Rotorcraft Blade Angle Calibration Methods

The most vital system of a rotorcraft is the rotor system due to its effects on the overall flight quality of the vehicle. Therefore, it is of importance to be able to accurately determine blade position during flight so that fine adjustments can be made to ensure a safe and efficient flight. In this study, a current calibration method focusing on the pitch, flap, and lead-lag blade angles is analyzed and found to have larger than acceptable error associated with the sensor calibrations. A literature review is conducted which reveals four novel methods that can potentially increase the accuracy of the sensor calibrations. An uncertainty analysis is conducted aiding in the decision of which of the four methods would best improve the calibration accuracy. The results conclude that a simpler method can be applied and calibration times can greatly be reduced while increasing the accuracy of the calibration. Finally, a new calibration method is proposed utilizing the newly chosen sensor that can be later implemented into the system

An Investigation of the Rotor Tip Path Height of the MH-60S Helicopter in View of Forklift Clearance in Support of the United States Navy Medium Lift Shipboard Logistics Mission

The purpose of this paper is to summarize Department of the Navy tests performed to measure rotor tip path height of the MH-60S helicopter and present an analysis of collected data to determine if safe cargo loading operations on the MH-60S can be conducted with a forklift while the rotor is engaged. Testing was conducted to measure the dynamic height of the rotor tip path plane during incremental cyclic displacements, rotor response to external disturbances, and pilot tendencies when centering the cyclic control stick. Additional information was gathered on representative forklifts in use on U.S. Navy ships, and shipboard operating procedures for cargo movement. A comparison between the forklift and rotor heights was conducted to evaluate the clearance available for forklifts transiting the rotor arc. While it cannot be concluded that cargo loading using a forklift with the rotor engaged can be conducted without incident, substantial data were gathered that indicated that current safety precautions coupled with the clearance from the engaged rotor would allow for safe conduct of the evolution. Specifically, if operations are conducted with low profile forklifts, which have an obstruction height shorter than the average male, rotor clearance is considered sufficient to preclude catastrophic interaction between the rotor and the equipment. Additional research, safety review, and equipment and publication changes are recommended to further increase the safety of conducting these operations

Swashplateless Helicopter Experimental Investigation: Primary Control with Trailing Edge Flaps Actuated with Piezobenders

Helicopter rotor primary control is conventionally carried out using a swashplate with pitch links. Eliminating the swashplate promises to reduce the helicopter's parasitic power in high speed forward flight, as well as may lead to a hydraulic-less vehicle. A Mach-scale swashplateless rotor is designed with integrated piezobender-actuated trailing edge flaps and systematically tested on the benchtop, in the vacuum chamber and on the hoverstand. The blade is nominally based on the UH-60 rotor with a hover tip Mach number of 0.64. The blade diameter is 66 inches requiring 2400 RPM for Mach scale simulation. The rotor hub is modified to reduce the blade fundamental torsional frequency to less than 2.0/rev by replacing the rigid pitch links with linear springs, which results in an increase of the blade pitching response to the trailing edge flaps. Piezoelectric multilayer benders provide the necessary bandwidth, stroke and stiffness to drive the flaps for primary control while fitting inside the blade profile and withstanding the high centrifugal forces. This work focuses on several key issues. A piezobender designed from a soft piezoelectric material, PZT-5K4, is constructed. The new material is used to construct multi-layer benders with increased stroke for the same stiffness relative to hard materials such as PZT-5H2. Each layer has a thickness of 10 mils. The soft material with gold electrodes requires a different bonding method than hard material with nickel electrodes. With this new bonding method, the measured stiffness matches precisely the predicted stiffness for a 12 layer bender with 1.26 inch length and 1.0 inch width with a stiffness of 1.04 lb/mil. The final in-blade bender has a length of 1.38 inches and 1.0 inch width with a stiffness of 0.325 lb/mil and stroke of 20.2 mils for an energy output of 66.3 lb-mil. The behavior of piezobenders under very high electric fields is investigated. High field means +18.9 kV/cm (limited by arcing in air) and -3.54kV/cm (limited by depoling). An undocumented phenomenon is found called bender relaxation where the benders lose over half of their initial DC stroke over time. While the bender stiffness is shown not to change with electric field, the DC stroke is significantly less than AC stroke. A two-bladed Mach-scale rotor is constructed with each blade containing 2 flaps each actuated by a single piezobender. Each flap is 26.5% chord and 14% span for a total of 28% span centered at 75% of the blade radius. Flap motion of greater than 10 degrees half peak-peak is obtained for all 4 flaps at 900 RPM on the hoverstand. So, the flaps show promise for the Mach-scale rotor speed of 2400 RPM. A PID loop is implemented for closed loop control of flap amplitude and mean position. On the hoverstand at 900 RPM, the swashplateless concept is demonstrated. The linear springs used to lower the torsional frequency are shown to have minimum friction during rotation. 1/rev blade pitching of ±1 degree is achieved at a torsional frequency of 1.5/rev for each blade. At resonance, the blade pitching for each blade is greater than ±4 degrees. Primary control is demonstrated by measuring hub forces and moments. At resonance state, the flaps in conjunction with the blade pitching provide ±15 lbs of normal force at a mean lift of 15 lbs yielding ±100% lift authority. Significant hub forces and moments are produced as well. For a production swashplateless helicopter, it may be prudent to eliminate the pitch links by reducing the blade structural stiffness. A novel wire sensor system network is proposed in order to measure blade elastic flap bending, lead-lag bending and torsion. The theory for measuring blade twist is rigorously derived. A blade is constructed with the wire sensor network and validated on the benchtop for blade elastic bending and twist. This work is a step forward in achieving a swashplateless rotor system. Not only would this reduce drag in high speed forward flight, but it would lead to a hydraulic-less rotorcraft. This would be a major step in vertical flight aviation

