This paper outlines the design, modeling and testing of a new class of tool intended for the treatment of crack-arrest holes to improve fatigue life. By integrating a stack of high-power piezoelectric elements in a compression caliper, this Piezoelectric Impact Compressive Kinetic (PICK) tool can be used to clamp very tightly on either side of an aluminum plug, which is inserted in a crack-arrest hole. Ultrasonic vibrations at high compression loads applied by the piezoelectric stack dynamically cold work both the aluminum plug and the inside of the crack-arrest hole. This paper describes the overall design of the tool, the configuration of the aluminum plug, and the effect of dynamic vibrations on the plug and on the surface of the crack-arrest hole. The system was driven at various resonance modes during the coldworking process. Several 3.2-mm (1/8-in.) thick steel specimens with 3.2-mm (1/8-in.) diameter crack-arrest holes were treated ultrasonically with the PICK tool. Dynamic fatigue tests showed that fatigue lives of the specimens was increased substantially as a result of the ultrasonic treatment. Microhardness and neutron diffraction testing confirmed that the tool induced high levels of cold working at the edge of the hole and increased the grain density, with a regular decay as a function of distance from the edge of the hole.
This paper outlines the design, fabrication and testing of a new, high performance piezoelectrically driven aircraft flutter
test vane. This flutter test vane utilizes low-net passive stiffness (LNPS) actuator configurations to produce deflection
amplification ratios on the order of 5:1 while maintaining full blocked moment generation capability. With an order of
magnitude lower weight than conventional vanes, the LNPS flight flutter test vane is capable of producing larger
amplitude structural deflections with smaller force levels because vane forcing waveforms, frequencies and phasing can
be very exactingly controlled with respect to each other. The paper covers the fundamental driving theories behind the
device, actuator geometry, test article layout, fabrication and testing. This device was wind tunnel tested at airspeeds up
to 110 ft/s with excellent correlation between theory and experiment. Experimental tests show an improvement in
angular deflection and delta lift forces from approximately ±1.8 deg. and 0.45 lbf to ±8.5 deg. and 1.45 lbf, respectively.
The flutter test vane consumes only 1W of peak power at max. actuation frequency, drastically reducing the impact of
electrical power supply lines on the modal mass of the wing. This paper describes the modeling, testing and evaluation
of the adaptive flutter test vane and quantifies the implications on the current state of flight flutter testing.
This paper outlines a new class of piezoelectric flight control actuators which are specifically intended for use in guided
hard-launched munitions from under 5.56mm to 40mm in caliber. In March of 2011, US Pat. 7,898,153 was issued,
describing this new class of actuators, how they are mounted, laminated, energized and used to control the flight of a
wide variety of munitions. This paper is the technical conference paper companion to the Patent. A Low Net Passive
Stiffness (LNPS) Post Buckled Precompressed (PBP) piezoelectric actuator element for a 0.40 caliber body, 0.50 caliber
round was built and tested. Aerodynamic modeling of the flight control actuator showed that canard deflections of just
±1° are more than sufficient to provide full flight control against 99% atmospherics to 2km of range while maintaining
just 10cm of dispersion with lethal energy pressure levels upon terminal contact. Supersonic wind tunnel testing was
conducted as well as a sweep of axial compression. The LNPS/PBP configuration exhibited an amplification factor of
3.8 while maintaining equivalent corner frequencies in excess of 100 Hz and deflection levels of ±1°. The paper
concludes with a fabrication and assembly cost analysis on a mass production scale.
The paper begins with a brief historical overview of pressure adaptive materials and structures. By examining avian
anatomy, it is seen that pressure-adaptive structures have been used successfully in the Natural world to hold structural
positions for extended periods of time and yet allow for dynamic shape changes from one flight state to the next. More
modern pneumatic actuators, including FAA certified autopilot servoactuators are frequently used by aircraft around the
world. Pneumatic artificial muscles (PAM) show good promise as aircraft actuators, but follow the traditional model of
load concentration and distribution commonly found in aircraft. A new system is proposed which leaves distributed
loads distributed and manipulates structures through a distributed actuator. By using Pressure Adaptive Honeycomb
(PAH), it is shown that large structural deformations in excess of 50% strains can be achieved while maintaining full
structural integrity and enabling secondary flight control mechanisms like flaps. The successful implementation of
pressure-adaptive honeycomb in the trailing edge of a wing section sparked the motivation for subsequent research into
the optimal topology of the pressure adaptive honeycomb within the trailing edge of a morphing flap. As an input for the
optimization two known shapes are required: a desired shape in cruise configuration and a desired shape in landing
configuration. In addition, the boundary conditions and load cases (including aerodynamic loads and internal pressure
loads) should be specified for each condition. Finally, a set of six design variables is specified relating to the honeycomb
and upper skin topology of the morphing flap. A finite-element model of the pressure-adaptive honeycomb structure is
developed specifically tailored to generate fast but reliable results for a given combination of external loading, input
variables, and boundary conditions. Based on two bench tests it is shown that this model correlates well to experimental
results. The optimization process finds the skin and honeycomb topology that minimizes the error between the acquired
shape and the desired shape in each configuration.
