SMART Rotor Development and Wind-Tunnel Test

SMART Rotor Development and Wind-Tunnel Test

Friedrich Straub, V.R. Anand, and Terry Birchette – The Boeing Company, Mesa, Arizona Benton H. Lau – NASA Ames Research Center, Moffett Field, California Summary Boeing and a team from Air Force, NASA, Army, Massachusetts Institute of Technology, University of California at Los Angeles, and University of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center, figure 1.

The SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing-edge flap on each blade. The development effort included design, fabrication, and component testing of the rotor blades, the trailing-edge flaps, the piezoelectric actuators, the switching power amplifiers, the actuator control system, and the data/power system. Development of the smart rotor culminated in a whirl-tower hover test which demonstrated the functionality, robustness, and required authority of the active flap system.

The eleven-week wind tunnel test program evaluated the forward flight characteristics of the active-flap rotor, gathered data to validate state-of-the-art codes for rotor noise analysis, and quantified the effects of open- and closed-loop active-flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness and the reliability of the flap actuation system were successfully demonstrated in more than 60 hours of wind-tunnel testing. The data acquired and lessons learned will be instrumental in maturing this technology and transitioning it into production.

The development effort, test hardware, wind-tunnel test program, and test results will be presented in the full paper. A brief description of the test hardware and test program is provided below.

Hardware and Test Program The 34-foot diameter rotor has five composite blades with 10in chord and a 25% chord trailing-edge flap from 74 to 92% radius, figure 2. The flap is mounted to the blade using five equally spaced hinges. A piezoelectric actuator is installed in the blade spar at 74% radius. It drives the flap via a linkage that is connected to a horn at the inboard end of the flap. Rotor instrumentation includes a five-component balance, drive-shaft torque, and control-system motions. Blade load measurements include 8 flap, 6 chord and 6 torsion moments at various radial stations, as well as pitch link load. The piezoelectric actuator stroke, force, voltage and current are measured on each blade. All rotating-system measurements are acquired and multiplexed in a hub-mounted data system. The multiplexed data and the actuator control power are transmitted through a conventional 36-channel slip ring to the ground- station data system. Acoustic measurements include 4 fixed and 8 traversing microphones on the advancing side of the rotor disk. Acoustic data and wind-tunnel operating conditions are recorded by the NFAC wind-tunnel data system but rotor data are recorded by the Boeing data system.

The 11-week long wind-tunnel test was sponsored by DARPA and NASA. The objectives of the DARPA funded portion of the test were to acquire loads, performance, and acoustic data for an advanced rotor system in validating high-fidelity physics-based rotor-noise prediction tools that had Abstract submitted for the 35th European Rotorcraft Forum, Hamburg, Germany, Sept. 22-25, 2009 been developed under the Helicopter Quieting Program (HQP). Test conditions for the validation included level-flight, descent, and high-speed cases with single and multiple harmonic flap inputs. Five test points were defined with speeds of 83, 123, and 155 knot. For each point the rotor thrust, the model pitch angle, the baseline (zero), and a flap-deflection schedule were specified. Flap inputs included amplitudes up to 3 deg at specified phasing and 2, 3 or 5/rev harmonic variation. All predictions were completed before the test. Data was successfully acquired at four test conditions. Blade loads were too high for the high-speed condition at 155 knot.

The primary objectives of the NASA funded portion of the test were to evaluate the effect of open- and closed-loop active flap control on rotor loads, noise, and performance. Open-loop flap inputs at 2 to 6/rev with fixed amplitude and phase sweeps were used to determine the optimum phase angle for noise reduction. Amplitude sweeps at optimal frequency and phase were then used to further reduce noise. For all noise test points, feedback of the actuator position was used to control the flap deflections precisely.

Feedback of rotor-balance forces and moments was used for closed-loop control of vibrations. The impact of flap-control on rotor dynamics (modal identification) and flight controls (rotor response) was assessed by performing various open-loop flap frequency sweeps. The effectiveness of the active flap on control power and rotor smoothing (i.e. blade tracking) was evaluated using open-loop steady-state flap inputs. Test conditions included hover, descent and level flight cases at 62, 68, 82, and 124 knots.

The effectiveness of the active flap control on noise and vibration was conclusively demonstrated.

Preliminary results show significant reductions in blade-vortex-interaction (BVI) and in-plane noise as well as vibratory hub loads. Up to 80% reduction in vibratory hub loads and up to 6dB noise reduction were measured. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power and rotor smoothing was also demonstrated.

Results illustrating the effectiveness of the active flap in various objectives noted above will be presented in the full paper. Sample results are discussed below.

Figure 1: SMART rotor in the 40- by-80- Figure 2: Close-up view of the SMART rotor, foot wind tunnel of the National Full-Scale blade, and flap Aerodynamic Complex at NASA Ames Research Center.

SMART Rotor Development and Wind Tunnel Test

Friedrich K. Straub Vaidyanathan R. Anand Terrence S. Birchette Benton H. Lau Boeing Technical Fellow Dynamics Engineer Design Engineer Aerospace Engineer The Boeing Company The Boeing Company The Boeing Company NASA Ames R.C.

Mesa, Arizona Mesa, Arizona Mesa, Arizona Moffett Field, California ABSTRACT Boeing and a team from NASA, Army, DARPA, Air Force, MIT, UCLA, and U. of Maryland have successfully completed a wind tunnel test of the smart material actuated rotor technology (SMART) active flap rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART active flap rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero- acoustic analysis, and quantified the effects of open and closed-loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full- scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed reductions up to 6dB in blade-vortex- interaction and in-plane noise, as well as reductions in vibratory hub loads of about 80%. Trailing- edge flap deflections were controlled with less than 0.2 deg rms error for commanded harmonic profiles of up to 3 deg amplitude. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind tunnel testing.

INTRODUCTION blade control (IBC) at the blade root [3-7], and trailing Vibration, noise, and aerodynamic design compromises edge, active flap control (AFC) on the blade [8,9] have are inherent barriers to significant improvements in been successfully flight tested. Significant reductions in effectiveness, productivity, and public acceptance of the vibrations of about 80%, in BVI noise during descent of helicopter. Specific rotary-wing challenges include about 5dB, and in rotor power during high speed level cyclic variations in free stream velocity, blade controls flight of about 6% were demonstrated. Similar results and motions, transonic flow on the advancing side, have also been reported from numerous model and full- reversed flow and dynamic stall on the retreating side, scale rotor wind tunnel tests.

blade/vortex interaction (BVI), blade/fuselage flow interactions, swashplate mechanical constraints, and On-blade active control using smart materials has flight control hydraulic actuator bandwidth. recently been applied in a number of model scale [10- 16] and full-scale rotor experimental programs Passive design techniques, such as optimized airfoils, [8,17,18]. Piezoelectric materials are used for high tip shapes, tuned blade structures, etc. are successful in frequency actuation at small amplitudes. Some of the providing incremental gains. Still, in many cases model rotor programs [13,14] used distributed piezo operational restrictions are necessary to mitigate the fibers to effect elastic twist actuation of the blade. All noise impact, vibration absorbers are required to reduce of the full-scale rotors used discrete piezo actuators and vibrations, and aerodynamic performance in various trailing edge flaps. This approach offers a number of mission segments is compromised by the constraints on advantages compared with HHC and root pitch IBC.

blade design. The undesirable loading is suppressed at the source, requiring less actuation power. The system can be Numerous active control concepts have been tailored to the blade aeromechanics through spanwise investigated that effect rotor blade motion or shape at placement and even multiple flaps [8,19]. It is frequencies above 1/rev in order to mitigate unsteady independent of the primary flight control system, not effects, or adapt the blade to mission segments with constrained by the swashplate, and uses electric power.

conflicting design requirements. Higher harmonic Solid state piezoelectric actuators have high bandwidth control (HHC) through the swashplate [1,2], individual and a minimum of moving parts.

Presented at the 35th European Rotorcraft Forum, Hamburg, Germany, Sept. 22-25, 2009.

and the data/power system. Development of the Modeling of rotors with on-blade controls has seen SMART rotor culminated in a whirl tower test in 2003 increased activity in recent years. A comprehensive (see Fig. 1c), which demonstrated the functionality, assessment of six different on-blade control schemes robustness, and required authority of the active flap and IBC for performance enhancement [20] used the system [18]. Additional details on the design, comprehensive code CAMRAD II [21]. Simultaneous development and testing of this active flap rotor system reductions of vibration and noise as well as vibration are provided in [26].

and rotor power for active flap rotors were shown to yield suboptimal results in the individual objectives The SMART rotor is a 33.85-ft diameter (blade radius, [22,23]. Most recently, a coupled CFD-CSD code has R of 203.1 inches), full-scale, bearingless, five-bladed been applied to model the SMART active flap rotor main rotor modified from the MD 900 Explorer rotor [24]. Correlation with the data reported here has shown system. Each blade consists of 12% thick HH-10 airfoil improved prediction of cyclic flap bending and torsion sections inboard up to 74% radius and 9.5% thick HH- moments over comprehensive codes, and the ability to 06 airfoil sections outboard beyond 84% radius, with a capture 3-dimensional flow effects at the edges of the linear twist of ñ 10 degrees. The blade tip region, from active flap. Correlation of CMARAD II predictions 93% radius to the tip has a parabolic leading edge sweep with the SMART rotor data set [25] has shown (22 degrees at the tip), straight trailing edge and a 2:1 reasonable results for flap bending moment and pitch taper ratio. The constant chord section of the blade has link load, but also shown the need to further refine the a 10 inch chord. Nominal rotation speed of the rotor is blade structural/inertia properties of the model. 392 RPM producing a tip speed of 695 ft/sec. At 5,811 pounds thrust, the rotor thrust coefficient normalized by Under a joint DARPA/NASA/Army-funded program, thrust-weighted rotor solidity is 0.075 at sea level Boeing and a team from the Air Force, NASA, Army, standard conditions. Rotor properties are summarized DARPA, Massachusetts Institute of Technology, in Table 1.