Development and Whirl Flutter Testing of Swept-Tip Tiltrotor Blades

This thesis describes the development and whirl flutter testing of swept-tip tiltrotor blades. The blades are tested on the Maryland Tiltrotor Rig (MTR). The MTR is also developed as part of this thesis. MTR is a new test facility developed to support the research and development of next-generation high-speed tiltrotors. It is a parametric test bed for developing a fundamental understanding of high speed tiltrotor flight and acquisition of test data to validate advanced simulation tools. The baseline MTR is a Froude-scaled, 4.75 feet diameter, 3-bladed, semi-span, floor-mounted, optionally-powered, gimballed flutter rig. In this work, the rotor blades are designed with the objective of gaining a fundamental understanding of the impact of a swept-tip on tiltrotor whirl flutter. Two sets of blades are fabricated for wind tunnel testing -- straight and swept. The blades have a VR-7 profile, chord of 3.15 inches, and linear twist of -37° per span. The blades have a uniform cross section with mass and stiffness properties loosely based on a 1/5.26 Froude scale XV-15 rotor properties. The swept-tip blades are identical to the straight blades up to 80% radius, where a 20° sweep back angle is introduced to assess the impact on whirl flutter. The cross-section was designed using in-house 2-D section analysis tools. The blade inertial and structural properties were carefully measured. A novel method was developed to attain more reliable measurements of the blade cross-sectional stiffness using accelerometers as tilt sensors. Full 3-D models of the blades were also developed concurrently. These models were built in CATIA, meshed in Cubit, and analyzed with X3D. These models were validated with test data of the measured blade properties. Experiments in a vacuum chamber were carried out to measure frequencies and strains. The 3-D model was validated with this data. Tests and predictions proved the blades have sufficient structural integrity and stress margins to allow for wind tunnel testing. The first whirl flutter test of the MTR was completed in the Naval Surface Warfare Center Carderock Division (NSWCCD) 8- by 10-ft large subsonic wind tunnel. Testing was performed in four configurations for both blade geometries. The configurations progressed step-by-step from the baseline configuration of a gimballed rotor, freewheeling flight, with wing fairings installed; the second configuration removed the wing fairings; the third then locked the rotor gimbal; the fourth configuration then operated in powered flight. The frequency and damping of the wing beam and chord bending were collected at a rotor speed of 1050 RPM and wind speeds up to 100 knots. The 100 knots speed was a limitation posed by NSWCCD. In freewheel flight, the swept-tip blades increase the damping and stability of the wing chord bending mode, even at lower flight speeds. However, in powered flight, the opposite effect is observed. The parametric testing results in a comprehensive data set for evaluating the effect of swept-tip blades on whirl flutter. Richer data is expected at higher speeds where the aerodynamic and inertial couplings of the swept-tip blades are expected to be more pronounced. Nevertheless, it is hoped that the results presented in this work will inspire further investigation using advanced computational tools to develop a more complete understanding of tiltrotor whirl flutter and ultimately eliminate it

Comprehensive aeromechanical measurements of a model-scale, coaxial, counter-rotating rotor system

With the renewed interest in rigid, counter-rotating coaxial rotor designs, and the increased fidelity of fully coupled CFD/CSD simulations, there exists a lack of comprehensive experimental data for a rotor system with which to validate analyses. The goal of this dissertation is to generate a new set of measurements on a model-scale rigid coaxial rotor systems in hover and high speed forward flight. A counter-rotating transmission was built, incorporating 6-component upper and lower rotor load cells for individual hub load measurements. Upper and lower rotor control systems, as well as complementary instrumentation including pushrod load cells, root pitch measurement and blade tip clearance sensors were developed. Two sets of rotor blades were fabricated and characterized using stereoscopic digital image correlation in combination with static and dynamic loads. A novel rotating-frame operational modal analysis successfully identified the first blade flap frequency and aerodynamic damping. Hover testing focused on quantifying the effects of upper and lower coaxial rotor interference when compared to isolated rotors. Statistical analysis of the measured data revealed clear trends with a known confidence level. Due to mutual interference, the upper and lower rotors of the coaxial configuration consumed 18% and 49% more induced power than that of an isolated two-bladed rotor. The coaxial counter-rotating configuration was found to consume 6% less induced power than an isolated, four-bladed single rotor of equal solidity. While torque balanced, the upper rotor was found to produce 54% of the total system thrust regardless of blade loading. Significant four-per-revolution vibratory thrust was observed in the lower rotor, with primary and secondary peaks corresponding to bound vortex and blade thickness interactions respectively. Wind tunnel testing examined the effects of lift offset and rotor phasing at high forward flight speeds. Rotor effective lift-to-drag ratio was found to increase with increasing advance ratio and lift offset, resulting in a 50% peak efficiency gain. The lower coaxial rotor was found to operate at higher lift-to-drag ratio than the upper rotor, due to the reversal of differential upper and lower rotor thrust compared to hover. Lift offset resulted in a decrease in blade tip clearance with a corresponding rise in rotor side force. Vibratory loads increased with advance ratio, with the largest occurring at two and four-per-revolution harmonics. Lift offset decreased vibratory forces while increasing vibratory in-plane moments. The coaxial system experienced reduced vibratory in-plane forces and torque compared to the isolated rotors due to cancellation between upper and lower rotor loads. Adjusting the inter-rotor index angle modified vibratory forces and moments transmitted to the fixed frame, increasing some components while decreasing others.Aerospace Engineerin