This paper outlines an investigation using shape memory alloy (SMA) filaments to drive a flight control system with
precision control in a real flight environment. An antagonistic SMA actuator was developed with an integrated
demodulator circuit from a JR NES 911 subscale UAV actuator. The SMA actuator was installed in the fuselage of a 2m
uninhabited aerial vehicle (UAV) and used to control the rudder through slips and coordinated turns. Rudimentary design
envelope modeling was performed to verify peak-to-peak throw levels and airload resistance capabilities. The actuator was capable of 20 degrees of positive and negative deflection and was capable of 7.5 in-oz (5.29N-cm) of torque at a bandwidth of 2.8Hz.
Sensor elements employing fine filaments are often vulnerable to particulate fouling when used in certain
operational field conditions. Depending on the size, attraction level, thermal and electrical conduction and charge
accumulation properties of the particles, erroneous readings can be easily generated in such "dirty" environments. This
paper describes the design, development and testing of an ultrasonic system which dynamically rejects highly tenacious
electrostatically charged particles of a wide variety of sizes and even water. The paper starts with a brief introduction to
the field of acoustic vector sensing, outlining its outstanding characteristics and history. Operational challenges
including a statistical analysis of typical Middle-Eastern wind-blown desert sand and charge density are laid out. Several
representative subscale hot-wire filaments were fouled with calibrated dust representing desert sand. The fouled elements were then exposed to airflows of 13 ft/s (4m/s) and showed highly erratic, shifted conduction levels with respect to baseline (clean) levels. An ultrasonic cleaning system was designed specifically resonate the filament and cantilever so as to mechanically reject foulants. When operated at resonance, the ultrasonic cleaning system showed 98.6% particulate rejection levels and associated restoration of uncorrupted filament resistance levels to within 2% of baseline resistance measurements. The study concludes with an assessment of such cleaning techniques in various environments.
The feasibility of piezoelectric-based Adaptive-Impedance Composites (AIC) as a method of protecting aircraft
equipment from lightning strike events and the resultant High-Intensity Radiated Fields (HIRF) was investigated.
Classical Laminated Plate Theory (CLPT) and sheet vibration theory were applied to analytically derive the performance
of the AIC. Multiple prototypes were built for high voltage testing which revealed closed- to open-circuit switching as
fast as 77 μs. It was observed that slight geometric variations of the AIC strongly influenced the activation voltage. The
voltage necessary to trigger the 85mm long, 3rd generation AIC's impedance could be set between 10 and 60 V. The test
data and the analytical predictions were compared with the lightning strike data gathered by ONERA. The comparison indicated the AIC switching speed was over 30 times faster than the necessary minimum to shield typical avionics and flight control mechanisms from lightning-strike induced electrical eddy currents and HIRF.
This paper is centered on a new actuation philosophy executed on an old rotor design. An adaptive rotor employing
twist-active piezoelectric root actuators was used as a testbed to investigate the new branch of structural mechanics
devoted to low- and zero-net passive stiffness (ZNPS) structures. One of the more common methods to achieve zero net
passive stiffnesses in structures is to employ "negative" springs: that is, mechanisms which when combined with the
baseline structure null the passive stiffness of the total structural element. This paper outlines the application of such a
system via a Post-Buckled Precompression (PBP) technique at the end of a twist-active piezoelectric rotor blade
actuator. The basic performance of the system is handily modeled by using laminated plate theory techniques. A dual
cantilevered spring system was used to increasingly null the passive stiffness of the root actuator along the feathering
axis of the rotor blade. As the precompression levels were increased, it was shown that corresponding blade pitch levels
also increased. The PBP cantilever spring system was designed so as to provide a high level of stabilizing pitch-flap
coupling and inherent resistance to rotor propeller moments. Experimental testing showed pitch deflections increasing
from just 8° peak-to-peak deflections at 650 V/mm field strength to more than 26° at the same field strength with design
precompression levels. Dynamic testing showed the corner frequency of the linear system coming down from 63 Hz (3.8/rev) to 53Hz (3.2/rev). Thrust coefficients manipulation levels were shown to increase from 0.01 to 0.028 with increasing precompression levels. The paper concludes with an overall assessment of the actuator design and conclusions on overall feasibility.
