Current European research activities in helicopter interactional aerodynamics

ERF91-35

SEVENTEENTH EUROPEAN ROTORCRAFT FORUM

Paper No. 91 - 35

CURRENT EUROPEAN RESEARCH ACTIVITIES IN

HELICOPTER INTERACTIONAL AERODYNAMICS

G. PAGNANO, A. SAPORITI

AGUSTA S.p.A. CASCINA COSTA, ITALY

SEPTEMBER 24 - 26, 1991

Berlin, Germany

Deutsche Gesellschaft ftir Luft- und Raumfahrt e.V. (DGLR) Godesberger Allee 70, 5300 Bonn 2, Germany

CURRENT EUROPEAN RESEARCH ACTIVITIES IN HELICOPTER INTERACTIONAL AERODYNAMICS

G. Pagnano, A. Saporiti AGOSTA S.p;A., ITALY

ERF91-35

An Ellropean mllaborative prcxp:anm:: for Interactional Aerodynamics study on helicopter configuration was started in 1990 in the BRITE/EllRAM activities under the name SCIA (_stujy and ~ t i o n of }nteractional ~ c s ) •

'1he involved Partners are:

Agusta Aercspatiale MBB Westland Researd.t Centers Alfapi . CIRA DI.R NIR oomA Unive:csi.ties Bristol Univ. 'IU-Braunschweig 'IU-Denmark UNIBW Munchen Rane Univ.

In this paper, the d.evelcpnent and the current results of this research

are presented.

'1he major goal of the participants in the SCIA project is to inprgye

existirg nethodolc:qies for in::lividual helicopter cx:mp:>nents (rotor, fuselage)

and to develop algorithms for predictirg the canplex rotor/fuselage interaction problem. '1he efforts in this project are directed toward code/technology develq;:nent for an irrlustrial envirorment rather than p.ire

research ?Jrp:)SeS.

What follc:,.,s

is

an overview of the activities performed by the mllaborati ve groop in the ccmp.rtational area and also in the experimental field, the latter includErl. in the program because a sourrl data base is an essential reg:uirenent for aII'f programne involvirg the develcpnent and

inprovement of theoretical algorithms; the partners agreed on setting up of experimental activities to obtain basic information on rotor/fuselage flow fields, not yet available or specific to the planned activities.

1. Int::ra:b:::ticn

'!he flCM aroorxi irxlividual helicopter cacponents such as main rotor, fuselage, tail .rotor etc. is cx:.nplex. Sate of the basic flow phencmena are

described in [1]. In aerodynamic isolation, each ccmponent exhibits an unique

flCM field an::i characteristics. In an integrated system the resultant flCM is the sum of the interactin;J irxlividual flCM fields. A number of charges - both

favorable an::i lmfavorable - may develop in the c:naracteristics of the

i.rdividual caip:>nents an::i consequently alter the behavior of the system as a 'Whole.

'!he relative nation between rotatin;J an::i non-rotatin;J caip:>nents of the

h e l i ~ i.rdJD?S primarily lCM frequency unsteadiness in the flCM. Mid.

frequercy unsteadiness is produced, for exanple, durin;J interactioo of each main rotor blade with the wake of preoed:i.rg blades. High ·frequency unsteadiness is generated by blade/tip vortex erx:nmters.

Con.siderin;J the role of the fuselage, its displacerrent effect distorts the onset flCM resultin;J in a non-mu.form an:,;Jle of attack distrib.Itioo in the

rotor disk. '!he rotor wake is defornei due to the presence of the fuselage

whidl in

turn alters the unsteady interaction p.ro:ess between rotor blades an::i

their wake. 'lbe fuselage can i.muce high vertical velocities into the forward area of the rotor

disc,

particularly

if

the rotor plane is relatively close to

the fuselage. 'Ibis uptlaSh affects local airfoil loadin;J an::i, in cases where

the blade is

near

stall, can irxiuce local blade stallin;J with consequent urnesirable vibratory effects an::i increase of rotor power required. Another effect of

this

localized uptlaSh is to nove the tip vortex path upwards.

'Ihis can further aggravate the localized aerofoil problem (as the forward blade intersects the vortex path gradually) an::i can significantly affect the

blade/vortex interaction }i1enanerlon where the blade intersects the vortex abruptly causin;J acalStic problems.

'!he fuselage is imne.rsed in the main rotor downwash whose near wake region irduces unsteady airloads at a frequercy equal to an::i higher than the

number of blade passages per revolution.

With transition fran hover to cruise flight, the primary interaction area of rotor wake

shifts

fran the fuselage to the tail rotor an::i the enpennage. An

in:lirect rotor /fuselage interaction develops due to distortion of rotor hub wake an::i powerplant exhaust plume by rotor downwash with ensui.rg charges in fuselage flCM field.

