Preliminary Testing of a Mach-Scaled Active Rotor Blade with a Trailing Edge Servo-Flap

Steven R. Hall, Massachusetts Institute of Technology 77 Massachusetts Ave Cambridge, MA 02139-4307 USA, Office: 33-213, E-mail: Steven_Hall@mit.edu , Telephone:, 617-253-0869, Fax: 617-253-7397

Eric F. Prechtl Massachusetts Institute of Technology 77 Massachusetts Ave Cambridge, MA 02139-4307 USA, Office: 37-391, E-mail: prechtl@mit.edu ; Telephone:, 617-253-3267, Fax: 617-258-5940

Keywords: helicopter, rotor blade, servo-flap, actuator, piezoelectric

ABSTRACT

The results of preliminary tests on an active helicopter rotor blade are presented. The blade, a Mach-scaled model of a CH-47D helicopter blade, has a discrete piezoelectric actuator embedded within the spar that controls a trailing edge flap via a pushrod. Ultimately, the blade will be tested on a helicopter rotor hover stand at MIT. In this paper, we describe the tests performed prior to hover testing. First, the actuator was tested on the bench to determine its control authority and frequency response. Second, the actuator was tested on a shake table to simulate the out-of-plane accelerations that would be encountered in a full-scale helicopter in forward flight. Third, the actuator was embedded in the model blade, and its response to low-frequency sinusoidal actuation was obtained and compared to the bench test results. Finally, the frequency response of the actuator in the blade was determined using swept sine excitation. All test results indicate that the actuator should produce the desired level of control authority in the model-scale rotor.

1 INTRODUCTION

Helicopters are subject to high levels of vibration, mostly due to the unsteady aerodynamics of the main rotor, especially blade-vortex interaction. The vibration increases maintenance requirements, reduces pilot effectiveness, and reduces passenger comfort. Feedback control can be used to reduce this vibration. Early studies on helicopter rotor control considered the use of actuators that control the blade pitch through the swashplate, using either Higher Harmonic Control¹ (HHC) or Individual Blade Control² (IBC) methodologies.

During the last decade, there has been considerable interest in the use of on-blade actuation for helicopter rotor control for vibration and noise, for a number of reasons. Generally, HHC and IBC techniques require high bandwidth (on the order of N times the rotor frequency, where N is the number of rotor blades), and low blade pitch amplitude (about ±2 deg). On the other hand, the swashplate and the actuators that drive it are designed for maneuvering control, which is at a much lower frequency, and requires higher amplitudes. Also, the swashplate is flight-critical, whereas vibration control is not. Thus, there is some understandable reluctance to use the swashplate system for vibration control. Also, several studies point to the use of trailing-edge flaps as an effective means to control vibration.^3,4

Spangler and Hall⁵ first suggested the use of active materials, such as piezoelectric ceramics, for on-blade actuation. Piezoelectric materials are in some ways ideal candidates for this application, because of their inherently high bandwidth and stiffness. Spangler and Hall proposed using a bimorph ender element, cantilevered behind the blade spar towards the trailing edge, that would control a hinged trailing edge flap. Later, Hall and Prechtl⁶ improved on that design with a tapered bender, and a flap connected to the bender and airfoil with flexures. Chopra et al.^7-10 have also investigated the use of piezoelectric bender actuators on Froude-scaled model rotors, and have begun testing of a Mach-scaled model rotor with these actuators. Fulton and Ormiston^11,12 used similar actuators on a Froude-scaled blade to determine the response of a model rotor to trailing-edge flap control.

Unfortunately, bender actuators have a number of limitations. Their placement in the rear of the airfoil imposes a severe weight penalty on the blade, since the bender mass must be counter-balanced with weight at the leading edge for aeroelastic stability. More importantly, the energy density of current piezoelectric materials, using the d₃₁ effect required for bending actuation, is about an order of magnitude too low for effective flap-actuation in full scale or Mach-scale blades. The energy density of piezoelectric materials using the d₃₃ effect is about sufficient for rotor control. However, the challenge has been to develop an efficient amplification mechanism to convert the small strains these materials produce (1000-2000 mstrain peak-to-peak) to usable motion. A number of researchers have developed actuators that amplify the motion of piezoelectric or magnetostrictive stacks.^13-17

In this paper, we report on ongoing research at MIT to develop such an actuator, known as the X-Frame Actuator, for use in full-scale or Mach-scale model rotors. Our research results to date are for model rotors. As noted by Friedmann,¹⁸ Mach scaling is appropriate for tests of model rotors using active materials for actuation. Indeed, the important nondimensional parameter that relates to the use of active materials is the ratio of dynamic pressure to material modulus.