This paper discusses the modeling, design and performance of SMA actuated Post-Buckled Precompressed (PBP) plates.
The paper begins with an outline of past approaches to PBP actuation and control which have mostly been centered on
piezoelectric actuators. To skirt problems associated with tensile failure on convex faces of PBP actuators, SMA
filaments are used. Because SMA's are far less sensitive to tensile failure, far greater deflections and work outputs are
both predicted and experimentally measured. Unlike conventional SMA actuated plates, the total moment generation
capability and deflections can be simultaneously amplified, indicating a substantial increase in total work output.
Experiments showed that the SMA actuators are capable of tip deflection of up to 45° and that the Post-Buckled Precompressed Mechanism improves tip rotation up to 40% compared to conventional antagonistically actuated SMA plates.
A new type of adaptive structure is presented that relies on pressurized honeycomb cells that extent a significant
length with respect to the plane of the hexagons. By varying the pressure inside each of the cells, the stiffness can
be altered. A variable stiffness in combination with an externally applied force field results in a fully embedded
pressure adaptive actuator that can yield strains well beyond the state-of-the-art in adaptive materials. The
stiffness change as a function of the pressure is modeled by assigning an equivalent material stiffness to the
honeycomb walls that accounts for both the inherent material stiffness as the pressure-induced stiffness. A finite
element analysis of a beam structure that relies on this model is shown to correlate well to experimental results of
a three-point bend test. To demonstrate the concept of embedded pressure adaptive honeycomb, an wind tunnel
test article with adaptive flap has been constructed and tested in a low speed wind tunnel. It has been proven
that by varying the cell pressure the flap changed its geometry and subsequently altered the lift coefficient.
This paper is intended to introduce the international adaptive aerostructures community to the tremendous
opportunities these structures can bring to uninhabited aerospace systems. The paper starts with an
overview of the most critical classes of adaptive aerostructures for uninhabited aerial vehicles (UAVs) and
the materials which are used to drive them. The paper describes several classes of UAVs that take
advantage of the various kinds of these technologies. Adaptive aerostructures are shown to be integrated
into hovering, high speed, low speed and ultra-high performance UAVs. These ultra-high performance
UAVs are shown to significantly benefit from newly invented Post-Buckled Precompressed (PBP)
piezoelectric actuators. These UAVs are capable of hovering for extended periods of time as a helicopter in
gusty, windy, dusty environments, then pop up, converting and dashing out like a missile at several
hundred km/hr. The paper shows photos of ultra-high performance UAV launches from armored vehicles, a
battle-damage assessment exercise and a live fire sequence with 40mm munitions. The paper concludes
with a description of the Visual Signature Suppression (VSS) system which was employed on a 2m UAV
operating at several hundred meters above ground level. The VSS system was shown to reduce the visual
cross section to below 1.8cm2 which is the threshold for human aircraft observation. Accordingly, the VSS
equipped aircraft is said to "disappear" in mid flight.
The dynamic response of a new class of flight control actuators that rely on post-buckled precompressed (PBP)
piezoelectric elements is investigated. While past research has proven that PBP actuators are capable of generating
deflections three times higher than conventional bimorph actuators, this paper quantifies the work output and power
consumption under various axial loads, at various frequencies. An analytical model is presented that supports the
experimental findings regarding the increasing work output and natural frequency shift under increasing axial loads.
Furthermore, increasing axial loads shows an increase in open-loop piezoelectric hysteresis, resulting in an increasing
phase lag in actuator response. Current measurements show an electromechanical coupling that leads to power peaks
around the natural frequency. Increasing axial loads has no effect on the power consumption, while increasing the work
output by a factor of three, which implies a significant increase in work density over the piezoelectric material itself.
