

The trend of the influence of the fuselage on the rotor with

speed is, of course, reversed as it is stronger _t higher speeds.

This is best illustrated by comparing the blade angle of attack

distribution at an inboard and outboard station (0.4 and 0.9

R/RTI P) for advance ratios of 0.05, 0.015 and 0.30 shown in

Figures 24 and 25. As expected, the strong upwash over the nose

and downwash over the aft fuselage are reflected %n regions of

increased and decreased angle of attack. At u - 0.05, the fuselage effect on the rotor is almost negligible; at _- 0.30, it

dominates the picture. These dramatic changes in angle of attack

result from the fuselage-induced flow field. Earlier analyses,

the work of Landgrebe et al.2 and Polzll,12 are typical, failing

to include coupling between fuselage and rotor flow fields,

produce an inflow distribution that is symmetrical. The present

approaches with full coupling shows a more nonuniform upwash,

Figure 26, and one which certainly, in reflecting the lateral

differences in rotor loading, is not symmetrical about the fuselage centerline.

34
ORIGINAL PAGE Ib

OFPOORQUALIT 

2.T

1

1o-- v

$ -

_ O" "

-10

6 - T

0-

-2

-4

-6

! [ I

f

OUTSI.!

i

BOARD SID

!

PE WAKE

_,,..

I_DOUBLEi JUMP

IACROSS WAKE

I

I

INSIDE

I !

WAKE

!

8 10 IR 14 t6 10 IN

X

(a) Calculated Pylon Doublet Distribution--Low Advance Ratio

(Waterline 3.00) .

i

)

°

_,B w

| ,S • "_

V L.0

X

0.|

-_.$

6

4

-- 0"

-4

-I.O "- -6

A STAIBOARD SIDI

HAKE EDGE

WAKE

1

i i l i ! !

1o l_t 14 te

I

I

tO Oo

i

!

(b) Calculated Pylon Horizontal Velocity Ratio, Vx/VINF--

Low Advance Ratio (Waterline 3.00).

Figure 22. Typical Effect of Rotor Wake/Pylon Intezsection--

Body and Rotor Alone.

35
v

Z

l.er

Z_ PORT l

STARBOARD

ORIG_AK PAGIE I1@

OF POOR QUALITY

I

10 20 3e

x

(a) Advance Ratio 0.05.

4e

(b) Advance Ratio 0.15.

Figure 23.

Calculated Vertical Velocity Ratio V z

Waterline 0.0; Body and Rotor Alone. /VINF;

36
I

I,

W

b

b

rT

i

r

Figure 24.

ALPHA

(DEG)

ALPHA

(DEG)

ORtGINAL PAGE. i_,

OE POOR QUALITY.

_5

'R =O,q

ISF)I;UFD ROTO I

WITH BnDY I

t

_I r/R = 0.9

J

l

-5 t T v _ I • v T T

o l@O

" _ i ! ! ' " ! ! f "!

zee 3ee 4ee

Calculated Effect of Body on Rotor Angle of Attack--

0.05 Advance Ratio.

tO-

... rlR = 0._

i r/R • 0,9

e

LSOLAIED ROTO_

WITH BODY

-/I\

toe 2ee 300 40_

c

I

t!

Figure 25. Calculated Effect of Body on Rotor Angle of Attack--

0.30 Advance Ratio.

37
FWD

/ ,"16t\ \,

; j t', ',

\ "I, 1

:24, i

ORdeAL PAO,£ iS

OF_ I_;)OR QUALITY

_R = 650.0 FT/SEC

Figure 26. Calculated Contours of Upwash Velocity in Rotor Plane;

Body and Rotor Alone; 0.30 Advance Ratio.

Figure 27. Body-Rotor Panel Model with Tail Assembly.

m_

38
4.3 The Effect of Confiquration Elements on the Rotor/

_uselage Flow Field

Adding representative horizontal and vertical stabiliser

surfaces to the basic configuration has only a weak influence on

the rotor flow field, but serves as a good example of the ability

of the analysis to handle calculations where fuselage elements

are embedded in and, in fact, pierce wake regions. For this

configuration the tail surfaces, shown superimposed on the basic

fuselage in Figure 27, were modelled using the "type 3" patch

_4_aDility of Droqram VSAERO. 23 This permits the use of simple

1_tlng surfade sections, modelled with a doublet lattice, to

economically represent components in a situation where only

lifting (as opposed to thickness) effects are important.