University of California at Los Angeles, and University of Maryland have recently completed a successful wind Of particular interest here is the design of the flap tunnel test of the SMART active flap rotor in the Air system and its integration into the MD 900 rotor blades.

Force National Full-Scale Aerodynamic Complex Aerodynamic and aeroelastic simulations were (NFAC) 40- by 80-foot anechoic wind tunnel at NASA conducted early on [27] to define the flap type, flap Ames Research Center. The eleven-week wind tunnel chord and flap span such that the flap system has test program evaluated the forward flight characteristics enough control authority to provide required dynamic of the full-scale active flap rotor, gathered data to lift variations for vibration and noise reduction and validate state-of-the-art codes for rotor aero-acoustic requires the minimum actuator power. The flap system analysis, and quantified the effects of open- and closed- selected has a flap chord of 25% (hinge to trailing edge) loop active-flap control on rotor loads, noise, and with an overhang of 40% (total flap length of 35% performance. chord) and flap span of 18% rotor radius with its center located at 83% rotor radius to provide required control The present paper briefly describes the SMART rotor authority while minimizing the flap hinge moments development and hardware. It focuses on the wind [26]. The flap is mounted to the blade using five tunnel test program, test setup, and provides an equally spaced hinges to minimize stresses (see Fig. 1c).

overview of test results. Its amplitude is mechanically limited to 6 deg. Flap properties are summarized in Table 2.

SMART ACTIVE FLAP ROTOR The Boeing SMART rotor, developed under the Each blade contains an embedded 2x-frame actuator sponsorship of DARPA, NASA, Army, and internal with four piezoelectric stack columns (see Fig. 1d) [28], funding, is a MD 900 helicopter (see Fig. 1a) full-scale, designed to drive the trailing edge flap at frequencies up five-bladed bearingless rotor, modified with to 11-per-rev (11P) with as much as 4 deg amplitude piezoelectric-actuated trailing edge flaps on each blade authority under load. Each actuator is powered by a 2- (see Fig. 1b). The objective of the development of this channel switching power amplifier. The piezo stacks rotor system was to demonstrate significant rotor- are driven by a DC bias voltage and a dynamic voltage.

induced vibration and BVI noise reductions and The dynamic voltage for the outboard x-frame is 180 aerodynamic performance improvements in wind tunnel deg out-of-phase relative to the voltage for the inboard and flight tests. The development effort included x-frame actuator. In this paper, references to voltage design, fabrication, and component testing of rotor refer to the dynamic voltage only. Actuator properties blades, trailing edge flaps, piezoelectric actuators, are shown in Table 3. Inputs to the five actuators are switching power amplifiers, actuator control system, controlled using a PC-based system. Typically, inputs are phased azimuthally such that each flap receives the When installed in the NFAC 40- by 80-foot wind tunnel same command at a given azimuth. The flap deflection test section, the LRTS is mounted on a three-strut th for the k blade is support system placing the rotor hub 23.7 ft above the tunnel floor at zero degree shaft tilt (see Fig. 3a). The ‰ = A sin (n  +  ), n = 0 Ö 11 fk n k n LRTS outrigger arms are mounted to two fixed, faired struts via ball joints and the tail sting is mounted to an where  =  ñ (k-1) 2  /5, k = 1-5.

k 1 extensible tail strut that provides shaft tilt. Three sets of Positive flap deflections are trailing edge down. For the fairings are used to enclose the sled/motor/gearbox, the rest of this paper, active flap settings will be described main strut, and the balance/controls.

in a three-parameter form, A /nP/  ñ where amplitude n n A and phase  are expressed in degrees, n is an integer n n Instrumentation multiple of the non-dimensional rotor speed, and P Blade load measurements include 6 flap, 4 chord and 4 stands for per rev.

torsion moments at various radial stations, as well as one set of flap, chord, torsion moments each on the The piezoelectric actuator is installed in the blade spar pitchcase and flexbeam of this bearingless rotor. Pitch at 74% radius. It drives the flap via a linkage that is link load and drive shaft torque are also measured. The connected to a horn at the inboard end of the flap (see piezoelectric actuator stroke and force are measured on Fig. 1e). A comparison of SMART and MD900 blade each blade. Piezo stack temperature and active flap lift mass properties shows that the flap/actuator system load at two intermediate hinges are measured on one increases blade weight by about 5 lb (see Table 4). A blade. All rotating system measurements are acquired comparison of SMART and MD900 blade modal and multiplexed in a hub-mounted data system (see Fig.

frequencies for computed rotating and measured free- 3b). The multiplexed data and the actuator control free conditions is shown in Table 5. In the analysis, the power are transmitted through a conventional 36- test stand control system stiffness was used for both channel slip ring. Test stand measurements include hub rotors, and the active flap was locked out. Results show accelerations, static mast bending, the five-component that the design goal to match SMART blade dynamics rotor balance, swashplate actuator motions, optical shaft to the baseline blade was achieved. In particular it encoder, and numerous test stand health parameters.

should be noted that no effort was made to lower the Wind tunnel test conditions are measured using a blade torsional stiffness and thus increase the active flap redundant set of analog as well as digital sensors.

effectiveness. Also shown is the fundamental mode of the actuator/flap (1TEF) when installed in the blade at For acoustic measurement, a series of microphones was about 96 Hz.

strategically placed around the model to capture rotor noise sources of interest (see Fig. 4a). These WIND TUNNEL TEST SETUP microphones were grouped into: a) out-of-plane fixed Large Rotor Test Stand microphones (M1 and M4) to correlate to microphones In preparation for the wind tunnel entry, a brief whirl used previously in the MDART test [29] b), traverse tower test was conducted (see Fig. 2). This system microphones (M5 through M12) that can be moved integration test verified operation of the SMART rotor, along guided rails for blade-vortex interaction noise large rotor test stand (LRTS), rotor control console mapping, c) in-plane microphones (M13, M15 and (RCC), test stand health monitoring system (HMS), flap M14) for low frequency, in-plane rotor noise actuator power amplifier and control system, and data measurement, and d) fixed microphone (M16) on the acquisition, processing, and display systems.

rotor balance fairing. Microphones M13, M15 and M14 were mounted on tower struts to be near in-plane of the Boeing í s LRTS (Fig. 2), consists of a sled structure that rotor (approx. 10 degrees below wind tunnel horizon).

supports a 1500-hp General Electric motor, 1500-hp With the exception of M14, all microphones are located gearbox, and tail sting. A vertical main strut is mounted within the acoustically-treated portion of the 40- by 80- to the gearbox and supports the balance housing, foot test section. The microphone traverse travel ranges balance, static mast, hydraulic actuators, swashplate, from 200 inches upstream of the center of the rotor hub and rotor. The static mast encloses the final drive shaft to 200 inches downstream with traverse stopping at that transfers torque to the rotor hub. Two horizontal every 40 inches. The plane of traverse microphone outrigger arms attach to the main strut just above the positions is located 89.4% radius below the rotor hub gearbox. The entire stand is mounted to a tripod center, extends from 41% to 141% radius across the test support. The LRTS was previously used during the section on the advancing side, and from 98.5% radius MDART (MD900 pre-production rotor) 40- by 80-ft upstream to 98.5% downstream. A top view of the wind tunnel test entry [29].

microphone layout is shown in Fig. 4b, and additional details are provided in two companion papers [30, 31].

real time monitoring. Test stand HMS data is monitored and recorded on a dedicated system. Typically 12 sec Active Flap Control The flap actuator high-voltage power amplifiers and of data are post-processed per test point; for active flap PC-based controller are located in the control room. frequency sweeps and closed-loop vibration control The user interface for providing control inputs and points, 40 sec are used. All acoustic data and wind- monitoring execution is developed using dSPACE tunnel operating conditions are recorded by the ControlDesk software. The open and closed-loop azimuth-based NFAC wind-tunnel data system, using control laws are implemented on a dSPACE DS1103 10 sec per point (within the 12 sec of Boeing data). The single-board controller. The board has 20 channels of NFAC acoustic data acquisition and reduction system analog input, 8 channels of analog output, and other enabled the near real-time acoustic processing of the digital I/O channels and external interrupt inputs. A data. After the tests, both data sets are aligned using a dedicated Simulink blockset is used to simplify triangular wave form alignment signal and rotor development of real-time controllers in the azimuth.