This paper describes a new class of flight control actuators using Post-Buckled Precompressed (PBP)
piezoelectric elements mounted within a transonic missile fin. These actuators are designed to produce
significantly higher deflection and force levels than conventional piezoelectric actuator elements. Classical
laminate plate theory (CLPT) models are shown to work very well in capturing the behavior of the free, unloaded
elements. A new high transverse deflection model which employs nonlinear structural relations is shown to
successfully predict the performance of the PBP actuators as they are exposed to higher and higher levels of axial
force, which induces post buckling deflections. A 6" (15.2cm) square rounded diamond transonic fin was made
with integral PBP actuator elements. Quasi-static bench testing showed deflection levels in excess of ±7° at rates
exceeding 21 Hz. The new solid state PBP actuator was shown to reduce the part count with respect to
conventional servoactuators by an order of magnitude. Power consumption dropped from 24W to 1.3W, slop
went from 1.6° to 0.02° and peak current draw went from 5A to 18mA. The PBP actuator was wind tunnel tested
and shown to possess no flutter, divergence or adverse aeroelastic coupling characteristics.
This paper describes a new class of flight control actuators using Post-Buckled Precompressed (PBP)
piezoelectric elements to provide much improved actuator performance. These PBP actuator elements are modeled
using basic large deflection Euler-beam estimations accounting for laminated plate effects. The deflection
estimations are then coupled to a high rotation kinematic model which translates PBP beam bending to stabilator
deflections. A test article using PZT-5H piezoceramic sheets built into an active bender element was fitted with an
elastic band which induced much improved deflection levels. Statically the bender element was capable of
producing unloaded end rotations on the order of ±2.6°. With axial compression, the end deflections were shown to
increase nearly 4-fold. The PBP element was then fitted with a graphite-epoxy aeroshell which was designed to
pitch around a tubular stainless steel main spar. Quasi-static bench testing showed excellent correlation between
theory and experiment through ±25° of pitch deflection. Finally, wind tunnel testing was conducted at airspeeds up
to 120kts (62m/s, 202ft/s). Testing showed that deflections up through ±20° could be maintained at even the highest
flight speed. The stabilator showed no flutter or divergence tendencies at all flight speeds. At higher deflection
levels, it was shown that a slight degradation deflection was induced by nose-down pitching moments generated by
separated flow conditions induced by extremely high angles of attack.
This paper presents the use of a new class of flight control actuators employing Post-Buckled Precompressed (PBP) piezoelectric elements in morphing wing Uninhabited Aerial Vehicles (UAVs). The new actuator relies on axial compression to amplify deflections and control forces simultaneously. Two designs employing morphing wing panels based on PBP actuators were conceived. One design employed PBP actuators in a membrane wing panel over the aft 60% of the chord to impose roll control on a 720mm span subscale UAV. This design relied on a change in curvature of the actuators to control the camber of the airfoil. Axial compression of the actuators was ensured by means of rubber bands and increased end rotation levels with almost a factor of two up to ±13.6° peak-to-peak, with excellent correlation between theory and experiment. Wind tunnel tests quantitatively proved that wing morphing induced roll acceleration levels in excess of 1500 deg/s2. A second design employed PBP actuators in a wing panel with significant thickness, relying on a highly compliant Latex skin to allow for shape deformation and at the same time induce an axial force on the actuators. Bench tests showed that due to the axial compression provided by the skin end rotations were increased with more than a factor of two up to ±15.8° peak-to-peak up to a break frequency of 34Hz. Compared to conventional electromechanical servoactuaters, the PBP actuators showed a net reduction in flight control system weight, slop and power consumption for minimal part count. Both morphing wing concepts showed that PBP piezoelectric actuators have significant benefits over conventional actuators and can be successfully applied to induce aircraft control.
This paper describes a new method for drag elimination and stall suppression via tangential synthetic jet actuators. This boundary layer control (BLC) method is shown to perform as well as continuous and normal synthetic jet BLC methods but without fouling difficulties, system-level complexity or extreme sensitivity to Reynolds number. Classical laminated plate theory (CLPT) models of the piezoelectric actuators were used to estimate diaphragm deflections and volume per stroke. A 12” (30.5cm) chord, 6” (15.3cm) span NACA 0012 profile wing section was designed with three unimorph 10 mil (254μm) thick, 3.25” (8.23cm) square piezoelectric diaphragm plenums and five 1 mil (25μm) thick stainless steel valves spaced from 15%c to the trailing edge of the airfoil. Static bench testing showed good correlation between CLPT and experiment. Plenum volume per stroke ranged up to 5cc at 500 V/mm field strength. Dynamic testing showed resonance peaks near 270 Hz, leading to flux rates of more than 60 cu in/s (1 l/s) through the dynamic valves. Wind tunnel testing was conducted at speeds up through 13.1 ft/s (4 m/s) showing more than doubling of Clmax. At low angles of attack and high flux rates, the airfoil produced net thrust for less than 4.1W of electrical power consumption.