In the rear fuselage region, the main rotor an::i the hub wakes are merged with the tail rotor wake: this leads to an extrelrely complex unsteady flCM field affectin;J tail rotor an::i ell'{)el1l'lage.

For many years, the vario.JS flow phencmena associated with helicopters have been addressed with a mixture of si.nplified linear aercdynamic theories, wind tunnel data, an::i design charts. But recently, analytical an::i experimental investigations have been oon:iucted, in

which

the interactions between a rotor

Sheridan and Smith [2]. offered a

review

of interactional aerodynamics and presented the results obtained testing different rotor/airframe mdels. Wilby

et al. [3] ptq,osed an investigatial of the fuselage effects al rotor

performance, by using predictial methods and flight test experiments. Rand and

Gessow [4] calculated the fuselage effects al main rotor. Freeman [5] applied a coupling methodology in order to predict the rotor/fuselage flow field.

McMahon et al. [6] and I.eishman et. al. [7] used IX)tential · flow methods for calculating the unsteady pressure al a silrplif ied fuselage. Dedicated· experimental activities have been also cxniucted in this field: isolated rotor

in grourrl effect, Light [8], silrplified i.nteractional m:xiels (Kanerath et al.

[9]), scaled powered m:xlels( I.eishman et al. [7, 10, 11], Hoad et al. [12]),

full scaled m:xlels (Nannan and Yamaudli [13]).

starting with the state of the art provided by the plblished material and considering the different levels of contrib.rt:ion and capabilities among the partners, a program was agreed with sharing of activities and resp::,nsibilities; the cxx:>rdination of the program is in charge to

Agusta.

'lh.e general scc:pe of the ptogram is the inprovement of methodologies and

prediction methods for i.nteractianal aerodynamics, in order to provide

benefits in the helicopter design phase, in particular for: - prediction of rotor performance and loads

- reduction in power cansunption

- reduction of pitch-link loads

- prediction of rotor wake develcpnent for blade vortex - interaction and IX>ise analysis

- prediction of fuselage aerodynamic loads - reduction of vibration level

2. Experilllental activities

Information is provided on the isolated rotor, in hover and forward flight, and on the powered m:xiels inclu:ii.n;J flow visualization, usin;J a laser

a.It technique, and measurem:mts, usin;J Hot Wire and three caxq:xment IIJV.

'lllese data bases are designed to provide the following - Aerodynamic rotor forces

- Aerodynamic fuselage forces - Fuselage pressure distrib..rtion.s

- Inflow velocity distrib..rtion near rotor

- Isolated rotor tests in hover has been CX>rrlucted at the

Agusta

low speed facility. 'lh.e rotor mdel is Madl scaled, fully articulated with four ~ blades and with a diameter of 1.5 m (Figure la). 'lh.e basic tests

provide a survey of the external flow field at different thrust and RPM: Hot

Wire measurements of the imuced velocity (Figm:e lb), flow visualizatial by means of laser a.It tedmique, this in order to

assess

the hover numerical methods.

- Tests on a

Da.1.JP'Uil

powered m::xiel. (scaled 1/7.7) have been perfarmed in t h e ~ S2 Olalais-Meudon wini tunnel. Unsteady velocity field measurements

aroum

the fuselage and the rotor in two different transversal planes

(perpen:licular to the fuselage axis) , usirg a three-dilrensional Laser Dqpler

Velocimeter, have been dcne. .

'!he fuselage leJ'l3tl} is about 1.5 m, the rotor diameter is 1.5 .m and_ the

rotatirg tip speed is 100 m/s. A view' of · the noiel in.side the S20l-wini tunnel is presented in Figure 2a. '!he configuration studied corresponds to a sim.ll.ated mass of 4000 Kg, at an advarx::e ratio of

o.

2.

'!he measurements

are

perfarmed

every

4

°

in azinuth; this means 90

different slots to

describe one

rotor revolution,

each

slot has a width of 1°

in azim.rt:h and far eadl slot the measurements are· averaged over 100 sanples. '!he instantaneaJs

vertical

velocities far different points located slightly above the rotor disk show that art:side the rotor

disc

(Y/R

>

1)

the velocity fluctuatioos are very small and the mean value positive (upwash) lffii.lst inside the rotor disk (y/R < 1, Figure 2b) the fluctuatioos due to the blades passage are inpart:ant an:l the mean value negative (downwash).

- Force and pressure measurements far isolated fuselage at different onset velocities and incidence and yaw argles have been also performed oo a

00105 scale noiel in the low speed wi.rxi tunnel of DI.Rat Braunschweig.