Figure 28 shows the chordwise and spanwise pressure distributions on the vertical stabiliser, with the lower portion immersed in the main rotor wake. The success of those portions of

the program dealing with wake cutting can be measured by the

smooth vertical development of pressures.

As might be expected the presence of the tail surfaces has

only a small effect on the rotor loads, most of this coming from

the slight retrimming required to compensate fOE the disturbance

a loading on the rear of the disc caused by the small amount of

circulation about the horizontal stabiliser. Figure 29 shows the

comparison of blade angle of attack before and after adding the

tail surfaces.

The further addition of nacelle units, Figure 30, however,

has a substantial effect on the rotor flow field. Mounted on

either side of the pylon they produce changes in the upwash field

with an interesting indirect effect illustrated in Figure 31.

This shows the angle of attack variations with azimuth at radial

stations of 40% and 90% of the blade radius. Although mounted at

900 and 2700 azimuth, their direct effect is seen all around the

azimuth. However, the retreating blade change occurs in the

relatively low energy region of the disc; for advance ratios

higher than the 0.15 of the present example, it would be in the

reverse flow region, and so the lateral rotor loading is put out

of balance. The rotor must, therefore, be retrimmed and this is

the indirect effect that is causing the observed changes at the

outboard station.

Adding a mass in the center of the disc has an even more

dramatic effect. The size of the blockage was chosen to represent the volume and frontal area of a rotor head with some kind

of vibration adsorption equipment superimposed. The separated

flow behind the rotor head was modelled and the wake transport

and deformation was calculated in parallel with the deformation

of the main rotor wake. Figure 32 shows an oblique view of the

configuration and the calculated wake trajectories. The main

rotor wake is present, but has been omitted from the picture for

reasons of clarity. The calculated path aft and down matches the

39

J
1

I

OUTSIDE

INSIDE WAKE EDGE

ORIGINAL PAQE IS

oE _ QUALrn'

Figure 28. Pressure Distribution on Lifting-Surface

Model of Tail Assembly--

0.15 Advance Ratio.

IsleALPHA

(DEG}

S

-5

ROTOR AND BODY

ROTOR + BODY ,.L

R/R - O.4 /_.._.,

dLS

R/R - 0.9

! ! ! u v , v 1 l v ! ! v 1 v T

e 1ee aee 3ee 4ee

BLADE AZIMUTH (DEG)

|

Figure 29. Effect of Tail Surfaces on Calculated Blade

Angle of Attack; 0.15 Advance Ratio.

4O
ORIGINAL PAGE I$

OF.POOR QuALr_

Figure 30. Basic Body Model with Nacelles and

Tail Surfaces Added.

Figure

IS--'

tO -

ALPHA

(O=G)

(

RpTOR AND ,3CDY

ROTOR ÷ BODY + NACELLES

I

I

_R = O.q

e/R - 0,9

T I" I

/

=,

I _ '1 I i I 1 i I" 1 I I

,,ee 3ee 4et

BLADE AZIMUTH (DEG,)

31. Effect of Nacelles on Calculated Blade Angle of

Attack--0.15 Advance Ratio.

41
-'--'-_" ___ REPRESENTATIVE CFd]SSSECTION

Figure 32. Calculated Rotor Head Model Wake Development--

0.15 Advance Ratio.

E:O-I

10

0

-LO

<_._

I 1 • ! ! ] r i "

30 -20 -10

H_O_ WAKE ' " '

EN61NE WAKE

'" _ , i v i, l v u v v ! I T T V ! !

0 10 2_ 30

Figure 33. Rotor Head Wake Cross Sections--0.15 Advance Ratio.

42

I

j,
observations of many authors. It should be pointed out here that

no attempt was made to model the effects of hub rotation and,

consequently, any lateral displacement is the result of asymmetry

in the steady flow about the fuselage and through the rotor.

Figure 33, showing cross sections through the aft fuselage and

all the wake elements at the mid-span and tip of the downstream

blade shows the wake convection more clearly.