MATLAB/Simulink environment. Rotor azimuth angle is determined from 1/rev and 512/rev signals provided When post-processing steady-state test points, sixty-four by the optical encoder. The commanded actuator revolutions of data (approximately 9.75 seconds) are voltages output from the dSPACE controller are used. All channels, except acoustic measurements, are amplified by high-voltage amplifiers, and the resulting re-sampled to 256 samples/rev on an azimuth basis. The actuator signals are passed via the test stand slip ring acoustic data channels are digitized at an effective into the rotating frame and to the individual actuators. sampling rate of 2048 samples/rev (equivalent to 13,380 One active flap command voltage, actual voltages and samples/sec at the nominal 392 RPM). The data currents on all flap actuators, and amplifier power exhibited good rev-to-rev repeatability; therefore, a supply voltage and current are measured. straightforward synchronous average of the time history data resulted in an averaged time history of one The active flaps are controlled in one of three modes. revolution duration of 256 or 2048 points.

For open-loop control, actuator voltage is specified and no feedback from the rotor is used. For closed-loop A hub weight tare, rotation tare, and aerodynamic tare position control, the flap deflection is specified in were taken before the test, in the configuration used degrees, and feedback loops are closed on 5 actuator during the test, except that the blades were replaced by strokes, which are kinematically related to flap spacers placed inside the pitchcase to firmly restrain the deflections. Either a continuous time higher harmonic flexbeam tips. During the weight and rotation tare, the controller (CTHHC) [32] or a modified version of the test section overhead and access doors were open. A discrete time controller (HHC) from [33] is used. For blade weight tare was taken before and after the test.

closed-loop control of vibratory hub loads, rotor balance The rotor was balanced and tracked using an Advanced loads are used for feedback with the CTHHC controller. Vibration Analyzer (AVA) and optical strobe. No commands were made to the active flaps. Final balance Data Acquisition and Processing was about 0.1 in/sec and track was well within one Data channels from the rotating portion of the model in chord thickness at the tip.

the tunnel (blades, hub, pitch links, flaps, etc.) are combined on the hub by a Metraplex Mini 770 data TEST OBJECTIVES AND APPROACH acquisition system into a pulse code modulated (PCM) The 11-week long wind-tunnel test was sponsored by serial data stream, and transmitted through one of the DARPA, NASA, and the U.S. Army. The objectives of slip ring channels into the fixed frame to the control the DARPA-funded portion of the test were to acquire room. Data channels from the non-rotating portion of loads, performance, and acoustic data for an advanced the model (test stand, balance, etc.) are combined in the rotor system in support of validating high-fidelity control room with flap actuator and amplifier voltage physics-based rotor noise prediction tools that had been and current measurements, rotor RPM and azimuth, the developed under the Helicopter Quieting Program wind tunnel operating condition analog measurements, (HQP).

and four microphones (M1,4,13,16) by another Metraplex Mini 770 system into a second, non-rotating Test conditions for the validation data base included PCM data stream. level-flight, descent, and high-speed cases with single and multiple harmonic flap inputs. Four test points Both PCM data streams are time-based and recorded were defined with speeds of 83, 123, and 155 knot, (see continuously on Boeing í s data system. Key parameters Table 6, Validation, condition 1-4). For each point the are displayed in various formats on four flat panel velocity, advancing tip Mach number, shaft angle, blade displays and two 18-channel electronic strip charts for loading, the baseline (0 deg) and a flap deflection schedule were specified. Flap inputs included (rotor response) were investigated by performing open- amplitudes up to 3 deg at specified phasing and 2, 3 or loop flap linear and logarithmic frequency sweeps up to 5/rev harmonic variation. All predictions were 80 or 200Hz or around rotor speed multiples of interest.

completed before the test, and therefore flap position Amplitudes of 150, 175, and 200V and the VSP control was used to closely follow the specified flap collective, longitudinal, and lateral cyclic inputs were deflections. Data was successfully acquired at three test used.

points. Blade loads were too high for the high-speed condition at 155 knot to exercise the specified flap For each test condition the rotor was trimmed to the deflection schedule, however, baseline data (with zero desired thrust and minimal rotor flapping as determined degree flap deflection) was acquired. from flexbeam cyclic flap bending. After trimming at the baseline condition, unless noted, the rotor was not The primary objectives of the NASA-funded portion of retrimmed during subsequent flap phase and amplitude the test were to evaluate the effect of open and closed- sweeps or microphone traverse.

loop active flap control on BVI noise, in-plane noise, and vibratory hub loads. Secondary objectives were TEST RESULTS control power, rotor smoothing, and performance, (see A sample of test results covering all the objectives is Table 6), as well as rotor dynamics and flight control presented here. Detailed results for the BVI noise, in- system identification from active flap frequency sweeps. plane noise, vibration reduction, and flap position Test conditions included hover, descent and level flight control are shown in the companion papers [30, 31, 32].

cases at 62, 68, 82, and 124 knots. For each test point All test results shown here were obtained at a nominal advance ratio, tip Mach number, shaft angle, and blade thrust coefficient-to-solidity ratio (C /  ) of 0.075 T loading were specified. The advance ratio and tip Mach (which corresponds to 5811 pounds of rotor thrust at sea number specified are based on the specified level flight level, standard day conditions). Exceptions were hover speed, nominal rotor tip speed, and sea level standard cases, typically run at Ë = 4 deg blade collective pitch day conditions.

and 10 deg forward shaft tilt to reduce recirculation, and the performance data with 2/rev inputs.

For noise reduction, single harmonic flap inputs at 2 to 5/rev with fixed amplitude (typically 1.5 deg) and phase Position Control sweeps (0 to 360 deg in 30 deg increments) were used One of the goals of the wind tunnel test program was to to determine the optimum phase angle. Amplitude precisely control blade flap position in a number of sweeps at optimal frequency and phase were then used specific flight conditions, to allow for correlation of to further reduce noise. For all noise test points, wind tunnel acoustic data with pretest predictions of feedback of the actuator position was used to control the rotor aeroacoustics. Analysis of an early SMART rotor flap deflections precisely. For vibration reduction, configuration in forward flight has shown that flap open-loop flap control at 2 to 6/rev with 250V dynamics can result in considerably larger than open- amplitude and phase sweeps were used to establish loop commanded flap deflections in the first quadrant controllability. Closed-loop vibration control used the [34]. Furthermore, SMART whirl tower data has shown CTHHC controller with feedback of rotor-balance differences in the open-loop active flap response on normal force, roll, and pitch moment (NF, RM, and each blade [35], possibly due to the combined effects of PM).

flap rigging, flap control system stiffness, and piezoelectric actuator performance. Thus, the need for The effectiveness of the active flap for control power closed-loop flap position control was indicated. All was evaluated using position control at 0 and 1/rev with results shown here were obtained with the CTHHC amplitudes of -3 to 3 deg and appropriate phasing.

controller. It was based on the actuator voltage to flap Flaps were controlled either individually, specifying position transfer function, identified from frequency harmonic and phase (IBC), or through software sweep data at 82kt, and controls harmonics 0-6/rev [32].

implementation of a virtual swashplate (VSP), using A single controller was used for all test points, since the amplitude and collective, longitudinal, and lateral cyclic transfer function was essentially invariant with flight commands. The effectiveness of the active flap for condition.

rotor smoothing (i.e. blade tracking) was evaluated using position control with steady flap amplitudes of -3 For the wind tunnel test, the flap rigging ( Ÿ =0) was to 3 deg on a single flap, either flap 1 or 2. The very tightly controlled to within 0.1deg for 5 flaps, (see effectiveness of the active flap for rotor performance Fig. 5a). In addition, it is noted that the nonrotating improvements was evaluated using position control at actuator/flap fundamental frequencies are within 2.5% 2/rev with 1.5 deg amplitude and phase sweeps. Rotor of their average value (not shown). Figure 5a shows the dynamics (modal identification) and flight controls flap deflection (min, max, and average of 5 flaps) versus speed for the uncontrolled case (0V), that is no voltage are crossed. Very similar results are seen in Fig. 8b for going to the flap actuators. The aero/inertia loading Ï =0.3.

deflects the flaps from -0.5 to 3 deg, with the range of values increasing with speed. Figure 5b shows that for From the above it is clear that the active flap affects closed-loop flap position control with a 0 deg command, overall blade response via the moment, or servo, effect.