With the rotor on configuration, powered nooel, the time averaged pressure distril::ution on fuselage surface (450 points) was measured far different thrust ratios and incidence: the Figure 3 shc::MS clearly the effect of the rotor downwash on

the

fuselage.

All these measurements (velocities, pressures and forces) constitute a

gcxxi data base for code validation in the BRITE/EX.JRAM group workirg on the helicopter interactional aerodynamics.

3. 'Jhecretical lllO[X

'Iheoretical work has been divided into three broad categories. 4 .1 :rsol.ated and tn:Xlllpled rcbr

In

this task different methods are investigated in order to assess their capabilities an:l to introduce sane significant inprovements in the case of both isolated rotor (without fuselage effects nooelled) and rotor with fuselage irrluced velocity field IOOdelled.

Isolated rotor

An isolated rotor code, based oo the liftirg surface

awroach,

[14], has

been ai;:plied in order to calculate (Figure 4) and to canpare the imuoed

velocities below the rotor disc of the isolated rotor nooel in hover O.G.E. at different planes and azim.rt:h. '!he

prescribed

wake nooels (Iandgrebe, Kocurek)

and the successive relaxations have been investigated. . Especially for the case of axial flight, a higher order mxlified liftin:;J surface method is ai;:plied in order to inprove the efficiency of the rotor flow field and airloads calculations.

A c:x::mparison of the calculated flow field for the isolated Dauphin rotor model with oomA wind tunnel measurements is shown in Figure 5: although the

~imental results are influenced by the fuselage of the powered model, the

agreement seems to be satisfactory. · 'Ihe rotor flow has been also calculated · using an unsteady vortex lattice met.hod with free wake fonnulation (19]. Rotor

and wake are represented by a distribution of vortex-doublet elements with

stepl,Tise constant strength. .

Moving the rotor blade (rotation), the vortex activity at blade's trailing edge sheds into the wake during each time step.

Boundary conditions at blade cxartrol points tmlSt be satisfied by doublet

strength of the lifting surfaces. other collocation points, like wake points,

are

m:m.rg

free tmde.r the influence of free stream and all existing

singularities. In that way, geanetrical wake up1ates require a special procedure, inclu:iing a near field method for the Biot-Savart law, to achieve a real wake behavior. Figure 6 presents a panelized 4-bladed rotor with a free vortex lattice wake after one of the blades. 'Ihe self-induced outer wake roll up, traced by the blade tip, is clearly visible.

A Boundary Element Met.hcxl (BEM) has been developed for the aerodynamic analysis of an isolated rotor in foi:ward flight for incanpressible and

sul:sonic c:x:rrpressible flow

regimes.

'!he met.hod solves an integral equation for

the velocity potential that is obtained by applying the generalized Green's function met.hod to the linearized govern.in;J equation for the velocity potential [20, 21].

'Ibis kind of approach requires an explicit treatment of the wake that therefore m.ist be modelled. Presently prescribed-wake georretry traced rut by

the path of the blade trailing edge durinJ its rrotion is used; a doublet layer, fully equivalent to a vortex layer, is distrib..lted on this prescribed surface. 'Ihe blades are represented by a source and doublet distribution. Both

blades and wake are discretized by zerot.h-order panels.

A first kind of validation of the met.hod has been performed by the

calculation of the vertical c:x:m:ponent of the induced velocity on a plane parallel to the tip path plane (Figure 7) for the NASA test case of SCIA

ccmm:>n exercise [12].

'!he results have been ccmpared both with the ~imental data and with

the rrumerical results of other partners: a good agreement was find rut in both cases.

'!he BEM methodology has been further developed [22, 23] and rrumerical investigations have been performed for isolated rotor in hover and foi:ward flight in c:x:rrpressible f l ~ . Free-wake results, Figure 8, for sul:sonic f l ~ , usinJ five-spiral wakes and a 12x12 element discretization over the blades,

are ccmpared with the ~imental data by caradonna and Tung (24] in hover

(Mtip=O. 727). '!hen, rrumerical results for forward flight rotors in sul:sonic f l ~ (Mtip=0.54,µ=0.17), usinJ 12 time step per revolution, 2 spiral wake and

a 4X9 element discretization over the blade, are compared, Figure 9, with those of

Tai

and Runyan [25].

Isolated rotor with fuselage effects

A ccnp.rt:ational efficient mcdel has been then developed, in order to

predict the effects of the fuselage on the rotor inflow arrl herx::e an rotor performanoe. '!he mathematical rotor model [15) is based on 2-0 strip aerodynamics arrl steady state rotor blade dynamics usin;J only cut-of-plane

ben:.tirg mcde shape, suitable for various types of rotor articulaticn.

originally the rotor in:inced velocity was calculated accord.m] to the

theory of Margler arrl Squire, expressin;J downwash as a Fourier series; the

code has been :m:xlified in such a way that an external flow field (e.g. fuselage i.rxiucm) can be superinpnsed.