This paper describes a new class of flight control actuators using Post-Buckled Precompressed (PBP) piezoelectric elements. These actuators are designed to produce significantly higher deflection and force levels than conventional piezoelectric actuator elements. Classical laminate plate theory (CLPT) models are shown to work very well in capturing the behavior of the free, unloaded elements. A new high transverse deflection model which employs nonlinear structural relations is shown to successfully predict the performance of the PBP actuators as they are exposed to higher and higher levels of axial force, which induces post buckling deflections. A proof-of-concept empennage assembly and actuator were fabricated using the principles of PBP actuation. A single grid-fin flight control effector was driven by a 3.5" (88.9mm) long piezoceramic bimorph PBP actuator. By using the PBP configuration, deflections were controllably magnified 4.5 times with excellent correlation between theory and experiment. Quasi-static bench testing showed deflection levels in excess of ±6° at rates exceeding 15 Hz. The new solid state PBP actuator was shown to reduce the part count with respect to conventional servoactuators by an order of magnitude. Power consumption dropped from 24W to 100mW, weight was cut from 108g to 14g, slop went from 1.6° to 0.02° and current draw went from 5A to 1.4mA. The result was that the XQ-138 subscale UAV family experienced nearly a 4% reduction in operating empty weight via the switch from conventional to PBP actuators while in every other measure, gross performance was significantly enhanced.
Visual signature suppression (VSS) methods for several classes of aircraft from WWII on are examined and historically summarized. This study shows that for some classes of uninhabited aerial vehicles (UAVs), primary mission threats do not stem from infrared or radar signatures, but from the amount that an aircraft visually stands out against the sky. The paper shows that such visual mismatch can often jeopardize mission success and/or induce the destruction of the entire aircraft. A psycho-physioptical study was conducted to establish the definition and benchmarks of a Visual Cross Section (VCS) for airborne objects. This study was centered on combining the effects of size, shape, color and luminosity or effective illumance (EI) of a given aircraft to arrive at a VCS. A series of tests were conducted with a 6.6ft (2m) UAV which was fitted with optically adaptive electroluminescent sheets at altitudes of up to 1000 ft (300m). It was shown that with proper tailoring of the color and luminosity, the VCS of the aircraft dropped from more than 4,200cm2 to less than 1.8cm2 at 100m (the observed lower limit of the 20-20 human eye in this study). In laypersons terms this indicated that the UAV essentially "disappeared". This study concludes with an assessment of the weight and volume impact of such a Visual Suppression System (VSS) on the UAV, showing that VCS levels on this class UAV can be suppressed to below 1.8cm2 for aircraft gross weight penalties of only 9.8%.
Although many subscale aircraft regularly fly with adaptive materials in sensors and small components in secondary subsystems, only a handful have flown with adaptive aerostructures as flight critical, enabling components. This paper reviews several families of adaptive aerostructures which have enabled or significantly enhanced flightworthy uninhabited aerial vehicles (UAVs), including rotary and fixed wing aircraft, missiles and munitions. More than 40 adaptive aerostructures programs which have had a direct connection to flight test and/or production UAVs, ranging from hover through hypersonic, sea-level to exo-stratospheric are examined. Adaptive material type, design Mach range, test methods, aircraft configuration and performance of each of the designs are presented. An historical analysis shows the evolution of flightworthy adaptive aerostructures from the earliest staggering flights in 1994 to modern adaptive UAVs supporting live-fire exercises in harsh military environments. Because there are profound differences between bench test, wind tunnel test, flight test and military grade flightworthy adaptive aerostructures, some of the most mature industrial design and fabrication techniques in use today will be outlined. The paper concludes with an example of the useful load and performance expansions which are seen on an industrial, military-grade UAV through the use of properly designed, flight-hardened adaptive aerostructures.