flap deflections are maintained within ±0.1 2 deg from 0 At frequencies below the torsion mode, flap down deg. The largest error is seen at the 62kt descent deflection increases lift at the trailing edge which results condition, possibly a result of BVI. For harmonic flap in a nose down pitching moment and blade torsion commands (Fig. 5c), the mean flap position is well response, and thus reduced section lift. Above the within ±0.1 deg from 0 deg and the largest rms error of torsion mode the blade torsion response reverses. Blade 0.2 deg occurs for the 2 deg 5/rev command at 124 kt. flapping at 0 and 1/rev responds to the reduced lift as Figures 5b-c show that the CTHHC controller was very expected with down flapping (about 90 deg phase lag effective in controlling flap position. Similar results relative torsion at 1/rev). At higher frequencies, the were seen for the HHC controller using on-line blade flapping response changes phase rapidly and identification and harmonics 0-5/rev. examination of flap bending and torsion at outboard stations, where the flap affects local airloading, may be Flap deflection time histories for three cases are shown required to better understand the flapping response and in Figures 6a-c. The uncontrolled case (0V, Fig. 6a) its impact on noise and vibration. Unsteady illustrates the differences between 5 flaps and aerodynamics also plays a role, however, unsteady lift is significant effect of aero/inertia loading on flap not expected to add more than 20-30 deg phase lag.

deflections around 90 deg azimuth. The harmonic flap deflections in Figures 6b-c illustrate the fidelity in BVI Noise Reduction matching commanded flap deflection profiles. Small High noise levels in helicopter descent or maneuvers are differences are seen from flap to flap and for 5/rev the caused by an impulsive noise-generating mechanism error at peak flap deflection may be as large as 0.4 deg. known as blade-vortex interaction (BVI) that results from the close proximity between the main rotor blades Rotor and Active Flap Dynamics and the vortices generated by them. BVI noise radiates Data was collected for about 200 active flap frequency out-of-plane and is the primary source of noise sweeps (chirps) in hover, 82, and 124 kt. Open-loop annoyance around heliports when the rotor is close to flap inputs were made using collective, longitudinal, the ground during landing approach. BVI noise can be and lateral cyclic modes. Sweep range from 0-200 Hz reduced by redistributing airloads near the blade tip to nd was used for flap actuator dynamics. A range from 0-80 reduce the tip vortex strength in the 2 quadrant or by Hz was used for rotor dynamics, and 0-9 Hz was used increasing the vertical separation (miss distance) for flight controls. Close-up sweeps were conducted between the blade and the tip vortex during the spanning (n ± 1/2)/rev for harmonics 2-6 and 10. Figure interaction.

7 shows pitchcase torsion response in hover to a 0.2-80 Hz sweep with logarithmic rate and 200V collective Effects of the active flap on BVI noise were evaluated at amplitude. The blade fundamental torsion mode is three descent flight conditions with Ï =0.15, 0.165, and readily seen at about 5.8/rev in the power spectral map 0.2. Shaft angle sweeps were conducted for the baseline and phase plot. rotor (0V flap command) to identify the shaft angles corresponding to maximum BVI. For condition 2 the Phasing of the active flap angle (+ TE down) relative to shaft angle was estimated at 1.8 deg aft for the 6 deg actuator voltage, and flexbeam flap bending (+ up) and glide slope used during MD900 FAA certification, and flexbeam torsion (+ LE up) relative to flap angle are rotor speed was set at 392 rpm. All subsequent BVI shown in Figures 8a-b for 82 and 124kt respectively. noise test points were obtained with flap position These are obtained from active flap excitation and control. A traverse sweep for the baseline rotor responses at integer harmonics. Phase lead is positive. (simulated with flap position at 0deg) was then Figure 8a ( Ï =0.2 ) shows that the flap deflection lags the conducted to identify the microphone location with the voltage by about 20 degrees across the entire range. highest BVISPL. BVISPL was determined from the The phase of flexbeam torsion to active flap decreases band-pass filtered spectrum between blade passage gradually from 180 deg at 0/rev to about 40 deg at 6/rev. harmonics 8-60. The optimum combination of active It crosses 90 deg between 5 and 6/rev, near the blade flap parameters for BVI noise reduction at the selected torsion mode. The phase of flexbeam flap bending to location was then determined through systematic phase, active flap is 180deg at 0/rev and decreases rapidly with frequency, and amplitude sweeps. For additional details increasing frequency as the three flap bending modes see [30].

Figure 9 shows the change in BVISPL from baseline for the directivity characteristics of thickness noise near in- microphone M7 at traverse station -120 as a function of plane of the rotor, it was proposed to use on-blade flap phase, with 1.5 deg amplitude at single harmonics controls to alter blade airloads and generate an in-plane 2-5 (condition 1, Ï =0.15, · =4 ˚ ). Noise reductions from loading noise profile that would negate or reduce the 3 to 6 dB are seen at the best phase for each of the four thickness noise pulse [31]. Achieving this ì anti -noise î harmonics. An amplitude sweep from 1 to 2 deg at the profile would require an increase in the in-plane loading best frequency and phase combination of 4P/30 ˚ as the blade approaches the advancing side near 90 ˚ determined that active flap actuation of 1.5 ˚ /4P/30 ˚ blade azimuth.

provided best BVISPL noise reduction at all microphones at this traverse location. Effects of the active flap on in-plane noise were evaluated for a level flight condition ( Ï = 0.3, · = -9.1 ˚ ), A traverse sweep was conducted with the 1.5 ˚ /4P/30 ˚ considering microphone M13 and a low frequency flap actuation. Figure 10a shows the carpet plots of sound pressure level (LFSPL) noise metric, which BVISPL contours for the baseline rotor, rotor with includes acoustic energy only in the first six blade- 1.5 ˚ / 4P/30 ˚ flap actuation , and the difference (active passing harmonics. All test points were obtained with flap rotor noise levels minus the baseline rotor noise flap position control.

levels) in noise levels between the two. BVISPL values used in these contours are averaged over repeat data Figure 11a shows the change in LFSPL from baseline points where available. Figure 10a clearly shows that for microphone M13 as a function of flap phase, with the active flap was able to reduce the BVI noise over a 1.5 deg amplitude at single harmonics 2, 3, and 5P.

wide range of directivity angles under the advancing Only a 1 deg flap amplitude was realized for the phase side with BVISPL reductions as high as 7 dB. With sweep at 4P due to high blade loads. Noise reductions active flap, the BVISPL at the baseline hot spot location from 4 to 5 dB are seen at the best phase for each of the was reduced by as much as 3.5 to 6 dB. Figure 10b four harmonics. Amplitude sweeps from 0.7 to 2 deg shows the time history and spectral data comparisons for the best frequency and phase combinations 2P/0 ˚ , between the baseline rotor and the rotor with 3P/250 ˚ , and 4P/180 ˚ were conducted to further explore 1.5 ˚ /4P/30 ˚ flap actuation for microphone M7 at the effectiveness of the active flap. Within this range of traverse station -120. Figure 10b clearly shows the flap amplitudes, Figure 11b shows that increasing flap reduced acoustic pressure and higher harmonic spectral amplitudes at 3P and 4P achieved more noise levels with flap actuation resulting in 7.1 dB reduction reductions. Best noise reduction of 5.1 dB and 5.7 dB in BVISPL relative to the baseline, thus demonstrating was achieved at 2.0 ˚ flap amplitude at 3P, and at 1.3 ˚ the effectiveness of harmonic flap actuation for flap amplitude for 4P, respectively. Beyond these reducing BVI noise. However, this active flap actuation measured flap amplitudes, extrapolated trends suggest which produced large BVI noise reductions also that there is an optimum point whereby a further produced large increases in vibratory hub loads relative increase in flap amplitude does not necessarily result in to those for the baseline rotor. more noise reductions. This is shown to be the case for 2P where noise reduction margin diminishes from 2.8 Similar results [30] were obtained for the simulated dB at 1.5 ˚ flap amplitude to 2.0 dB at 2.0 ˚ flap amplitude.