2 THE X-FRAME ACTUATOR CONCEPT

The design considerations and in-depth description of the X-Frame actuator were presented in two previous papers by the authors.^19,20 The basic actuator concept is shown in Figure 1. The X-Frame actuator consists of two active material stacks and two criss-crossed frames. A rotational degree of freedom serves as the interface between the two frames at one end, called the pivot end. As the stacks extend, the frames rotate with respect to each other about the pivot end of the actuator. The shallow angle between the stacks and the frames geometrically amplifies the stroke at the opposite end, called the output end. Restraining one of the frames at the output end of the actuator causes the opposite frame to undergo the entire motion.

Refinements were made to the basic actuator design to facilitate its incorporation into the Mach scaled rotor blades. These refinements were described in depth in a previous paper.²⁰ A brief overview of the design features is provided here. The rotor blade actuation system is shown in Figure 2.

The stacks used in the actuator can be composed of any type of active material. The stacks used in this research were purchased from EDO Corporation. The active material is designated by EDO as EC-98, which is a lead magnesium niobate-lead titanate (PMN-PT) composition. Large radius spherical end caps were used on each stack to allow relative stack/frame rotational motions. Small alignment pins were also included to maintain stack alignment despite large transverse accelerations.

The rotational degree of freedom between the frames is realized through a flexural connection at the pivot end of the actuator. At the output end, the outer frame includes guide arms between which the inner frame can slide. These guides both keep the two frames aligned with each other and help react transverse forces on the inner frame.

Two restraints are used to mount the actuator within the rotor blade spar. The outboard end of the actuator is bolted to the outboard spar restraint, constraining that end in the flapwise, chordwise and spanwise directions. An additional flexure, oriented in the spanwise direction, is used between this bolted mounting point and the actuator frames to equilibrate the centrifugal loads on the two stacks. The inboard side of the actuator is restrained in the chordwise and flapwise directions by a sliding interface between the outer frame and the inboard spar restraint. The actuator is free to move in the spanwise direction at the inboard side to allow for stack expansion and contraction.

The actuator forces and deflections are transferred to the servo-flap by a control rod. Threads at either end of the control rod are used to connect it to the inner frame and a clevis at the trailing edge. The clevis/control rod attachment is accomplished using an 0-80 threaded engagement. This engagement acts as the adjustment/trimming mechanism for the servo-flap. A pinned connection is used to connect the clevis to the flap horn, which is bonded to the servo-flap during its initial cure. A spherical protrusion on the pin is used to allow for small relative flapwise and chordwise rotations between the clevis and the horn.

A graphite reaction rib is laid up in the composite rotor blade, oriented in the chordwise direction. This rib serves to react actuation forces, as the blade structure is not normally designed to be stiff in the aft fairing.

A pre-stress wire runs through the entire length of the servo-flap. It interfaces with the outboard side of the flap through two keys and with a pre-stress wire flange at the inboard side. This wire has three functions; the flap rotational shaft, the flap thrust bearing, and the pre-stress element for the actuator. A detailed description of how this wire performs these three functions is presented in the previous paper.²⁰

A slotted flap was used for the servo-flap design. The airfoil/flap contour was designed using MSES, a multi-element, viscous, compressible airfoil analysis code. MSES is a derivative of the single element code, ISES.^21-23 The code was also used to compare the redesigned section to the original CH-47 section. While the slotted flap does develop slightly higher drag on the blade, it requires substantially lower hinge moments for operation than a plain flap, which would have a lower drag penalty. The aeroelastic properties of a slotted flap are also beneficial, since it is easier to mass balance the flap for stability. Finally, because the slotted flap pivots about an axis very close to its centroid, the inertia is smaller, which results in greater actuation bandwidth.