The presence of the rotor head has a very strong effect on

the blade behavior. Forced to accommodate not only the flow

distortion associated with the presence of the rotor head mass

itself, it must also pass through the "dead" region of the separated wake. The blade _esponse to this perturbation domimates

Figure 34 where the azimuthal variation of blade angle of attack

at two stations is presented. Clearly, this will have a profound

effect on the calculated aeroelastic response of the blade and

could well be softened when elastic effects are included. The

substantial changes in cyclic pitch required to retrim the rotor

in the presence of the rotor head can be seen in the differences

in local angle of attack at the outboard station also shown in

Figure 34. It should be pointed out that this treatment of the

rotor head as a simple bluff shape with no base ventilation is

almost c_rtainly exaggerating the effect. Further work is required in this area.

Several studies, Reference 24 is typical, have shown how the

roto_ head wake and regions of upper body separated flow can be

controlled by the addition of a rotor head cap or "beanie" or by

contouring of the aft pylon to provide an eoge separation which

rolls up and convicts the separated flow out of the region of

harm. Both these devices were studied using the analysis.

Figure 35 shows the panel models of rotor head cap and the aft

pylon modification and their wakes. The success of both devices

in depressing the disturbed flow and the cente_ of the main roto_

wake is clearly seen in the wake cross sections at the trailing

edge of the rotor disc shown in Figure 36. The strong roll-up of

the wakes on the rotor cap and modified pylon shapes contrasts

sharply with that noted on the basic rotor head. Again, although

the main rotor wake was included in the calculation, it was

omitted from the drawing for reasons of clarity.

The presence of the flow control devices was also felt in

the blade aerodynamics. Figures 37(a) and (b) show the azimuthal

variation of blade angle of attack both inboard and outboard and

the softening effect of the flow control can be clearly seen in a

comparison with the untreated, basic rotor head result in Figure

34. Plots of blade angle contours over the whole disc show again

the softening effect of the flow control devices.

Again, it shoul@ be stressed that the rotor head effect in

this study has been exaggerated to illustrate the analysis. A

study of the actual flow field would require a mora detailed and

representative model especially regarding the base _low.

°

43
ORIGINAL PAGE. t_

OF, I_OOR QUALITY

L_

I

r

(DEG)

_,iti

0 100

!

- RESPONSE TO ROTOR HEAD WA_E

.... ROTOR * BODY

I _,IT_RoTo_.EAO

:,",( .... t .... "

2ee ]ee 4ee

BLADE AZIMUTH (DEG.)

Figure 34. Effect of Rotor Head and Wake on Calculated Blade

Angle of Attack--0.15 Advance Ratio.

Figure 35(a). Calculated Rotor Head Cap Wake Development--

0.15 Advance Ratio.

44
i

Figure 35(b).

2°I

t0 _

-Ig

ORIGINAL PAGIG. |b

oF. POORQU/UJ_

REPRESENTATIVE CROSS SECTION

Calculated Rotor Head Wake Development with

Modified Pylon--0.15 Advance Ratio.

f ROTO,R C_PWAK_

ROTOR WAKE _ _T_ © j ST. 2Q.O

L:.Ij

i i ' I r _ ' T

3g -20

ST. 3tl.0

ENGINE A ES

1 ! -' , ! T I u 1 _ I T 1 r ir ! v

-Lg 0 10 80 30

Figure 36(a). Calculated Wake Cross Section with

Rotor Head Cap Wake--C.15 Advance Ratio.

45
ORIGINAL PAGE it

I OF POOR QUAL

.

o:. _o,.o,_,-o__ .o

0 34.0

* i { i

-30 -20 -10 0 I_ 2e 3e

Ik

= Figure 36(b) .

°

Calculated Wake Cross Section with

Modified Pylon--0.15 Advance Ratio.

46
ORIGINAL PAGE IS

OE POOR QUALIT'_

Figure

ALPHA

15

J

r/R --0.4

BO Y + ROTOR

÷ ENGINES

W IIFIROTOR HEAD C

i _ r I i 1 1 i I I

2ee lee 4ee

_P

BLADE AZIMUTH (DEG.)

37(a). Effect of Rotor Head Cap on Calculated Blade

Angle of Attack--0.15 Advance Ratio.

ALPHA

(DEG)

Isle

-s

r/R = 0,4

• _A __ J

9

W[' H PYLON MOP

ADDED

-_IY*RUFOR

• NACELLES

lit! IYTI lilt II11

I _il aO0 300 4_

BLADE AZIMUTH _DEG.)

Figure 37(b). _ffect of Pylon Modification on Calculated

Blade Angle of Attack--0.15 Advance Ratio.