FAA noise certification descent flight (  =0.165). A flap schedule of 1.5 ˚ /3P/180 ˚ was able to reduce BVI Examination of the active flap motion and blade torsion noise levels by as much as 5 dB. Noise reductions at moment at 0.82R showed that all reduced-noise the baseline rotor BVI hot spot locations varied between 3 and 5 dB. For the moderate high speed test case conditions exhibit decreasing active flap deflection (maximum flap up rate) near 90 ˚ azimuth, while (  =0.2), a flap schedule of 1.5 ˚ /3P/180 ˚ produced simultaneously the blade twists LE up [31]. It was smaller BVISPL reductions, with a maximum of about 3 concluded that reduced noise was likely due to active dB.

flap induced changes in aerodynamic loading and blade torsion response at the blade tip near 90 ˚ azimuth. The In-plane Noise Reduction net effect was to increase in-plane blade forces towards Low frequency rotor harmonic tones, emitted from near the trailing edge and thereby reduce in-plane noise. Not in-plane of the rotor, are particularly of concern for surprisingly then, reduced noise was accompanied by an military operations, as they tend to propagate long increase of in-plane hub loads at 5/rev.

distances without substantial attenuation by atmospheric absorption. At positions near in-plane and forward of The underlying mechanism of these reduced in-plane the rotor, the radiated noise is primarily due to thickness noise levels is illustrated for the three best cases from and in-plane loading mechanisms. Because the Figure 11b. Figure 12 shows the ability of the active directivity of the in-plane loading noise nearly matches flap to change in-plane loading and thus generate The effectiveness of the active flap to modify the appropriate ì anti -noise î pulses that partially cancel the aerodynamic loading and reduce vibratory hub loads negative pressure peak commonly associated with was evaluated at two flight conditions, descent ( Ï =0. 2, steady thickness noise, and thereby reduce the net · =2 deg) and level flight ( Ï =0. 3, · = -9.1 deg). Open- acoustic radiation forward of the rotor [31]. Active flap loop active flap control phase sweeps were conducted at control was found to reduce negative acoustic pressure both conditions with 250V amplitude for a single peaks by at least 50%. These reductions, however, did harmonic from 2-6P in order to establish control not occur uniformly for the pulses emanating from all sensitivity. Figure 13 shows the sine and cosine 5P five blades. Compared to the baseline acoustic balance loads at Ï =0. 3 for 4, 5 and 6P inputs with points signature, as well as the more benign 2P case, actuating forming a circle as phase is swept from 0 to 360 in 30 the trailing-edge flap at 3P and 4P appear to generate deg increments. Baseline (0V) loads are also shown, much stronger blade-to-blade differences. generally near the middle of the circle. The larger the circle, the more sensitive the load is to the specific input harmonic. From this it is clear that 6P is least effective Vibration Reduction The source of rotor-induced helicopter vibration is the while 4P is most effective. For the normal force 5P is unsteady environment experienced by the blades. The as effective as 4P, however somewhat less so for the unsteady blade forces are then transmitted through the other loads. Not shown, 2P was about as ineffective as hub to the fixed frame, and are felt as vibration in the 6P, while 3P was more effective than 5P for the in-plane fuselage. Generally, the forces on the blades are forces and the moments. For 4P and 5P the circle encloses the origin for all loads, indicating that any harmonics of the rotor frequency, Ÿ , since the rotor aerodynamics are (nearly) periodic. Theoretically, only single load can be zeroed out with a moderate voltage.

those harmonics that are multiples of the fundamental frequency at iN and iN ±1 (where N is the number of Also shown in Figure 13 are results for a vibration blades and i is an integer) produce vibration in the fixed index, which is a weighted sum of the squared 5P frame, due to the symmetry of the rotor. However, in balance loads. From the plot of vibration index practice, all harmonics contribute to vibration, due to magnitude versus phase, it is clear that control of asymmetries in the rotor, such as blade-to-blade multiple loads simultaneously with a single harmonic is imbalance and tracking error. not effective, because of the different phases required to zero out different loads.

Both open-loop (voltage command) and closed-loop control using CTHHC were applied. Assuming that the Closed-loop vibration control used feedback of a single effects of rotor dynamic periodicity are small and can be balance load at a time. At Ï =0. 2 normal force and roll neglected, the input-output relationship between moment at 5, 1, or 1-5P were used; at Ï =0. 3 balance controls and rotor loads should be time-invariant and normal force at 1-5P or 10P and pitch moment at 1-5P approximately linear. In that case, classical, discrete- were used. Active flap frequency sweeps were time HHC [36] can be extended to obtain a continuous conducted at both conditions using collective, time higher harmonic controller (CTHHC) of the form longitudinal, and lateral cyclic inputs to acquire the data needed to identify the transfer functions and determine 2 2 K (s) = 2/T (A s + B n Ÿ ) / ( s + (n Ÿ ) ) n n n n the controller constants [32].

where n is the harmonic to be controlled, T is the n desired time constant of the control loop, and A and B Figure 14a shows the spectrum (i.e. magnitude of the n n fast Fourier transform) of the balance normal force for are constants obtained from off-line identification of an appropriate rotor transfer function. The benefits of this Ï =0. 2, · =2 deg. Arbitrary units are used, since the magnitude depends on the number of samples in the approach are generally better phase and gain margins FFT. There are significant impulses in the spectrum at and the ability to use classical control techniques to all the integer harmonics from 1-11P, except 8P. This evaluate performance and stability. If from the Nichols plot it turns out that the choice of time constant T for indicates the presence of some asymmetry in the rotor, n due to blade-to-blade differences, rotor track and any of the controlled harmonics results in low gain or phase margins, one or more of the time constants may balance, or non-rotor induced loading. Also, there is a slightly broadened peak at about 6.2/rev, which is a have to be increased in order to achieve acceptable margins. In addition, the controller uses an integral transmission gear-tooth mesh frequency.

term K (s) = A / ( T s) to control steady forces or 0 0 0 displacements, and a modulation/demodulation scheme Figure 14b shows the transfer function identified for flap actuator voltage to normal balance load from a 0-80 to track small variations in rotor speed.

Hz collective sweep. Several features of the transfer function are noteworthy. First, several rotor modal frequencies are clearly visible, at approximately 1, 2.7, Surprisingly, the 10P vibration index is reduced by 76% 3.5, 4.3, 7.0, and 10.2 per rev. Second, the phase lag in (T=1).

the transfer function is 900 deg over the frequency range from 0-12/rev. The large increase in phase lag Closed-loop active flap control of vibratory normal with frequency has been seen in other actively force using CTHHC was extremely effective, reducing controlled rotors [11]. This large phase delay in the harmonics 1-5P by 95% for both the level flight and transfer function limits the achievable performance of descent condition. Control of vibratory roll moment in feedback controllers. descent and pitch moment in level flight was shown to be slightly less effective, reducing harmonics 1-5P by In order to control the harmonic vibrations in the normal 68% and 73%, respectively [32].

force, a feedback control law with balance normal load as the measurement signal, and the collective flap Closed-loop, simultaneous control of multiple loads voltage as the control signal is used. It targets the first using discrete time HHC or an advanced, discrete time five harmonics of vibration, with a sum of individual controller that takes blade-to-blade dissimilarities into components K(s) = ” K (s), where n=1-5, and the account [37] was planned for this entry, but no data was n controller constants are obtained from the transfer acquired in the available time.

function.

Control Power Figure 14c shows the resulting closed-loop spectrum of Control power from the active flaps was evaluated by the balance normal force. The spectrum is almost applying equivalent steady-state collective, lateral, or identical to the spectrum in the open-loop case (Fig. longitudinal cyclic flap deflections and observing the 14a), except that the impulses in the spectrum resulting changes in normal force, roll, and pitch corresponding to the first five harmonics are completely moment. Flap deflections were generated using a absent, which is to be expected, since the feedback virtual swashplate (VSP), i.e. software mixing for flap control K(s) is infinite at those harmonics. Harmonic inputs on each blade, or by applying 0P for collective load values are obtained by summing the energy in the 1P/90 ˚ for roll and 1P/180 ˚ for pitch inputs (IBC).

spectrum in a narrow frequency range ( ±0.5/rev) and Results for position control flap inputs, both VSP and converting it to rms vibration levels. The dominant IBC, and comparable swashplate inputs were obtained vibratory harmonic (1/rev) is reduced in magnitude by at Ï =0. 2 for level flight and descent, and for level flight 98%, and 4/rev and 5/rev are reduced in magnitude by at Ï =0. 3.

90%. The overall reduction of harmonic vibration (of the first five harmonics) is 95%. Further, the total Figure 16 shows that the active flap generates reduction in the normal force vibratory loads (including substantial changes in lift and hub moments at Ï =0. 3.

both harmonic and broadband vibration) over the range Thrust decreased by 1500 lb when moving the flaps 0.5 ñ 5.5/rev is 84.5%. collectively down from -3 to +3 deg. In comparison, thrust increased by 2700 lb when raising rotor collective Figure 15 shows the normal force and vibration index pitch from -1 to +1 deg relative to the trimmed position.

for baseline (0V) and five closed-loop controllers for Similarly, roll moment increased by 15000 in-lb for flap normal force (NF) at Ï =0. 3, · = -9.1 deg. The three inputs and 23500 in-lb for swashplate inputs. Pitch controllers for 1-5P use different transfer function and moment increased by 33700 in-lb for flap inputs and time constants. Harmonic load values are obtained from 45000 in-lb for swashplate inputs. These results the time history averaged over one rotor revolution. indicate substantial control power from the active flaps.

Result show that that the controlled normal force This is particularly noteworthy since the test stand has a components 1-5P are almost entirely eliminated, with very stiff control system (blade torsion mode near reductions of 80% for 5P, 98% for 1P and 73% for the 6/rev), thus limiting the flap í s moment control rms value of all harmonics. Corresponding results for effectiveness (servo effect) and resulting blade pitching the vibration index, including all five hub loads, are less motion. On the flight vehicle, with a much lower impressive, with no reduction at 5P, 29% at 1P and 26% control system stiffness, active flap effectiveness can be for rms. The uncontrolled normal force harmonics 6-9P expected to increase significantly [34].

appear unaffected, whereas the 10P harmonic increases slightly. The two controllers for 10P use different time Data was also acquired from multiple frequency sweeps constants. Results for normal force show good over a 0-9 Hz range and is being evaluated to assess reduction of the 10P normal force when using the first potential improvements in helicopter handling qualities controller (T=5) and excellent reduction of 98% when with high rate active flap actuators.

using the more aggressive controller (T=1). Again uncontrolled harmonics remain essentially unchanged.