3 EXPERIMENTAL RESULTS

In this section, we present experimental results with the actuator described above. Three types of experiments were performed. First, the actuator was tested on a shake table, to simulate the out-of-plane loads that would be encountered by a blade-mounted actuator due to unsteady aerodynamic loads. Second, the actuator was tested in the model rotor blade at low frequency, to compare its performance to bench top experiments. Finally, the actuator was tested in the blade with a swept-sine excitation, to determine its frequency response. The experimental setup and results are described below:

3.1 Shake Tests

One issue of concern is whether a blade-mounted actuator is sensitive to blade accelerations. Such a sensitivity can take two forms. First, the actuator may have induced deflections due to acceleration, which would produce an undesirable (perhaps even destabilizing) coupling between blade motion and flap deflection. Second, blade accelerations may cause the actuator to have reduced performance. For example, increased friction due to inertial forces might reduce the output deflection. In a typical rotor, the highest (unsteady) accelerations by far occur in the out-of-plane direction. Therefore, the X-Frame actuator was tested for acceleration sensitivity by simultaneously operating the actuator while shaking in the out-of-plane direction at the frequencies and amplitudes that would occur in forward flight, appropriately scaled to the model rotor.

To perform the tests, an apparatus was constructed that allowed the actuator to be operated on a shake table, while simulating as much as possible the conditions that the actuator experiences in the blade. In particular, inboard and outboard restraints were manufactured identical to those in the blade, except for flanges necessary to mount them to the apparatus. To simulate the load of the flap, the pushrod was connected to a steel load flexure with stiffness matching the expected aerodynamic hinge stiffness of the flap. Note that the inertia of the flap was not incorporated into the apparatus, which has implications for the frequency response testing later.

The actuator was operated under four different acceleration amplitudes (7.1g, 14.5g, 43g, and 69g); four different acceleration frequencies (22.5 Hz, 45 Hz, 67.5 Hz, and 135 Hz); and four different actuation frequencies (3 Hz, 22.5 Hz, 45 Hz, and 67.5 Hz). In addition, the actuator was operated prior to and after shake testing with no acceleration. In all, 48 test runs were performed.

Figure 3 shows the results from two of the tests, which are roughly the two extreme cases. The plot on the left shows actuator excitation at 3 Hz, with no shaking. The plot on the right shows the result with the same excitation, and acceleration at 69g amplitude and 45 Hz frequency. In the two cases, the time responses of the actuator are nearly identical. The peak-to-peak amplitude of motion is the same, as is the amount of hysteresis due to material properties. The only slight difference visible is that in the case with shaking, there is a very small ripple visible, apparently at 90 Hz. We believe that this is due to the side members of the frames bending at 45 Hz, which produces foreshortening with fundamental frequency at 90 Hz. In any event, the ripple is small, and is only noticeable in the highest acceleration cases.

The results in Figure 3 are typical. That is, all 48 cases have similar time traces, except for a slight frequency-dependent amplitude variation, probably due to material properties. The results of all test cases are shown in Table 1.

Shake Amplitude (g)	Shake Frequency (Hz)	Actuation Frequency (Hz)
		3	22.5	45	67
		Actuator Deflection (mil)
0	0	15.30	14.65	14.50	14.51
7.1	22.5	15.05	14.44	14.30	14.28
	45.0	15.11	14.48	14.33	14.28
	67.5	15.12	14.51	14.33	14.29
14.5	22.5	15.05	14.44	14.27	14.22
	45.0	15.13	14.55	14.30	14.29
	67.5	15.16	14.56	14.34	14.31
	135.5	15.18	14.55	14.37	14.39
43	45.0	15.09	14.50	14.19	14.12
	67.5	15.09	14.52	14.31	14.31
69.0	45.0	15.08	7.01^a	7.09^a	6.95^a
0	0	15.10	14.48	14.33	14.28

Table 1. Shake Test Data. All cases were performed with stack voltage amplitudes between 390.5V and 396.5V amplitude. The deflection data in this table is linearly scaled to obtain the deflection at 400V amplitude.

Note: ^a For these cases, the leads to one of the two stacks in the actuator were inadvertently disconnected. As a result, the amplitude of the response is approximately one-half that of the other cases.