In Figure 10,

oontaur

plots of the calculated inflow ratio for the

Margler-Squire mcdel with (right) an:l witha.It (left) the influence of the

fuselage are carpared (test case [12]) • 'lhese results show that fuselage upwash disturbs the in:iuoed velocity field mainly alon;J the lon;Jituiinal axis near the rotor hub, an:l the velocity dist.rirutions cx,rrelate well with regard to exper.unental data

trem.

A carp3rative analysis of the ccnp.rt:ational results shows that the

disturbed flow field, for the configuration urrler study, influences rotor blade fl.al=Pin;J · arrl rotor loads. Rotor torque coefficient predictions an:l carparison with experimental results, at the same q>eratin;J cx,n::litions, show a better agreement carpared to results of the original code withcut fuselage

inflow. ·

Another a.wroach used for isolated rotor is a liftin;J line method 'Where

the rotor wake is discretized ~ a lattice linear vortex segnents [16) whose intensity is related to the variation in circulation span an:l azi.mrt:h. wise.

Once the marginal vortex has rolled up

(Betz

theory), this lattice is reduced

to a tip arrl root vortex formirg the far wake. '!he wake geometry is prescribed with the conventional cycloidal trajectories (enpirical fo:anulae

inspired

fran Egolf arrl I..arx:igrebe work); the in:iuced velocity at each cx,ntrol point on the

blade is calplted ~ the Biot an:l Savart law an:l the lift is obtained trc:u;;Jh

2D airfoil tables.

Blade m::rt:ion is calculated sinultaneously ~ cx,nsiderin;J rigid blades an:l

hinJed

in flap only: this aerodynamics/dynamics problem is solved iteratively

with a relaxation method where in:iuced velocities are the unknowns.

'!he next step is the addin3' to this isolated m:xiel the effect of the

fuselage inflow obtained ~ a low order panel method [17], with ex>nstant

sa.iroe an:l doublet distrirution. '!he CCllp.lted results obtained for: a) the

isolated rotor an:l b) for the same with the effect of the velocity field catp.Ited for an isolated fuselage, are carpared with exper.unental results

[12): a better agreement can be ooticed when the fuselage effect is taken into

acooont, in particular an the inboard part of the blade (Figure 11).

'!he effect of fuselage on rotor is usually calculated~ panel methods,

rut in parallel, a ccnp.rt:ationally efficient method for determininJ fuselage displacenent flow effects on a main

rotor

has been developed. '!he necessity for ccnp.rt:ationally efficiency is because fuselage effects form only a small element in the ccnp.rt:ation of rotor perfonnarx:e within a comprehensive

rotor

analysis, which nust recognize all significant plysical J;ilenc:mena· to a cx>nSistent degree of awroxilnation within the overall calculations.

'!he mcdel for the fuselage displacerrent flow effects is bein;J develcprl

'!he work to date

has

concentrate:i on two aspects:

&

the developnent of a

distributed

scurce/sink irodel for approximating the

displacement flow of a fuselage,

121

the developnent of a sinple method. for m:xiellin:J cross-flow effects arr a

fuselage. .

Initial correlations of the

distrib.Ited.

source/sink representatioo for the SCIA

test case [12] showed that

the sinple m:xiel can adequately reproduce the inflow perturbatioo · at the rotor plane predicted by panel method. '!he sillple model facilitates the calculation of wake element vertical displacements due to the fuselage flow field, which can be inportant 'When

calc::ulatirg the local incidence due to ~ices in the flow field, Figure 12. A cross-flow m:xiel for the fuselage is currently urrler develcpnent (to be validate:i by

carparisoo

with the panel method), again with the primary

am

of developing a sinplified m:xiel which can be used within canplex flow fields with high c:xmp.rtational efficiency.

In this task c:xmp.rtational activities are perforne:i in applying arrl

inprovirg UllCCAlpled-approach methods for the calculations of the flow field in

the case of both isolate:i fuselage arrl fuselage in presence of prescribed rotor dow:nwash.

Isolate:i

fuselage

A panel method [lB] has been applied to the SCIA reference fuselage

test

case [5], in order to analyze the effects of flow separations on surface pressure distril:ution arrl in order to harrlle the fuselage wake irodels.

'!he potential no-wake solution shows the presence of an unrealistic stagnation point on the rear side of the pylon, Figure 13. By applying a two-dimensional l::aJrrlary layer analysis on the calculate:i streamlines, a

vi.scx:AJs/

inviscid ca.1pling highlights a separation area on the rear of the

pylon: a wake has been then m:xielled arrl attached at the calculate:i separation line (dooblet panels with prescribed vorticity gradient), an:i the solutioo was inproved.