Rotor Smoothing Rotor smoothing, or blade tracking, was evaluated Of particular interest is the use of active flap control to making steady-state inputs to individual flaps, either potentially improve rotor performance. Two/rev active flap 1 or 2, using position control to move the active flap inputs were made at 1.5 deg amplitude and varying flap down from -3 to +3 deg in 1 deg steps. Since no phase at Ï =0. 3, · = -9.1 deg and two thrust settings, C / Û T direct measurement of blade track was available, the =0.075 and 0.09. In this case the rotor was trimmed at effectiveness of the active flap for rotor smoothing was each test point. Results are presented for the rotor lift- assessed by considering the mean thrust and 1/rev roll to-drag ratio L/D = L / (P/V-D ), where D is the p p and pitch moments resulting from the rotating lift force parasite drag which is obtained from the propulsive due to flap deflections on one blade. Results were force.

obtained in hover, Ï =0.2 , · =2 ˚ , and Ï =0.3 , · = -9.1 ˚ .

Figure 18b shows the rotor lift-to-drag ratio L/D as a function of flap input phase. Results for the baseline Changes in hub loads from deflecting a single flap are case (0deg) before and after the 2P phase sweeps are shown in Figure 17. Figure 17a shows that mean thrust shown at 10 and 350 deg phase, respectively, for clarity.

decreases by 450 lb and sin1P roll and cos1P pitch Not unexpectedly, the results raise questions regarding moment increase by 4000 in-lb when moving the flap data consistency and repeatability. However, on blade 1 down from -3 to 3 deg. This corresponds to comparing the results from 2P flap inputs with the 2P about 7.5% of the nominal rotor thrust and 10% of the averaged L/D for each thrust condition, there is a clear rotor balance oscillatory moment limit loads. Figure trend. Higher L/D is seen at 0-180 deg phase, and lower 17b shows the roll moment sine and cosine 1/rev L/D at 180-360 deg phase. The difference between L/D components for three tunnel speeds, when moving flap at 90 and 270 deg is about 4%. Thus, a definite cause- 1. Clearly, the active flap effectiveness increases effect relationship between 2/rev flap inputs and rotor considerably with tunnel velocity; little effect is seen in performance is seen. Ignoring the outlier point at 60 hover. Also shown are results when moving flap 2, for deg phase, a 1% increase in L/D versus baseline is seen the Ï =0.3 case. The phase relationship indicated by the around 90 deg phase. Such a small change may be data properly reflects that blade 2 follows blade 1 by 72 within the measurement accuracy, but the change is in deg in azimuth.

the right direction and consistent with prediction [20].

Results show that, as customary, pitch link adjustments CONCLUSIONS should be used for track changes in hover. However, A wind-tunnel test of the SMART active flap rotor was measured thrust changes and observed blade track conducted in the 40- by 80-foot wind tunnel anechoic changes due to flap inputs also indicate that use of test section of the NFAC at NASA Ames. Loads, active flaps for rotor smoothing in forward flight performance, and acoustic data were acquired in support appears to be feasible.

of validating high-fidelity physics-based CFD-CSD rotor-noise prediction tools. The effectiveness of the Rotor Performance active flap control on noise and vibration was A limited evaluation of rotor performance was conclusively demonstrated. Results show reductions in conducted, considering two aspects. As noted before, blade-vortex interaction (BVI) and in-plane noise as when the active flap motion is not controlled (open- well as vibratory hub loads. Noise reductions up to loop, 0V) the flaps experience considerable deflections 6dB, as well as vibratory hub load reductions of about and some blade-to-blade variations in flap response.

80% were measured. Trailing-edge flap deflections While the blade-to-blade variations did show a were controlled with less than 0.2 deg rms error for reduction in BVI noise when compared to the case with commanded harmonic profiles of up to 3 deg amplitude.

flap position controlled to 0 deg (BVISPL lowered by The impact of the active flap on control power, rotor more than 2dB) [30], the opposite was expected for smoothing, and aerodynamic performance was also performance.

demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than Rotor performance comparison for 0 deg and 0V control 60 hours of wind-tunnel testing. Specific conclusions is shown in Figure 18a for a thrust sweep in level flight are as follows.

at Ï =0. 3, · = -9.1 deg. Results are presented for the rotor lift coefficient versus power coefficient, both 1. Both CTHHC and HHC were effective in controlling normalized by thrust-weighted solidity. In this case the active flap position, using position feedback and rotor was trimmed at 0deg but not retrimmed for 0V. It applying individual control to each flap. When is seen that the uncontrolled flap case carries a 1% commanding zero deflection, flap deflections were performance penalty, i.e. lower rotor lift at the same within 0.12 deg for all speeds tested. When power, for nominal and higher thrust values.

commanding harmonic deflection profiles, the rms MIT, UCLA, and University of Maryland are gratefully error was less than 0.2 deg. acknowledged. In particular, the authors would like to 2. Data for the validation of physics-based aero- thank Mr. Daniel Newman, Dr. William Warmbrodt, acoustic prediction codes was successfully acquired and Dr. Ram JanakiRam for their support and many at three test points. Blade loads were too high for the helpful discussions. Special thanks go to the authors of high-speed condition at 155 knot to exercise the the companion papers, Drs. Ben Sim, Ram JanakiRam, specified flap deflection schedule, however, baseline and Steven Hall for sharing their material. Thanks also data (with zero degree flap deflection) was acquired. go to Mr. Roger Smith for his evaluation and 3. Reductions in BVI noise in descending flight and in- assessment of SMART rotor performance.

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Elevon Rotor, î AHS 61st Annual Forum, 32. Hall, S.R., Anand, V.R., Straub, F.K., and Lau, Grapevine, TX, 2005. B.H., ì Active Flap Control of the SMART Rotor th 20. Yeo, H., ì Assessment of Active Controls for for Vibration Reduction, î AHS 65 Annual Forum, nd Performance Enhancement, î AHS 62 Annual Grapevine, TX, May 2009.

Forum, Phoenix, AZ, 2006. 33. Molusis, J.A., Hammond, C.E., and Cline, J.H., ì A 21. Johnson, W., ì CAMRAD II, Comprehensive Unified Approach to the Optimal Design of Analytical Model of Rotorcraft Aerodynamics and Adaptive and Gain Scheduled Controllers to Dynamics, î Release 4.7 , Johnson Aeronautics, Achieve Minimum Helicopter Vibrations, î AHS th Palo Alto, 2008. 37 Annual Forum, New Orleans, May 1981.

22. Patt, D., Liu, L., and Friedmann, P.P., ì Rotorcraft 34. Straub, F.K., Charles, B.D., ì Comprehensive Vibration Reduction and Noise Prediction Using a modeling of rotors with trailing edge flaps, î AHS th Unified Aeroelastic response Simulation, î JAHS 55 Annual Forum, Montreal, May 1999, JAHS Vol 50(1), Jan. 2005. Vol 46(3), July 2001.

23. Liu, L., Friedmann, P.P, Kim, I., and Bernstein, 35. Straub, F.K., and Anand, V.R, ì Whirl Tower Test D.S., ì Simualtenous Vibration reduction and and Analysis of the SMART Active Flap Rotor, î rd Performance Enhancement in Rotorcraft Using Presented at the AHS 63 Annual Forum, Virginia nd Actively Controlled Flaps, î AHS 62 Annual Beach, VA, 2007.

Forum, Phoenix, AZ, May 2006. 36. Shaw, J., J., ì A Feasibility Study of Helicopter 24. Ananthan, S., and Baeder, J.D., ì Prediction and Vibration Reduction by Self-Optimizing Higher Validation of Loads on Bearingless Rotors Using a Harmonic Blade Pitch Control, î Master í s thesis, th Coupled CFD-CSD Methodology, î AHS 64 MIT, Dept. of Aeronautics and Astronautics, 1967.

Annual Forum, Montreal, May 2008. 37. Roget, B., and Chopra, I., ì Individual Blade Control Methodologies for a Rotor with Dissimilar Blades, î JAHS, Vol 48(3), July 2003, pp.176 -185.

