Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc. AIAA Meeting Papers on Disc, April 1996, pp. 492-501 A9630777, AIAA Paper 96-1221 Active control of helicopter blade stall Khanh Nguyen NASA, Ames Research Center, Moffett Field, CA AIAA, Dynamics Specialists Conference, Salt Lake City, UT, Apr. 18, 19, 1996 This paper describes the numerical analysis of an automatic stall suppression system for helicopters. The analysis employs a FEM and includes unsteady aerodynamic effects (dynamic stall) and a nonuniform inflow model. The stall suppression system, based on a transfer matrix approach, uses blade root actuation to suppress stall directly. The results show that stall can effectively be suppressed using higher harmonic blade root pitch at both cruise and high speed flight conditions. The control amplitude was small, less than 1 deg. In a high thrust, low speed flight condition, stall is fairly insensitive to higher harmonic inputs. In general, stall suppression does not guarantee performance improvements. The results also show the distinction between stall suppression and performance improvement with active control. When the controller aims to reduce the shaft torque, rotor performance improvement can be achieved with a small degradation in stall behavior. (Author) Page 1 A96-30777 ACTIVE CONTROL OF HELICOPTER BLADE STALL Khanh Nguyen* NASA Ames Research Center, Moffett FieldA C , forward flight blada , e encounter sdifferena t dynamic Abstract pressure due to the combination of blade rotation and ro- tor translation speed. Thus, the dynamic pressure is This paper describes the numerical analysis of an auto- greateadvancine th n ro g sid eretreatine thath n no g side. matic stall suppression system for helicopters. The r rolFo l moment balance blade th , e operate t anglea s f so analysis employs a finite element method and includes attack that are low on the advancing side and high on the unsteady aerodynamic effects (dynamica stall d an ) retreating side. At high blade loading or at high forward nonuniform inflow model. The stall suppression sys- speeds locae th , l blade section angl attacf eo becomn kca e tem, based on a transfer matrix approach, uses blade root large enough to stall. For untwisted blades, the stall area actuation to suppress stall directly. The results show occurs nea blade rth e tip, growing inboar loadine th s da g that stal effectiveln ca l suppressee b y d using higher har- forware oth r d speed increases r twiste[1]Fo . d blades, monic blade root pitch at both cruise and high speed effecte th reversee sar d stale —th l area spreads froe mth flight conditions controThe . l amplitud smallewas , less blade root outboard. than 1 deg. In a high thrust, low speed flight condition, stal s fairli l y insensitiv higheo et r harmonic inputsn I . Operating in an unsteady environment, the most severe general, stall suppression does not guarantee performance typ f staleo l encountere rotoa y b dr blad s dynamii e c improvements. The results also show the distinction be- stall forwarn I . d flight blade th , e experiences time-vary- tween stall suppressio performancd nan e improvement ing dynamic pressure and angle of attack changes arising with active control. Whe controllee nth r aim reduco st e from blade pitch inputs, blade elastic response nond an , - the shaft torque, rotor performance improvement can be uniform rotor inflow. If supercritical flow develops un- achieved with a small degradation in stall behavior. dynamir de c conditions, then dynamic stal initiates i l y db leading edge or shock-induced separation. Supercritical Introduction flow is associated with the bursting of the separation bubbl bubble th s ea e encounters tiie large adverse pres- Suppression of retreating blade stall has been proposed as sure gradient near the blade leading edge [2]. Dynamic a means of helicopter flight envelope expansion, thereby stal characterizes i l sheddine th y b d f stron go g vortices enhancin utilite gth f thesyo e aircraft. Unlike fixed-wing fro leadine mth g edge region leadine Th . g edge vortex aircraft, stall doe t limi loe no s wth t speed operatiof no produces a large pressure wave moving aft on the airfoil helicopters. Stal roton o l r blades, however, limite sth upper surface and creating abrupt changes in the flow helicopter maximum speed as well as the loading capa- field. The pressure wave also contributes to large lift and bilities. Stall places a loading limit on most of the heli- moment overshoots in excess of static values and signif- copter flight envelope at low and medium speed, and at icant nonlinear hysteresi airfoie th n si l behavior. high speed, either stal r compressibilito l y effectn ca s limit helicopter operations. A rotor experiencing stall e otheTh r typ f staleo l typically encountere rotoy db r can require