47

_°
4.4 The Potential to Study Full Confiourations

The section above has shown how the analysis can De used to

explore the influences of detailed configuration changes. If

program capacity were increased Zrom the current limit of i000

panels, the analysis would have the capability to explore full

vehicles, including multiple rotors. This is illustrated in

Figure 38. Here, as an illustration, the full machine is modelled, fuselage, pylon, engine nacelles, horizontal and vertical

stabiliser, rotor head cap and, finally, main and tail rotors.

For the lifting components vortex sheath wakes, not shown in the

d_awing, are attached around the edges. The rotor head cap and

the stabilisers are modelled, in the interest of economy, using a

lifting surface rather than a full surface singularity model. If

machine capacity had been available, a full panel model, including thickness effects, would have been used. Even with this

simple model further stripped down to fit within the 1000-panel

constraint by removing the nacelles, the effect of the tail rotor

on the vertical stabiliser in the presence of the main wake can

be demonstrated.

In Figure 39 the chordwise loading on the vertical fin in

the presence of the main rotor with and without tail rotor is

shown. The tail rotor is mounted and is operating such that the

fin is on the inflow side. The graph shows the difference in

pressure coefficient between the port and starboard surfaces of

the vertical stabiliser. No attempt has been made to trim the

lateral forces on the uail rotor/fin combination and the tail

rotor is simply operating at a set fixed collective. As a consequence, the vertical stabiliser is operating in an inflow field

which tends to generate side force counter to that of the tail

rotor. This is, of course, unproductive and in practice the two

would be adjusted to complement rather than fight each other.

At the 0.15 advance ratio of the calculation, the influence

of the tail rotor on the main rotor was almost negligible (less

than 0.1% of the total). This could be caused by a direct tail

rotor on main rotor effect; by a second-order tail rotor on tail

surface and then main rotor effect; or it may be the result of

the cycle to cycle imprecision in the iterative process.

48
ORIGINAL PAGE I_

OF. POOR QOALI'r_

Figure 38. Mock-up of Full Configuration with Rotor Head Cap.

(Actual Run Exceeds Present Program Panel Capacity)

k) 4 !

• ', T'7, i

" _ I MAIN ROTOR

0 --. OT _ _ I '

-0.2 3S 36 37 38 39 40 41 42

X

Figure 39. Effect of Tail Rotor on Vertical Stabiliser

Loads at Waterline 5.75 (Main Rotor Present--

0.15 Advance Ratio).

49
5.0 CONCLUSIONS AND RECOMMENDATIONS

An analysis has been developed which permits a fully coupled

solution of the rotor and airframe behavior of realistic helicopters. The effect of the rotor/body coupling on the rotor

inflow has been illustrated and the significance of the effect of

configuration elements, particularly the rotor head and rotor

head cap devices, demonstrated- The role of the rotor head cap

and pylon flow control devices in moving the separated wake and

the center of the main rotor wake was shown for a typical helicoper configuration.

Several additional steps must be taken to verify the usefulhess of the analysis beyond the performance and handling qualities applications of the present study. In order to be useful to

the dynamicist, the higher harmonics of loading associated with

discrete blade vortex encounters must be added (how this can be

achieved was discussed in the earlier report on this work by the

present authorl8), and the resulting blade loadings coupled to an

aeroelastic analysis. This would be further facilitated if the

program capacity were expanded beyond the current 1000-panel

version. If this were done, much greater detail in rotor disc

modelling would be possible than is possible with the present,

_elatively crude, 32 azimuthal steps in the rotor solution.

Despite the relative crudity of the model, the author feels

that with this analysis it is now possible to explore the highly

interactive flow field around the helicopter and that other

configurations previously analysed with relatively empirical

approaches can now be studied in detail with a representative

model. The most likely candidate in this category is the tiltrotor concept, where now, with the demonstrated wake cutting

capability of the present program, a full analysis of the dual

rotor, wing, fuselage flow field becomes possible.

Although the initial results of the study are encouraging,

one area of weakness has been identified. In the discussion of

the correlation it was noted, particularly at low advance ratio,

that the wake skew angle did not follow the expected values

beyond the rotor near field. This has been identified a consequence of the way in which the type-4 (higher-energy) wake is

modelled in the basic VSAERO program. In this report the wakes

were modelled using a piecewise constant doublet segment between

each wake grid plane. Relative to the way in which the wake

behaves in the zeal world, the stepwise constant model deemphasises the far-wake effects. Modifications to the wake model

which would correct the problem have been identified and should

be incorporated if work is continued on the body/rotor program.