Table 1: SMART Rotor Characteristics Table 3: 2X-Frame actuator characteristics Rotor blade modified MD900 Blocked force 113 lb Hub type bearingless (MD900) Free stroke 0.081 in No. of blades 5 Maximum work 2.28 in-lb Rotor Diameter 33.85 ft (R=203.1in) Voltage, max (nom) 475 ± 725V (400 ±5 00V) Rotor Speed 392 rpm Weight 2.16 lb Tip Speed 695 ft/s Specific work 1.1 in-lb/lb Chord 10 in Airfoils HH-10, t/c=12%, to 0.74R Table 4: MD900 / SMART Blade Mass Properties HH-06, t/c=9.5%, from 0.84R MD900 SMART Tip Sweep parabolic LE, Weight (lbs.) 39.16 44.22 from 0.93R; 22deg at tip Span Moment (in.-lbs.) 4550 5244 Tip Taper 2:1, straight trailing edge CG chordwise 27.3% 26.7% Twist -10 deg Torsion frequency 5.8/rev Table 5: Blade Frequencies, SMART versus MD900 Analysis, cyc/rev Measured, Hz Table 2: Flap Characteristics Flat pitch, vacuum Free-free blade Radial station 150 ñ 186 in Mode SMART MD900 Mode SMART MD900 1C 0.58 0.59 Span length 36 in 1F 1.05 1.05 1F 6.56 6.63 Chord length 3.5 in (c + c ) f o 2F 2.8 2.7 2F 18.1 18.9 Hinge location 75% of blade chord 2C 4.4 4.5 3F 36.8 36.6 Flap twist axis 1.0 in aft of flap LE 3F 4.6 4.7 1C 40.8 41.1 Control horn length 0.75 in 1T 6.4 6.0 1T 69.7 68.5 Max. flap angle ± 6 deg 1TEF 95.8 -- Flap weight 1.26 lbs Table 6: Active flap test conditions Active Flap Control Tip Adv Tip Shaft Angle Blade Objective Condi Velocity*, Advance Mach # Mach # uncorrected Loading, Harmonic Amplitude A, Phase tion V, kt Ratio*, Ï MT MAT · , deg CT/ Û Number, n deg [V] ˆ , deg Validation 1 123 0.3 0.805 -9.4 0.08 5 0, 2 90 0.3 2 123 0.805 -9.4 0.08 3 0, 2 60 3 155 0.375 0.852 -9.3 .07, (.075) 5 0, (1) 180 4 83 0.2 0.746 0.9 0.075 2 & 5 0, 2 & 1 240 & 330 BVI Noise 1 62 0.15 0.623 4 0.075 2,3,4,5 1 - 2 sweep (RPM=392) 2 68 0.165 0.617 1.8 " 3,4 1 - 2 " 3 82 0.2 0.623 2 " 2,3,4,5 1 - 2.5 " Inplane Noise 1 124 0.3 0.623 0.809 -9.1 0.075 2,3,4,5 1 - 2 " Vibration 1 82 0.2 0.623 2 0.075 2,3,4,5,6 250V sweep 2 " " " " " Closed Loop: NF 5,1,1-5P; RM 5,1,1-5P 3 124 0.3 " -9.1 " 2,3,4,5,6 250V sweep 4 " " " " " Closed Loop: 1-5P, 10P; PM 1-5P Control Power 1 82 0.2 0.623 -5.5 0.075 0,1 -3 to 3 0,90 2 " " " 2 " " " " 3 124 0.3 " -9.1 " " " " Rotor Smoothing 1 0 0 0.623 -10 0.028 0 A1=-3 to +3 2 82 0.2 " 2 0.075 " " 3 124 0.3 " -9.1 " " " 4 " " " " " " A2=-3 to +3 Performance 1 82 0.2 0.623 2 0.075 n/a 0V, 0 2 124 0.3 " -9.1 0.075 " 0V, 0 3 " " " " 0.075 2 0, 1.5 sweep 4 " " " " 0.09 " " " NF - normal force, RM - roll moment, PM - pitch moment, [alternate units], (target condition, not achieved) * Condition is set to value shown in large type font 34 ft 34 ft Rod End Bearing Flexural Linkage 11.92 ft 11.92 ft (3.38 m) (3.38 m) 12.0 ft 12.0 ft 9.17 ft 9.17 ft (3.65 m) (3.65 m) (2.80 m) (2.80 m) Flap Access Cover Inboard Actuator Figure 1e: SMART Blade, flap, actuator cross-section 32.33 ft 6.53 ft 32.33 ft 6.53 ft (9.87 m) (1.99 m) (9.87 m) (1.99 m) 1.25 ft 1.25 ft (0.38 m) (0.38 m) Figure 1a: MD900 Explorer Rotor Data Rotor Data Rotor Rotor Flap Flap Balance Balance Actuator Flap Main Main Strut Strut Blade Tail Sting Tail Sting Outrigger Arm Outrigger Arm Motor Gearbox Motor Gearbox Figure 1b: SMART blade with embedded piezoelectric actuator and trailing edge flap Rotor Data Rotor Data Test Test Stand Stand Tripod Support Tripod Support Sled Sled Flap Horn, Flap Horn, Figure 2: System integration test at whirl tower Linkage Linkage Intermediate Intermediate Hinges Hinges Rotor Data Rotor Data Flap Flap Rotor Balance Rotor Balance In-plane Mic. In-plane Mic.

Main Strut Main Strut Flap Flap Traversing Traversing Mic. Array Mic. Array Figure 1c: SMART rotor blade on whirl tower Inboard X-Frame Actuator, Assembled Inboard X-Frame Actuator, Assembled Gearbox Gearbox Load Link Load Link BVI Mic. BVI Mic.