'!he pylon wake has been also rn.merically simulate:i by a set of interactirg vortex filaments, Figure 14, then, a three dimensional l::aJrrlary

layer axle c:x:mprting laminar, transition arrl turb..llent state has been also applied to that configuration in order to inprove the separatioo line prediction.

Isolate:i fuselage with prescribed rotor downwash

An :iJiportant part of the research into rotor /fuselage interactions is to develop methods for predictirg the pressure over the fuselage in the presence of rotor dow:nwash. 'llris again uses panel methods to represent the fuselage,

rut

-raw the rotor arrl its wake are modelled by a vortex system that m:xlifies the flow at the fuselage.

A typical result for the pressure distril:ution over the

uwer

fuselage

center

line for low advance ratio is shown in the Figure 15. 'Ille dotted o.n:ve

shows the pressure distril:ution predicted using a constant total head for the

flow (consistent with the increased induced velocities irrluced by the rotor

'!his is seen to markedly under-estimate the pressure loads on the

uwer

fuselage when cmpared with experimental data. '!his is because the

rotor

vortex m:x:lel, althcujl correctly m:x:lellin;J the velocity field in:iuced by the rotor, does not represent the in::rement in total head irrl.uced by rotor systen.

'1he solid curve shows the result of in::lulirq an empirical con:ectioo to

the total head. '!his is the lr.'eakest part -of the theoretical method, especially at l<Jfll advance ratios, an:i required further develqment. In particular it is necessary to identify a means of iooorporatin;J increments in total head in the

rotor l!Ddel, an:i then tradcin;J the regions of flow that contain these . .in::renents.

3 .3 Qq>]ed rr:b:r/flJSPJage calDJlatims

In this task methodologies are bein;J develcped to obtain a rotor/fuselage cx:np.rt:atiooal method by the canbination of the distinct rotor an:i fuselage fl<Jfllfield prediction algorithms ai:plied in the previous tasks an:i by usin;J global aerodynamic

awroac:t:ies.

A cx:np.rt:ational method is un:Jer develqment by couplin;J the isolated fuselage an:i the isolated rotor methods. '!he couplin;J technique is thrcujl a time mardl.irg

awroacn:

a each azina.rth step a fuselage calculation is

performed with the velocity perturbation due to the rotor, a.rrl then the

velocities due to the fuselage are cx:rrp.rted at the level of the rotor a.rrl rotor wake to be taken into ac:x::omt at next time step.

F\n:thernore a

preliminary

work for the develqment a.rrl the awlication of global interactional aerodynamic m:x:lels will be umertaken. '!he

awroacn

described in [18], VSAER), is bein;J also ai:plied to a powered m:xiel in order

to predict the in:iuced inflow 1-chard above the tip path plane of the rotor

disc [12). Usin;J a blade element m:xiel of the rotor, the program calculates

the radial an:i azina.rth variation of blade loads an:i c:xmverts these in order to · calculate the normal velocity 1:::nJniaey corxlition at the doublet rotor

disc

in

the panel sdleme: then the oarplete helicq,ter c:xmfiguration is calculated by solvin;J the dooblet stren:;Jths oo the fuselage an:i on the rotor.

'!he stuiies in progress in Furope in the helicopter aerodynamic

interactions have been described. In particular, a survey of the activities in

the experimental an:i cx:np.rt:ational field, both in the isolated a.rrl in the

oarplete configuratioo, were presented.

'!his research an:i collaborative work is improving the

i:tiencmena

urrlerstand.in; a.rrl

is leaciin;J to a develqment of a dedicated prediction tools.

'1he activities described here need to be exterrled in order to analyze

other inport:ant an:i oarplex aerodynamic interactional Iilenomena, . that occur in

the helicq,ter flight envelope such as main rotor /tail rotor wakes, tail rotor

wake/enpennage, wake develqmant in gram:i effect.

Future works will in::lu:ie links with mre sophisticated C.F.D. solvers (Full-Potential, Elller, Navier-stokes) to sinulate the complex flows aroord a helicq,ter, in::lulirq. viscoos effects, a.rrl to provide nore -detailed aerodynamic forces for performance a.rrl load analysis.

'!his "'1tlrk was carried

out

with

furw.n::J

fran &Iropean &:xn:mic Ommmi.ty

under the BRITE/El.JRAM Contract No. AER:>-0011-C {A) "Helicopter Rotor/Fuselage Interactional Aerodynamics", acronym SCIA.