UAV's, UCAV's, miniaturized munitions and smart bombs have a variety of objectives clamoring for easement of weight/volume restrictions. These include anti-jam, explosive, servo control, electronics packaging, GPS and other required functions. The possibility of freeing up valuable real estate in the missile itself is most attractive for such applications. QorTek has developed the first self-contained high authority control surface to replace externally activated steering fins or canards. These flight actuation systems require only external control signal and power. Moreover, the technology easily scales to micro munitions. Because of their unique composite structure, these powerful solid-state devices offer exceptional performance in a durable package suitable for miniature munitions. The purpose of this paper is to discuss new breakthroughs in piezo-actuated technology that minimize vol./weight enabling a self-contained flight control actuation system that eliminates the need for servo controls. The presentation will focus on the new design that enables integration into high angular displacement actuation into a graphite epoxy fabricated RALA flight control actuator that can handle the aerodynamic loading conditions.
A new technique is presented for designing actuators for guided hard-launch adaptive munitions by using actuator and substrate strain limits, static analysis methods and matching the local actuator strains along its length by varying the width. This Load-Matched design technique leads to an exponential area distribution as a function of length which is contrasted against the conventional rectangular actuator shapes that have been used in all adaptive hard-launch munitions up till now. To demonstrate the viability of this new Load-Matched actuator design, ten 600mg, 100mm long rectangular
and ten identical mass and length, exponentially shaped, Load-Matched actuator specimens were designed and built to withstand the maximum possible accelerations. Predicted design static strain distributions are presented along with limits, showing that rectangular actuators exhibit a strong strain peak at the root while Load-Matched actuators have a much more even distribution and a gentle maximum near the middle. Shock table testing showed that the rectangular specimens were
predicted to fail at 3,500g's, but survived acceleration levels 9.5-
12% higher than expected (3,833 to 3,931g's). The exponentially shaped Load-Matched actuators proved that they could withstand shocks from 17 to 21% over the predicted failure acceleration level of 8,000g's (9,377 to 9,670g's).
This paper outlines the design, fabrication and testing techniques used for a new class of active wings. In support of recent efforts to shrink high authority, high speed aircraft flight controls for Micro Aerial Vehicles (MAVs), a new wing design was conceived. Quasi-vortex lattice methods were used to determine roll control power of unswept wings with (1) conventional ailerons, (2) active twist, and (3) a new method which employs root twist only. It was shown that linear wing twist provides 3.9 times more roll control than conventional ailerons. Root twist manipulations are shown to provide 1.6 times more control than twist. A comparison of control authority per unit actuator weight also shows that root actuators are more than 20% lighter than the other two systems. To drive the new root pitch actuator, a directionally attached piezoelectric (DAP) torque-late actuator was built. A laminated plate theory analysis of the 4' span, 3' chord DAP torque-plate showed good correlation between theory and experiment. The DAP torque plate was integrated into a 30' span, 4' chord graphite-epoxy wing. Structural testing showed wing pitch deflections up to +/- 1.5 degree(s) at rates in excess of 70 Hz. Wind tunnel tests were conducted between 1 and 20 ft/s (typical MAV flight speeds) and showed rolling moment coefficients up to +/- 0.026 (equivalent to +/- 8.4 degree(s) of aileron deflection), indicating that the new wing is well suited to flight control of micro-sized aircraft.
This study outlines the design principles, analytical models and testing procedures for a new type of solid state adaptive rotor (SSAR) which was specifically intended for use in mini and micro-unmanned aerial rotorcraft ((mu) UAR). This new SSAR employed a pair of torque-plates which were structurally integrated at the helicopter hub and connected to a two-bladed rotor assembly of 22' (55.9 cm) diameter. These plates were constructed from symmetrically oriented directionally attached piezoelectric actuator elements which were bonded to an aluminum substrate. As the electrical field was changed across the elements, the twist of the plate changed accordingly. Because this new arrangement may be controlled as a function of azimuth, both collective and cyclic commands were available. Unlike earlier designs, this new arrangement used a Hiller servopaddle configuration to achieve flight control. Analytical modeling of the torque- plate performance was accomplished through laminated plate theory and showed good correlation between theory and experiment. Rotor-dynamic models included propeller and aerodynamic moments. Rotor testing showed servopaddle deflection levels in excess of +/- 5.8 degree(s) at rates up to 2.5/rev. Not only is this system effective in achieving flight control, but it is also very simple and lightweight. Indeed, the torque-late, electrical leads and contacts weight 40% less, have a much cleaner hub and replace more than 94 individual components which are found on the conventional flight control system.