50
p

6.0 EE2._J_tC_

lo Sheridan, P.F° and Smith, R°P°, "Interactional Aerodynamics:

A New Challenge to Helicopter Technology', Presented at the

35th Annual Forum of the AHS, 1979.

° Landgrebe, A.J., Moffit, R.C. and Clark, D.R., "Aerodynamic

Technology for Advanced Rotorcraft', Paper Presented at the

AHS Symposium, Essington, PA, 1976.

3. Wilson, J.C° and Mineck, R.£., "Wind Tunnel Investigation of

Helicopter-Rotor Wake Effects on Three Helicopter Fuselage

Models', NASA T_, 1975.

°

°

Freeman, C.E. and Xineck, R.E., "Fuselage Surface Pressure

Measurements of a Helicopter Wind Tunnel Model with a 3.15-

Meter Diameter Single Rotor', NASA TM-80051, 1979.

Freeman, C.E. and Wilson, J.C., "Rotor-Body Interference

(ROBIN) Analysis and Test", Paper Presented at the 36th

Annual National Forum, AHS, Washington, D.C., May 1980.

6. Betzina, M.D. and Shinoda, P., "Aerodynamic Interaction between a 1/6-Scale Helicopter Rotor and a Body of Revolution", NASA y_, Ames Research Center, Moffett Field,

CA, June 1982.

7. Arcidiancano, P.J. and Sopher, R., "A Review of Rotor Loads

Prediction Methods', AGARD Conference Preprint AGARD-CP-334-

14, Fluid Dynamics Specialists Meeting, London, May 1982.

°

°

Smith, R°V°, "Some Effects of Wake Distortion due to a Fuselage Flowfield on Rotor Thrust Limits', Paper Presented at

ARO Workshop on Rotor Wake Technology, Raleigh, N.C., 1979.

Young, C., "Development of the Vortex Ring Model and its Influence on the Prediction of Rotor Loads', AGARD Conference

Preprint, AGARD-CPP-334-11, Fluid Dynamics Specialists

Meeting, London, May 1982.

I0. Jepson, D., Moffitt, R., Hilzinger, K. and Bissell, J.,

"Analysis and Correlation of Test Data from an Advanced

Technology Rotor System', NASA _,_c.l_, July 1980.

Ii. Huber, H. and Polz, G., "Studies of Blade to Blade and

Rotor-Fuselage-Tail Interferences', AGARD Conference Reprint, May 1982.

12. Stricker, R. and Polz, G., "Calculation of the Viscous Flow

around Helicopter Bodies", Paper Presented at the 3rd

European Rotorcraft and Powered Lift Forum, Aix-en-Provence,

France, September 1977.

51
13. Clark, D.R., Dvorak, F.A. et al., "Helicopter Flow Field

Analysis", Final Report for Period April 1977 - September

1978, USARTL-RT-79-4, Ft. Eustis, VA 1979.

14. Clark, D.R., "A Study of the Effect of Aft Fuselage Shape on

Helicopter Drag", Paper Presented at the 6th European Rotorcraft and Powered Lift Aircraft Forum, Bristol, England,

September 1980.

15. Freeman, C.E., "Development and Validation of a Combined

Rotor-Fuselage Induced Flowfield Computational Method', NASA

TP-1656, 1980.

16. Heyson, H.H. and Katzoff, S., "Induced Velocities near a

Lifting Rotor with Nonuniform Disk Loading", NACA Report

1319, 1957.

i

17. Clark, D.P., "An Outline of a Method for Predicting Fully

Coupled Body/Rotor Interferenco", Paper Presented at the ARO

Workshop on Rotor Wake Technology", Raleigh, N.C., 1979.

18. Clark, D.R. and Maskew, B., "An Analysis of Airframe/Rotor

Interference in Forward Flight", Paper Presented to the 7th

European Rotorcraft Powered Lift Aircraft Forum, GarmischPartenkirchen, FRG, September 1981.

19. Maskew, B. and Dvorak, F.A., "Analysis of Vortex/Surface

Interactions using Panel Methods", Paper Presented at the

1981 U.S. Air Force - FRG DEA Meeting, Gottingen, West

Germany, April 1981.

20. Maskew, B., "Prediction of Subsonic Aerodynamic Characteristics--A Case for Low-Order Panel Methods", AIAA I.

_, Vol. 19, No. 2, February 1982.