more shaft power than is available from the blades involves trailing edge separation. The phe- engine. Also, the excessive control loads on a stalled ro- nomenon of trailing edge separation is associated with ei- tor blade, together with the changes in blade aerodynamic ther static or dynamic conditions. Separation starts from behavior, adversely affect aircraft handling qualities. the airfoil trailing edge, and with increasing angle of at- Stall-induced loads, possibl combination yi n with blade tack, the separation point progresses towards the leading dynamics as in stall flutter, can severely damage blade edge region. Trailing edge separation contribute nono st - structural components and cause excessive cabin vibra- linear behavior, suc hysteresiss ha liftn i , , dra pitchd gan - tion. in glose momen circulationth n s i o t e tdu contrasn I . o t t dynamic stall tha characterizes i t abrupy db t changen i s A unique characteristi f helicopteco roccure stalth s i l - airfoil behavior, trailing edge stall progresse moder-a t sa renc retreatine th f stal eo n o l rotoge sidth f r eo disk n I . ate rate. Passive control of blade stall typically involves the tai- loring of blade twist and planform for efficient blade load •Aerospace Engineer distribution. Another method employs blade construc- Copyright © American Institute of Aero- tion with multi-airfoil sections — thick, high-lift sec- nautic d Astronauticsan s , Inc., 1996H A . tions inboard and thin, transonic sections for the tip re- rights reserved. gion. These method provido t m sai e efficient rotor disk computed transfer functions relating higher harmonic draw loadinthusd lo gan d , employegan d primarilr yfo contro changeo t l bladn si e angl f attackeo , Arcidiacono performance benefits; however, they also provide stal- al l derived a blade pitch schedule that approximated an "ideal leviation. schedule" for stall alleviation. The analysis showed that the blade pitch schedule, which includeP 3 d botan P h2 As an alternative to passive methods, active control of components with a combined maximum amplitude of blade pitch has the potential to alleviate blade stall. deg3 capabls 4. wa , f avoidineo g retreating blade stall. Recent developmen high-frequencyf o t , blade-mounte- dac resultine Th g effects could rais speee eth d limihelia f o t- tuator ] make[3 s s this concept feasible e operatinTh . g coptepercen0 3 y b r t ove baseline th r e maximum speed. frequencies for blade pitch control are not limited by the additionae Th l power requiremen speee th do t gaie du tn blade-integer harmonics swashplatn i s a , e oscillationt bu , would be large, however, to compensate for the increase e bandwidtbth y e actuatorsth f o h . RecentlyF Z , in fuselage and rotor profile drag (about 100 percent). Luftfahrttechnik, GmbH of Germany built and wind tun- nel tested, together with NASA Ames Research Center, In 1961, a flight test program was conducted to investi- an individual-blade-control system on a full-scale BO-105 gat feasibilite eth f usinyo g higher harmonic contron o l rotor. These actuators were tested at harmonics from 2P an UH-1A helicopter [7]. Using a rotor head mechanism to 6P (42.5 Hz) and amplitudes up to 3 deg. Although capable f generatino bladP g2 e pitch, Bell Helicopter no stall suppression study was attempted, the benefits of conducted a series of flight tests to determine the effects IBC input on rotor performance at high forward speed of active contro roton o l r performanc loadsd ean . Test (advance ratio p. = 0.4) were encouraging [3]. results indicated that the 2P control at different ampli- tudes and phases did not produce any reduction in rotor Previous Work shaft torque. Determined to resolve this variance with theoretical prediction investigatore ,th s conducte dposta - In 1952, Stewar ] suggeste[4 t d per-rethao tw t v (2P) test analysis. Analytical results indicated that the drag blade pitch applied to rotors in forward flight could be reduction in the retreating side due to 2P control was off- use delao dt onsee yth retreatinf o t g blade stall. Basen do set by an increase in profile drag in the fore and aft por- the analysis that included a rigid flapping blade, quasi- tions of the rotor disk. Such conclusions confirmed pre- steady aerodynamics and uniform inflow models, Stewart vious analytical predictions that 2P control could reshape derived an approximate transfer function relating the the rotor disk loading. change in 2P blade angle of attack due to 2P control. Results indicated that rotor disk loading coul effie db - Kretz [8, 9] reported the wind tunnel test results of a ciently re-distributed using higher harmonic Wade pitch. Stall Barrier Feedback (SBF) syste six-fooa n mo t diame- For a particular flight condition, the loading redistribu- ter two-bladed rotor.