The design principles, analytical models, construction methods and test results for a new type of solid state adaptive rotor (SSAR) are presented. A pair of directionally attached piezoelectric (DAP) torque-plates were fabricated and attached to the root of a 23.5' diameter helicopter rotor assembly. The DAP torque-plate tips were joined to a pair of graphite-epoxy servopaddles which were moved in pitch by the action of the torque-plates. The torque-plates were constructed from a single aluminum substrate and PZT-5H DAP elements mounted symmetrically at 45 degrees. Electrical signals were carried to the DAP torque-plates via a shielded brush and rotating contact assembly. A series of non-rotating static tests were conducted on the rotor, demonstrating servopaddle pitch deflections up to plus or minus 5.8 degrees and good correlation with classical laminated plate theory. Non rotating dynamic testing showed a system natural frequency in excess of 2.5/rev and good correlation with inertial models. Because the servopaddles were aeroelastically tailored to balance out propeller moments, deflection degradation with increasing rotor speed was barely noticeable up to plus or minus 1 degree pitch levels. However, as rotor speed increased, total servopaddle deflections in the rotating frame at 1600 rpm (full speed) were degraded, but still operated up to plus or minus 2.7 degrees in pitch. To conclude the study, the rotor was attached to a converted Kyosho Hyperfly electric helicopter. Flight tests demonstrated fundamental controllability. A system-level comparison showed that the SSAR Hyperfly experienced a 40% drop in flight control system weight, an 8% cut in total gross weight, a 26% decrease in parasite drag and a part count reduction from 94 components to 5.
Structural and aerodynamic models for a 10 degree half-angle, articulated, conical, barrel-launched adaptive munition (BLAM) are presented. The forward half of the test BLAM was gimballed and actuated by two orthogonal pairs of piezoceramic tendons. Laminated plate theory models showed that piezoceramic actuators could survive hard-launch conditions if they were properly precompressed by using an elevated temperature cure cycle and a high, favorable mismatch in coefficient of thermal expansion between the substrate and the piezoceramic. Aerodynamic models based on supersonic pressure distributions over cones were used to predict free-flight trim angles and normal force range. A series of bench tests on a 13.4 cm (5.28 in) long BLAM showed plus or minus 0.12 degree steady articulation angles and a first natural frequency of 198 Hz with close agreement between theory and experiment. A series of wind tunnel tests at Mach 3.26 demonstrated a steady deflection increase with field, producing untrimmed normal force coefficient changes of plus or minus 0.0019. Extrapolation of the data to a 105 mm (4.13 in) caliber PGU-28 shaped round fired at 610 m/s (2,000 ft/s) showed that effective range increases by more than a factor of 15 over the conventional PGU-28.
This study outlines active flight control materials, structural arrangements, and several new active flight control methods for rotorcraft, airplanes and missiles. A system-level comparison shows that flight control actuator systems using materials like piezoceramics have approximately double the mass-specific energy and 4 to 6 times the volume specific energy of conventional actuators. New fabrication techniques centered on the principal of directional attachment allow wings and rotor blades to become twist active. Using these new methods, directionally attached piezoelectric (DAP) actuator elements were built into graphite-epoxy sandwich structures. When compared to conventionally attached piezoelectric (CAP) elements, twist deflections (important for flight control) of DAP plates were an order of magnitude greater. By using such twist-active elements in a torque-plate configuration, an active helicopter rotor was built. This Froude-scaled solid state rotor was whirl-stand tested and showed steady blade pitch deflections in excess of plus or minus 8 degrees with good correlation between theory and experiment rates up to 42 Hz (which corresponded to 2.5/rev) and no degradation in deflection as RPM was increased. DAP elements were also used in high aspect ratio subsonic and supersonic wings, demonstrating static twist deflections of plus or minus 2 degrees and plus or minus 6 degrees respectively, with good correlation between experiment and finite element results. The final section compares all-moving active stabilator structural arrangements and pitch deflections, which range up to plus or minus 12 degrees, generating lift coefficient changes in excess of plus or minus 0.8.