21. Landgrebe, A.J. and CheynEy M.C., Jr., "Rotor Wakes--Key to

Performance Prediction", AGARD Conference Proceedings No.

111, Marseilles, France, September 1972.

22. Clark, D.R., "Use of Analytic Tools in the Design and

Development of Rotorcraft', Paper No. I, Presented at the

4th European Rotorcraft and Powered Lift Aircraft Forum,

Stresa, Italy, September 1978.

23. Maskew, Bo, "Program VSAERO, A Computer Program for

Calculating the Nonlinear Aerodynamic Characteristics of

Arbitrary Configurations. User's Manual", Prepared for NASA

Ames Research Center, Moffett Field, CA, under Contract

NAS2-8788, April 1982.

24. Roesch, P. and Vuillet, A., "New Designs for Improved

Aerodynamic Stability on Recent Aerospatiale Helicopters",

Paper Presented at the 37th Annual Forum of the AHS, New

Orleans, LA, 1981.

P

_ 52
APPENDIX 1: TEST DATA CORRELATION

TEST DATA DRAWN FROM REFERENCE 4

DATA IS PRESENTE_ FOR CASES WHERE CTE 0,005

ADVANCE RATIO 0,05: RUN 22/POINT 139

CT : 0,00518

ADVANCE RATIO 0.15: RUN 25/POINT lq8

CT : 0,00500

LOWER SURFACE (DOWNWARD FACING)_ ANALYSIS

UPPER SURFACE (UPWARD FACING)

• TEST DATA

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CR- 177340 .....

4. Title aml Ilulmle

Study for Prediction of Rotor/Wake/Fuselage

Interference

Part I: Technlcal Revort

7. Au_Id_

D. R. Clark

B. Naekew

s. _,q.,..,,,,,oo._..,_.eo. _ ,.d

Analytlcal Methods, Inc.

2047- 152nd Avenue, N.E.

Redmond, WA 98052

'I?. _pomo_nql ,_ Nmm a_l J_Id.l_

NASA Ames Research Center

Hoffett Field, CA 94035

I_ Sul_t.mmutuV Notes

Point of Contact:

L _ rb..

March 1985

L ,'mm_ o,s,.tm_ cm.

..

8304-1

I_ Wer_ Uet No.

11. Ce_aet or Gnu_ Mo.

NAS2-I0620

13. Type of liieDon Ind Period Covaed

Final, 611180-1111183

14. sm._tw _ ceo

505-42-11

Technical Monitor, Charles A. Smith

M/S 247-1. NASA Ames Research Center

Moffett Field, CA 94035 (415)694-6714

16 A_aCt

A method has been developed whlch allows the fully coupled calculatlon of

fuselage and rotor alrloads for typlcal helicopter configurations In forward

flight. To do this, an Iteratlve solution is carried out based on a conventional panel representation of the fuselage and a blade element rep_sentatlon of the rotor where fuselage and rotor slngularlty strengths are

determined simultaneously at each step and the rotor wake Is allowed to relax

(deform) in response to changes in rotor wake loadlng and fuselage presence.

On completion of the iteration, rotor loadlng and Inflow, fuselage slngularlty

strength (and, hence, pressure and velocity distributions) and rotor wake are

all consistent.

The results of a fully coupled calculatlon of the flow around representative

hellcopter configurations are presented. The effect of fuselage coNponents on

the rotor flow fleld and the overall wake structure is detailed and the aerodynamic interference between the different parts of the aircraft is discussed.

Tn particular, the flow field developed by the rotor head is followed and the

effect of a rotor head cap and pylon modifications in redirecting the rotor

head flow are illustrated. Good correlation between measured and calculated

fuselage airloads in low-speed flight is achieved and correspondence with

observed flow field behavior is demonstrated.

17. K_ Wwm |Sutllmd _ Aullwld)

Rotor Air Loads

Fuselage Rotor Loads

Helicopter

Aerodynamic Interference

19. Sl¢_ Omit iM I_ ,_t!

UNCLASSIFIED

14. O_tril_IGn Sqmlemem

Un I Imi ted

Subject category 01

20. kCwi_.,, ¢lmif. Iof It_ IIIII} 21. No. o_ PaPs 2"_. Iq_,e*

UHCLASSI FI El) 82

*For sl_ by the NstmmH 'l'_'tm_d Int_m'mli_m Ik,_m. _, V_ TJI01