3.2 Quasistatic Response

Typically, active materials such as piezoelectric ceramics exhibit nonlinear behavior, such as hysteresis, as well as voltage-dependent nonlinearities (increasing or decreasing d₃₃). In addition, there can be other sources of nonlinear behavior, such as binding or friction in the actuator mechanism, or nonlinear gain. Thus, it is important to characterize the quasistatic response of the actuator.

Figure 4 shows the quasistatic responses of the actuator on the bench top driving a light load (to match that of the prestress wire in the blade), and in the rotor blade. The data in Figure 4a is normalized to yield equivalent flap deflection, so that it can be compared directly to Figure 4b. In Figure 4a, the hysteresis is moderate, and is primarily due to material effects. In contrast, friction in the flap hinge (i.e., the prestress wire) and clevis increase the hysteresis significantly, reducing the free deflection by about 4 deg. Under aerodynamic loads, the effect of friction will be less important, due to the increased hinge stiffness of the flap. We expect that we will lose only about 2 deg of motion due to friction at the aerodynamic design point.

The friction problem is largely a result of the difficulties associated with building at model scale. In particular, we were unable to locate suitable bearings for the hinge line, and as a result our hinge has metal-to-metal contact without a bearing.

Also, the friction problem could be reduced significantly in a configuration where the actuator preload is not applied through the flap itself.

3.3 Frequency Response

Finally, the transfer functions of the actuator and actuator/flap system were determined, for two cases, as shown in Figure 5.

Figure 5a shows the frequency response for the actuator alone in the shake test apparatus. The deflection of the actuator was measured using a strain gage attached to the load flexure, which had been calibrated using a laser interferometer. The response is very nearly second order, with natural frequency at about 650 Hz, and lightly damped. Also, there is slightly increased phase lag as compared to a second-order system, due to material hysteresis.

Figure 5b shows the response of the actuator in the blade with the flap attached. The flap deflection was determined by measuring the actuation deflection using a Hall effect sensor, which was calibrated by comparing the Hall effect sensor output voltage to flap angle measured using a laser light lever. The actuator was actuated using a 0-800 V sinusoidal excitation. Because of significant levels of friction, the response of the flap was more nonlinear than in the shake test apparatus. Hence, the data is plotted showing deflection at 800 V excitation, rather than as a transfer function.

Figure 5b exhibits two natural modes within the frequency range tested. The first mode is at about 140 Hz. The mode is at a lower frequency than in the shake test apparatus, which did not include the effect of the flap inertia. Also, note that this mode is more heavily damped, due to the friction effects. Friction and hysteresis effects also cause significant phase lag. The second mode is at about 180 Hz. This mode corresponds to the first torsional mode of the rotor blade for the test condition, i.e., with the root end free.

Note that the bandwidth of the actuator in the blade is greater than 6/rev, which is more than adequate for rotor control on the CH-47, which requires up to 4/rev actuation.

Finally, the flap was actuated at several frequencies, and videotaped while illuminated with a strobe light, in order to visualize the flap motion. Interested readers can view the video at http://web.mit.edu/srhall/www/blade.html.

4 CONCLUSIONS

The X-Frame actuator has been incorporated into a Mach-scaled model rotor system, and preliminary (non-rotating) tests demonstrate that the device has the expected control authority and bandwidth. The actuator has been shown to work well in the vibration environment expected in the full-scale rotor. Other tests, not reported on here, demonstrate that the actuator should work in the centrifugal field expected. Hover testing of the actuator should begin at MIT in April 1999.

ACKNOWLEDGEMENTS

This research was supported by DARPA under Contract Number F49620-95-2-0097 and MDA972-98-3-0001, monitored by Bob Crowe, Bill Coblenz, and Ephrahim Garcia. Additional support was provided by the Army Research Office, under contract DAAH04-95-0104, monitored by Gary Anderson. The authors would like to acknowledge Aaron Julin of MIT and Rich Bussom of Boeing Philadelphia for their support of the shake test and apparatus; Prof. Mark Drela of MIT for aiding with the aerodynamic analysis of the flap and airfoil section, as well as code for blade manufacturing; John Rodgers and Mads Schmidt for assistance with blade manufacture; and Terry Deane at Advanced Machining and Tooling, Inc. for machining the model scale actuator components.

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