A new type of very low stiffness super-active composite material is presented. This laminate uses shape-memory alloy (SMA) filaments which are embedded within a low Durometer silicone matrix. The purpose is to develop an active composite in which the local strains within the SMA actuator material will be approximately 1% while the laminate strains will be at least an order of magnitude larger. This type of laminate will be useful for biomimetic, biomedical, surgical and prosthetic applications in which the very high actuator strength of conventional SMA filaments is too great for biological tissues. A modified form of moment and force-balance analysis is used to model the performance of the super-active shape-memory alloy composite (SASMAC). The analytical models are used to predict the performance of a SASMAC pull-pull actuator which uses 10 mil diameter Tinel alloy K actuators embedded in a 0.10' thick, 25 Durometer silicon matrix. The results of testing demonstrate that the laminate is capable of straining up to 10% with theory and experiment in good agreement. Fatigue testing was conducted on the actuator for 1,000 cycles. Because the local strains within the SMA were kept to less than 1%, the element showed no degradation in performance.
A new type of subsonic missile flight control surface using piezoelectric flexspar actuators is presented. The flexspar design uses an aerodynamic shell which is pivoted at the quarter-chord about a graphite main spar. The shell is pitched up and down by a piezoelectric bender element which is rigidly attached to a base mount and allowed to rotate freely at the tip. The element curvature, shell pitch deflection and torsional stiffness are modeled using laminated plate theory. A one-third scale TOW 2B missile model was used as a demonstration platform. A static wing of the missile was replaced with an active flexspar wing. The 1' X 2.7' active flight control surface was powered by a bi-morph bender with 5-mil PZT-5H sheets. Bench and wind tunnel testing showed good correlation between theory and experiment and static pitch deflections in excess of +/- 14 degree(s). A natural frequency of 78.5 rad/s with a break frequency of 157 rad/s was measured. Wind tunnel tests revealed no flutter or divergence tendencies. Maximum changes in lift coefficient were measured at (Delta) CL equals +/- .73 which indicates that terminal and initial missile load factors may be increased by approximately 3.1 and 12.6 g's respectively, leading to a greatly reduced turn radius of only 2,400 ft.
The structural and aerodynamic characteristics of a new class of active flight control surface are presented. This new type of surface uses a symmetric, subsonic aerodynamic shell which is supported at the quarter-chord by a main spar and actively pitched by an adaptive torque- plate. The structural mechanics of the torque-plate and several actuator elements are detailed, including newly invented interdigitated electrode and constrained directionally attached piezoelectric elements. Laminated plate models demonstrate that both generate similar deflections with comparable torsional stiffnesses. An experimental torque-plate specimen constructed from PSI-5A-S2 piezoceramic shows high torsional deflections and stiffness as well as excellent correlation with theory. The constrained torque-plate was integrated into a 12.5 cm plan X 5 cm chord adaptive missile fan which was designed for Mach 0.6 flight under standard conditions. The specimen showed static pitch deflections up to +/- 8.1 degree(s) and dynamic deflections of +/- 19 degree(s) at resonance. The active surface was also wind tunnel tested up to 40 m/s and demonstrated invariant pitch deflections as a function of airspeed, a steady break frequency of 50 Hz, no flutter, buffet or divergence tendencies and steady lift coefficient changes up to +/- 0.51.
The design of a new aeroservoelastic wing configuration is detailed. The design uses a torque plate mounted within an aerodynamic shell. As the plate is actively twisted, the shell, which is connected to the plate at the tip, undergoes a pitch deflection. An active torque plate for a low aspect ratio subsonic wing section is analyzed and designed with laminated plate theory. Several actuator elements were considered for the plate, including: conventionally attached lead zirconate titanate (PZT), polyvinylidene fluoride (PVDF), piezoelectric fiber composites and directionally attached piezoelectric (DAP) elements. The analytical studies demonstrate that the highest twist deflections are obtained by DAP elements bonded to a beryllium substrate which produces twist deflections and restrained moments that are 67% greater than the next closest actuator material. An aeroservoelastic wing using the DAP torque plate was constructed to demonstrate the concept. The wing used an active plate made from DAP elements bonded to an AISI 1010 steel substrate. The torque plate was mounted within a graphite-epoxy wing with a modified NACA 0012 profile, measuring 2.6' in chord with a 3.66' span. The wing was tested to 160 ft/s and exhibited a stable increase in pitch deflection (in addition to the +/- 3.5 degree(s) static deflection) with theory and experiment in close agreement.
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