Motor Motor Piezo Stack Column Piezo Stack Column Slip Ring Slip Ring Tail Strut Tail Strut Port Port Starboard Strut Starboard Strut Flexure Flexure Mount Mount Figure 3a: SMART rotor in the NFAC 40- by 80- X-Frame Actuator, Frames X-Frame Actuator, Frames Lower view shown Lower view shown foot wind tunnel (looking upstream) Outboard X-Frame Actuator, Disassembled Outboard X-Frame Actuator, Disassembled Figure 1d: 2x-Frame piezoelectric actuator 3.5 2.5 1.5 0.5 Flap Deflection, deg Ÿ = 0 -0.5 -1 0 20 40 60 80 100 120 140 160 Tunnel Speed, knots Figure 3b: Close-up view of the SMART rotor, blade, Figure 5a: Active flap deflection versus speed, open- and flap in the tunnel loop flap with 0 Volt applied 0.25 M14 M14 0.2 In-plane In-plane 0.15 Microphones Microphones 0.1 0.05 M15 M15 M13 M13 -0.05 M16 M16 · = -10 -0.1 Flap Deflection, deg Ë = 4 -0.15 · = 4 deg -9.3 -9.4 0.9 -0.2 Traverse Mics Traverse Mics M12 M5 M12 M5 -0.25 0 20 40 60 80 100 120 140 160 M4 M1 M4 M1 Tunnel Speed, knots Fixed Mics Fixed Mics Figure 5b: Active flap deflection versus speed, closed-loop flap position control with 0deg command a) SMART rotor in 40- by 80-foot anechoic test section 0.2 and microphone configuration (looking downstream) 2deg sin(2P+240) + 2deg sin(5P+90) 1deg sin(5P+330) 0.15 2deg sin(3P+60) 0.1 RMS Error 0.05 0 Mean Flap Deflection, deg -0.05 1.5deg sin(4P+30) -0.1 20 40 60 80 100 120 140 Tunnel Speed, knots b) Microphone layout (top view) Figure 5c: Active flap mean deflection and RMS error versus speed, closed-loop flap position control Figure 4. Acoustic test setup: a) SMART rotor and with four harmonic commands microphone configuration, b) microphone layout High High High 1.5 Pitchcase Torsion Pitchcase Torsion Magnitude Magnitude 0.5 Low Low Low -0.5 Flap Deflection, deg -1 B lade 1 Blade 2 B lade 3 Blade 4 -1.5 Pitchcase Torsion Pitchcase Torsion B lade 5 Command Phase Phase -2 0 90 180 270 360 Azimuth of kth Blade - ¯ k, deg Figure 6a: Active Flap deflection versus azimuth at 83 knots, · = 0.89 deg, uncontrolled (0V) 5.8/rev 5.8/rev 5.8/rev 5.8/rev Figure 7: Pitchcase torsion magnitude (power spectral map, log scale) and phase (deg) from active -1 flap frequency sweep; 200V collective, 0.2-80 Hz Flap Deflection, deg log sweep, hover, · = -10 ˚ , collective pitch = 4 ˚ -2 Blade 1 Blade 2 Blade 3 Blade 4 Blade 5 Command FlapAngle - Act Volt a) mu=0.2, alfa=2 -3 FlxBm FB - FlapAngle 0 90 180 270 360 FlxBm T - Flap Angle Azimuth of kth Blade - ¯ k, deg Figure 6b: Active flap deflection versus azimuth at 83 knots, · = 0.89 deg, ‰ = 1.5deg sin(2 ¯ +240 ) + f 1deg sin(5 ¯ +330) -45 -90 Phase, deg 2.5 -135 -180 0 1 2 3 4 5 6 1.5 Harmonic Excitation, per rev 0.5 FlapAngle - Act Volt b) mu=0.3, alfa=-9.1 FlxBm FB - FlapAngle -0.5 FlxBm T - Flap Angle -1 Flap Deflection, deg -1.5 -2 1 2 -2.5 3 4 -45 5 Cmd 0 90 180 270 360 -90 Azimuth of kth Blade - ¯ k, deg Phase, deg -135 -180 Figure 6c: Active flap deflection versus azimuth at 0 1 2 3 4 5 6 123 knots, · = -9.4 deg, ‰ = 2 deg sin(5 ¯ +90 ) f Harmonic Excitation, per rev Figure 8: Phase relationships for excitation at rotor speed multiples; active flap vs voltage and flexbeam flap bending and torsion vs active flap Figure 9: Effect of active flap excitation on BVISPL for Condition 1 (  = 0.150,  = +4.0 ˚ ) at microphone M7 (traverse station: ñ 120) Figure 10a: BVISPL contours for baseline (0 deg) and active flap actuation (1.5 ˚ /4P/30 ˚ ) for Condition 1 (  = 0.150,  = +4.0 ˚ ) Figure 10b: Acoustic pressure time histories and spectral data for baseline (0 deg) and active flap actuation (1.5 ˚ /4P/30 ˚ ) for Condition 1 (  = 0.150,  = +4.0 ˚ ) at microphone M 7 (traverse station: -120) Baseline 109.76 dB 2P0 3P250 4P180 -3 Noise Reduction -6 LFSPL Change from Baseline, dB -9 0.0 0.5 1.0 1.5 2.0 2.5 Active Flap Amplitude, deg Figure 11b: Effect of active flap excitation frequency Figure 11a: Effect of active flap excitation frequency and amplitude for ì best î phase on LFSPL at and phase on LFSPL at microphone M13;  = 0.3,  microphone M13;  = 0.3,  = -9.1 ˚ = -9.1 ˚ Figure12: Measured acoustic time histories and frequency spectra at ì best î phase, ì best î amplitude conditions (microphone M13): a) 1.5 ˚ /2P/0 ˚ , b) 2.0 ˚ /3P/250 ˚ , c) 1.3 ˚ /4P/180 ˚ ;  = 0.3,  = -9.1 ˚ (OASPL ñ overall SPL; th MFSPL ñ medium frequency SPL, > 6 blade-passing harmonic) 100 500 4P 5P 4P 50 400 6P 0deg 5P 6P -200 -100 0 100 200 300 30deg -50 Baseline Baseline -400 -200 0 200 400 600 800 -100 100 -100 30deg -150 -200 -600 -400 -200 0 200 400 600 -300 Side - Sin 5P, lb Thrust - sin 5P, lb Axial - Sin 5P, lb -200 -100 Baseline -400 0deg -250 -200 -500 4P 0deg -300 30deg 5P -600 -300 6P Thrust Axial Side -350 -700 -400 Thust - Cos 5P, lb Side - Cos 5P, lb Axial - Cos 5P, lb 8000 6000 4P 4P 30deg 4P 5P 5P 600 5P 6P 6P 6P 0deg Baseline Baseline 500 4000 2000 -10000 -8000 -6000 -4000 -2000 0 2000 4000 Baseline Pitch - Sin 5P, in-lb Roll - Sin 5P, in-lb -2000 -6000 -4000 -2000 0 2000 4000 Vibration Index Magnitude- 5P -2000 -4000 Vibration Index 30deg Roll Pitch 0deg 0 -4000 -6000 0 90 180 270 360 Pitch - Cos 5P, in-lb Roll - Cos 5P, in-lb Phase, deg Figure 13: Vibratory hub loads (5P) for baseline case (0V) and open-loop active flap control phase with 250V amplitude at 4, 5, and 6P;  = 0.3, · = -9.1 ˚ Figure 14a: Fast Fourier transform (spectrum) of balance normal force for open-loop case (0V); Figure 14b: Estimated transfer function from  = 0.2, · = 2 ˚ actuator voltage to balance normal force; 200V collective, 0-80 Hz linear sweep;  = 0.2, · = 2 ˚ Figure 14c: Fast Fourier transform (spectrum) of balance normal force for closed-loop active flap control of normal force 1-5P;  = 0.2, · = 2 ˚ 1200 1200 1200 700 700 700 Normal Force (Lift, lb) at mu=0.30, alfa=-9.1 Run 33, 39, 42 Normal Force (Lift, lb) at mu=0.30, alfa=-9.1 Run 33, 39, 42 Normal Force (Lift, lb) at mu=0.30, alfa=-9.1 Run 33, 39, 42 Vibration Index at mu=0.30, alfa=-9.1 Run 33, 39, 42 Vibration Index at mu=0.30, alfa=-9.1 Run 33, 39, 42 Vibration Index at mu=0.30, alfa=-9.1 Run 33, 39, 42 Vibration Index Vibration Index Normal Force, lb Normal Force, lb 0V Baseline and CTHHC control 0V Baseline and CTHHC control 0V Baseline and CTHHC control 0V Baseline and CTHHC control 0V Baseline and CTHHC control 0V Baseline and CTHHC control 600 600 600 1000 1000 1000 -26% -29% -26% -29% -26% -29% -98% -98% -98% -98% -73% -73% -73% -73% 0V Average NF 1-5P 0V Average NF 1-5P 0V Average NF 1-5P 500 500 500 0V Average NF 1-5P 0V Average NF 1-5P 0V Average NF 1-5P 800 800 800 NF1-5P,T=10 NF1-5P,T=1 NF1-5P,T=10 NF1-5P,T=1 NF1-5P,T=10 NF1-5P,T=1 400 400 400 NF1-5P,T=10 NF1-5P,T=1 NF1-5P,T=10 NF1-5P,T=1 NF1-5P,T=10 NF1-5P,T=1 600 600 600 NF 10P,T=5 NF 10P,T=1 NF 10P,T=5 NF 10P,T=1 NF 10P,T=5 NF 10P,T=1 NF 10P,T=5 NF 10P,T=1 NF 10P,T=5 NF 10P,T=1 NF 10P,T=5 NF 10P,T=1 300 300 300 -80% -80% -80% -80% 400 400 400 200 200 200 -76% -76% -76% -98% -98% -98% -98% 200 200 200 100 100 100 0 0 0 0 0 0 CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P CYC RMS 1P 2P 3P 4P 5P 6P 7P 8P 9P 10P Figure 15: Vibratory hub normal force and vibration index for baseline case (0V) and closed-loop active flap control of normal force (NF);  = 0.3, · = -9.1 ˚ -25 2 fwd Swashplate 1deg Coll up -20 Flap IBC -15 + Roll right 3deg 1P0 -10 3deg 0P + Pitch aft -5 1.5deg Flap Coll 1 right 1 left Swashplate 3deg 1P90 Thrust Force, 1000 lb 3deg 1P270 10 Pitch Moment, 1000 in-lb 2 Flap VSP Flap IBC + Collective up 1 aft + TE Flap down 3deg 1P180 2 aft -4 -3 -2 -1 0 1 2 3 4 -25 -20 -15 -10 -5 0 5 10 15 20 25 Control Input, deg Roll Moment , 1000 in-lb Figure 16: Control power: steady hub load changes from swashplate and active flap inputs;  = 0.3, · = -9.1 ˚ a) Thrust from collective inputs, and b) Hub moments from cyclic inputs 124 Knot 82 Knot df2=-3 124kt, alfa=-9.1 Hover df1=3 -1 -1 df1=3 df1=3 -2 ROLL - Cos 1P, 1000 in-lb Thrust/100 - mean df1=-3 -2 Roll/1000 - sin 1P -3 Hub Loads Change from Baseline, lb, in-lb Pitch /1000- cos 1P df2=3 df1=-3 -4 -3 -4 -3 -2 -1 0 1 2 3 4 -12 -11 -10 -9 -8 -7 -6 Flap Deflection, deg ROLL - Sin 1P, 1000 in-lb Figure 17: Rotor smoothing: hub load changes from steady inputs of -3 to +3 deg on a single active flap.

a) thrust (mean), roll moment (sin1P), and pitch moment (cos1P) relative to baseline with flap 1;  = 0.3, · = -9.1 ˚ ; b) roll moment (1P) with flap 1 or flap 2 (df1, df2);  = 0, · = -10 ˚ ;  = 0.2, · = 2 ˚ ;  = 0.3, · = -9.1 ˚ 0.10 9 2P (0.075) 2P (0.090) Baseline (0.075) Baseline (0.090) 2P Average (0.075) 2P Average (0.090) 0.08 8.5 0.06 Û 0deg Lift / Drag CLR / 0 V 0.04 Baseline results before/after 0.02 2P phase sweep plotted at 10/350deg phase for clarity 7.5 0 90 180 270 360 0.00 0.000 0.002 0.004 0.006 0.008 0.010 Active Flap 2P Phase Angle, deg Û CP / Figure 18b: Rotor performance with active flap Figure 18a: Rotor performance with open-loop (0V) and inputs of 0 deg and 1.5 deg sin(2P +  ); active flap position control (0 deg);  = 0.3, · = -9.1 ˚  = 0.3, · = -9.1 ˚ , C /  = 0.075, 0.09 T