'!he authors woold like to ac:::knowled;Je the contri.b.ltions of

their

colleagues in the steerirg Ccmnittee of the programme: F.T.Wilson, S.Fiddes,

D. Papanikas, A. Descpper, F. Tou.l.may, G. Folz, P. Renzoni, s. Weigner, S.R.

Ahmed, F. W. Meyer, J .N. Sarensen, C. Hermans.

1. F.X. caradonna, 'lhe application of CID to rotary wjn;J fla.,, problem,

AGl\R>-FtP-VKI Special Q:Qrse on aerodynam.i.c of RDtarcra.ft, VKI, Belgium, April 1990.

2. P.F. Sheridan

am

R.P. smith, Interactional aerodynamics - A new dlallerge to helicq,t:er technology, 35th A.H.S. Fm:un, May 1979.

3. P.G. Wilby,

c.

Yoorr;J

am

J. Grant, An investigation of the influence of fuselage fla.,, field on rotor loads,

am

the effects of vehicle ex>nfiguration, 4th &Iropean Rotcrcraft Forum, Paper No 8, September 1978.

4.

o.

Rarxl

am

A. Gessow, :ftk'ldel for investigation of helicopter fuselage influerx::e on rotor fla.,rfields, Jouma.l of Ai.:rc:ra:ft, Vol. 26, No 5, May 1989.

5. c.E. Freeman, Developnent

am

validation of a combined rotor-fuselage irrluced fla.,, field CXllpltational methcxi, ~ TP 1656, June 1980.

6. H.M. M:Mahon, N.M. Kanerath.

am

D.N. Mavris, Prediction of aerodynamic rotor-airframe interactions in forward flight, Journal of the A.H.S., october 1989.

7. J .G. Leishman, N. Bi

am

G. L. Croose, 'lheoretical

am

experimental study

of unsteady rotor/}::ajy aerodynamic interaction, 46th A.H.S. F'ca:1.a, May 1990.

8. J

.s.

Light, Tip vortex gearetry of a hoverin;J helicopter rotor in grourrl effect, 45th A.H.S. Forum, May 1989.

9. N. Kaoerath, H. McMahon, A. Bram,

s.

Liou

am

D. Mavris, Prediction

am

measurenent of the aerodynamic interactions between a rotor

am

airframe

in forward flight, 45 A.H.S. Farullll, May 1989.

10. J.G. Leishman

am

N. Bi, Aerodynamic interactions between a rotor

am

a fuselage in forward flight, 45th A.H.S. Forum, May 1989.

11. J. G. Leishman

am

N. Bi, Exper:ilnental study of aerodynamic interactions

between a rotor

am

a fuselage, A1M 7th Applied Aerodynamic Q::nfei.euce,

July 1989.

12. D.R. Hoad, S.L. Althoff

am

J.W. Elliott, Rotor inflow variability with

advance ratio, 44th A.H.S., June 1988.

13. T.R. Norman

am

G.K. Yamauchi, Full-scale investigation of aerodynamic interaction between a rotor

am

fuselage, 47th A.H.S. Fm:un, May 1991.

14. Analytical Methods, Evaluation of blade tip planform effects on hover

performance, Analytical Methods Report No 7908, 1979.

15. H. Have?."d.irgs, A control m::idel for maneuverin;J flight for· application to a carp.rter flight testjn;J programme, 7th &Jropean Rotarcraft Farum, Paper

No 38, September 1981.

16. A. Dehont

am

F.

TouJ.m:w,

Inf luenc::e on rotor inflow performance

am

trim, 15th &Iropean Rotarcraft Far:um, September 1989.

17. J. Ryan, G. Falerpin arxi T .H. le, Rotor plane velocities by a heliocpter

fuselage, 2rd Int:ematirm.1. Ccmfe:ceuce m Basic Rotarc:ra:ft Res rd>.,

February 1988.

18. B. Masker.,.r, A carprt:er pz:ogtam for calculatirg the non-linear aerodynamic

characteristics

of

arbitrary o:t'lfiguration,

NAS2-11945, December 1984.

19. A. Rot:tgenna.nn, . R. Behr,

c.

Sd:lottl arxi

s.

Wagner, calculation of

blade-vortex interaction of rotary wirgs in incarpressible flow by

unsteady vortex lattice, 7th GMM seminar

on

Numerical Techniques for

Boondary Element Methoos, Kiel, January 1991.

20. L. Morino arxi K. ~ ' A general theory of unsteady m1pressible potential flows with applications to airplanes

am

rotors, Deve.lqlEnts in Bamdaey Integral. Pquat:irn &:I I o:1 Volume 6, Elsevier At:Plied Scien:::e

Publishers, Barkin:;J, UK, 1990.

21.

o.

Macina, Analisi aerodinamica agli elementi al contorno per rotori di

elicottero

in volo di

avanzamento,

Tesi

di Iaurea,

Universita

di

:Rema I.a

Sapienza,

1990.

22. L. Morino, M. Gennaretti arxi

o.

Macina, A new integral equation for potential caxpressible aerodynamics of rotors in forward flight, Intec:natirm1 ~iaJ ist•s

Heet:inJ

m aJt:an:raft Basic Research, Atlanta,

1991.

23. L. Morino, M. Gennaretti arxi P. Petl:occhi, A general theory of potential

aerodynamics with applications to heliccpter rotor-fuselage interaction, By1p:Jsi1n of the Intemat.ialal. Associatim far Boondary Element &:llcds, Kyoto, 1991.

24 • F. X. Caradonna arxi

c.

'l\1n;J, Experimental arxi analytical studies of a

m:del heliccpter rotor in hover, ts\AVIWXXII 'IR-81-A-23, 1981.

25. H.L. Rlmyan arxi H. Tai, Ccilpl.essible unsteady liftirg surface theory for a heliccpter rotor in forward flight, NASA TP 2305, 1985.

Figure la. - Isolated rotar: lllOdel

HOVERING ROTOR WAKE

X-Wire probe measurements

rpm 1500 - coll 5.9 deg - Ct/sigmo 0.075

,-... V) V C 0 a. Q)

...

:, V) 0 Q) E ->J c Q)

...

Q)

--

'6 ... >-:4:::'. u .Q Q) > "O Q) u :, "O C 0.20 0.40 0.60 0.80 1.00

r/R

- - - z/R 0.075 ...-- z/R 0.150 ____. z/R 0.225 - - - z/R 0 . .300 __... z/R 0 . .375 - - - z/R 0.450 ..., z/R 0.525 ... z/R 0.600 ... z/R 0.675 - - spline fit

I. 0. _, -z -3 -4. -6 -e 2. I 0. ·I. -2 -3 ·4 -6 -e 4 3. 2. 0. -1. -t. -3. -4. -G -e

Figure 2a - DP:p>in powered m:ldeJ •

VERTICAL VELOCITIES W

x/R =O z/R

=

0.066,. X:O X: 315mm y/R

=

0.37 0.53 0.69

2.0 1.0 .s 0.0

-.,

-t.0 a ₌ O' ; µ. = 0.15 C " Isolated F'usaloge 0

•

_CT

=

0.002 0 z _{Cy = 0.005}

-~

\ I I

. ~ l Section V1 ~ I I _{I I I}

~,

I ·\ '/'., I

.\

Top

'

front J ' ., r,.__. ~ :__,,. L.__..;

,\

/ V

u

- -.)

-

I ~

_{~'K I}

j ~

V

~

l

It'

l'i

\ !

II

f. rJI,

\I

_V

//

'

:.,-

r~, ,.,.,.,. , '1 _~~

_It

_7,

~

\7

1 3 5 7 I I I 13 1 S 1 7 19 21 2:S 25 27 PRESS.TAP NO. 45 43 41 39 37 :SS 33 :S1 29 28 •

Figure 3 - Influence of

rotor

thrust on the fuselage

Cp.

0. 1S o., 0.05 Vz / Vt ·0.05 ·O. 1 ·0.1S ·0.2 :

.T

, I

···---+-~-

- ~ · ~ , 5 ; 1 - - - f - - - 4 - - ~ I I, . ! 11:~ · t·

_w--1

1

~

i·

,, . I I

: 1.

1 .

r--

_i

, ~ ~ I ~ . - - ~ - - ___ J_ ___ ·----+----+---~---4 0.2 I ---t--- ··--·+---+----+--- - + - - - " ' I i

I

I · - - - r - · . -·--t-- - - + - - - + 0.4 0.6

•••

r/a 1.Z 1.4 ~ 1'11 • 0.0 (Lon) 01'11 • JO.O (Lon) .1'11 • 60.0 (Lon) .

____

.,.

Figure 4 - calculated induced velocity below an isolated

·rotor in hover at different blade azimuth.

~

~I.

0.00

W•F(y,z)

Verical velocities at x/R•O.O

I I I I 0.20 I I j I 0.20 coaputationa aeasureNnts

,.,

...

:;-:.

~ I j I I I j I I I I I I I j I 0.40 0.60 O.IIO 1.00 y/R I I 1.20 1 i 1 1 1 i 1 1 1 i a I I J r , T 1 0.40 0.60 O.IIO 1.00 1.20 y/R

~

_;5'1:J--..,..,,...,.., ..-, "T,..,,-, ..

~

... I "T,..,,,..,,..., "T,..,,,..-,_, TJ ..,.,.,,-..., TJ "T,..,,_,...,, I 0.00 0.20 0.40 0.60 O.IIO 1.00 1.20 y/R

~~1

_{SS"~,..,,....,..., .,., ~,..,,....,..., .,., ~,..,, ...}

~

_{, ..,.., ~,..,,....,..., .,., ~ • ...-.. , .,., ~.-.--... , I} 0.00 0.20 0.40 0.60 O.IIO 1.00 1.20 y/R ~}~51-t-°T'l"Tl-ii~jr-r-lT""T"iTj"Ti-,i~i,-j~-,-i-Ti-~l~"T~-,i-,1~,~je~i..,...,.i-,j 0.00 0.20 0.40 0.60 0.80 1.00 1.20 y/R Z/R•0.0667 &/R•O.OllJ &/R--0.040 z/R•-0.0933 &/R•-0.1467

Figure 5 - calculated vertical velocity on an isolated rotor in forward flight. Comparison with experiments.

y

Figure 6 - Vortex lattice wake of one blade at starting process.

lll!IICE M.UEI • 8, 11411E•DI ., .. -8,44DIE•D1

1 ..

1211£•01 11, IGIIE•IIZ

··--•, 1811[-IJt •, SIIIIE-01

. m.. -.

4!8(-01

. ''lit

•,419£•01 - · • 211 · t1ldthll lt.311E•IIZ -,a•,18111E•D1

Figure 7 - calculated induced velocity above TPP of an

isolated rotor in forward flight.

o

.35

.3

.25 .2 . 15 . 1

.05

classic woke free woke Caradonna-Tung woke • exp.results , , , , , , , , , , , ,'

.

"

.2

.4 , , ,

.6

, ,

.8

1

r/R

I I I 0 0 0 4

..--*

b

3.5

3

2.5

2

1.5 ~ I

.5

JI'

'

'

'

'

'

'

'

' ' 'q '

'

• PRESENT -WORK c TAI-RUNYAN

300 350

degree

Figure 9 - B.E.K •ethod: results in forward flight.

50

100

150 200 250

LAIIIDA X/C I e J

..

'

•

"I

•

11

•••••

ISDD l'fH • U H • HH • U H • DDH IUI HH n1n111r / •-ulro 0111riou11on

Ron111r / •-u1r1 u11n ru•••••• o1o1urD0nco

• I 01

••••

_{I I I}

2MRTS ROTOR CHARACTERISTICS C MU= 0.15 )

INDuCED INFLOW RATIO

µ=0.30

-Measurement azimuth : PSI m

=

180°.

• • • • • • • • • • Calculallon with fu8elaoo

v.f V • (Positive up) r/R=0.40 -.016 -.030 - 0$6 -.040 ...,101,

zo. eo. 100. UO. IIO. 220. IIO. SIO. "8.

Figure 11 - calculated effect of fuselage on the rotor flow field: comparison with experimental data •

' 81 ade Vortew ; t n t e, 'l ec t I on • BI ade St• 1 1 • t ' o l l y Ser,er1Jted I='" I o ...

...

18~ Inr.1,Je1)CP. <Alpha-Alpha0J 10.00 9.000 8. fJE\0 ; . 00(' G.000 ".000 , .nnr l. \H-H"l l \j)l\ I. 00n 90

x/R•O x/R•0.4 0.8 0.6

~I

_[

z/R•O 0.4 0.2 (-Cp) 0 -0.2

-0 ~

_{h..,_-r::}

::i-o

-~

tJ

--

...

~

,.

.

?

--

c._

.

_r • ~L- " 1111 -0.4 -0.6

R

0

_•

•o

~

0 IJD

...

_~

_o

-0.8 -1

b

-

••

r, 0 0.2 0.4 0.6 0.8 1.2 XIA x/R•0.8 _x/R•1.9

.

I

x/R-1.018

::::J

I I

l ::> .

I

_:::::::::=--·

11 IUL _DOD 1.4 1.6 1.8 2 a WAKE ONO WAKE

Figure 13 - Wake effect on the top •idline fuselage pressure distribution.

e

::,

i

Q. 0

j

I

-1

....

-, 0

.

_.

.

_..

2 .... 3

4----_..._--~-~---0 0.25 0.5 0.75 1 1.25 1.5 1.75 X/R Panel Methcx:1 with

Rotor Pressure Additions Freeman Experimental Data

Panel Methcx:I with no Rotor Pressure Addition

0

Coefficient of Rotor Thrust 0.0033 Advance Ratio 0.05

Zero Incidence and Yaw

2

Figure 15 - calculated pressure distribution with rotor

downvash on the top fuselage surface (midline).