Evaluation of at-sea flight testing of the MV-22 Osprey for operational employment

The MV-22 Osprey tiltrotor aircraft is a radically new air vehicle designed to replace aging helicopters and support the US Marine Corp\u27s future concept of operational maneuver from the sea. Unfortunately the aircraft has been plagued with political and programmatic delays throughout its 19-year history that prevented early and comprehensive at-sea testing. With an operational evaluation in October 1999, shortened at-sea test period was required late in the aircraft development in January 1999.This thesis analyzes the compressed developmental test process used to prepare this novelair vehicle for sea service in a short time period.The dynamic interface testing of Naval aircraft and ships is not new, although the advent of tiltrotors incorporating digital fly-by-wire technology has challenged traditional developmental procedures. The MV-22 required extensive test planning, flying qualities evaluations and engineering tests to define safe operational limits in the shipboard environment. An analysis of a lateral control instability problem encountered during the testing and the subsequent test process innovations for this unique aircraft substantiate the need to conduct comprehensive and extensive developmental testing.It is the author\u27s opinion that at-sea testing is risky and the final exam for a Naval Aircraft. The risks of a shortened test process were that deficiencies would be uncovered and that uncharted capabilities would not be exploited for operational employment. The Documented successes and failures of the MV-22 at-sea test process yield lessons that should be put into practice by future amphibious Vertical Take Off and Landing (VTOL)aircraft such as the Joint Strike Fighter (JSF) and other follow-on VSTOL aircraft

Development of helicopter attitude axes controlled hover flight without pilot assistance and vehicle crashes.

In this work, we show how to computerize a helicopter to fly attitude axes controlled hover flight without the assistance of a pilot and without ever crashing. We start by developing a helicopter research test bed system including all hardware, software, and means for testing and training the helicopter to fly by computer. We select a Remote Controlled helicopter with a 5 ft. diameter rotor and 2.2 hp engine. We equip the helicopter with a payload of sensors, computers, navigation and telemetry equipment, and batteries. We develop a differential GPS system with cm accuracy and a ground computerized navigation system for six degrees of freedom (6-DoF) free flight while tracking navigation commands. We design feedback control loops with yet-to-be-determined gains for the five control "knobs" available to a flying radio-controlled (RC) miniature helicopter: engine throttle, main rotor collective pitch, longitudinal cyclic pitch, lateral cyclic pitch, and tail rotor collective pitch.We develop helicopter flight equations using fundamental dynamics, helicopter momentum theory and blade element theory. The helicopter flight equations include helicopter rotor equations of motions, helicopter rotor forces and moments, helicopter trim equations, helicopter stability derivatives, and a coupled fuselage-rotor helicopter 6-DoF model. The helicopter simulation also includes helicopter engine control equations, a helicopter aerodynamic model, and finally helicopter stability and control equations. The derivation of a set of non-linear equations of motion for the main rotor is a contribution of this thesis work.After discussing the integration of hardware and software elements of our helicopter research test bed system, we perform a number of experiments and tests using the two specially built test stands. Feedback gains are derived for controlling the following: (1) engine throttle to maintain prescribed main rotor angular speed, (2) main rotor collective pitch to maintain constant elevation, (3) longitudinal cyclic pitch to maintain prescribed pitch angle, (4) lateral cyclic pitch to maintain prescribed roll angle, and (5) yaw axis to maintain prescribed compass direction. (Abstract shortened by UMI.)We design and build two special test stands for training and testing the helicopter to fly attitude axes controlled hover flight, starting with one axis at a time and progressing to multiple axes. The first test stand is built for teaching and testing controlled flight of elevation and yaw (i.e., directional control). The second test stand is built for teaching and testing any one or combination of the following attitude axes controlled flight: (1) pitch, (2) roll and (3) yaw. The subsequent development of a novel method to decouple, stabilize and teach the helicopter hover flight is a primary contribution of this thesis.The novel method included the development of a non-linear modeling technique for linearizing the RPM state equation dynamics so that a simple but accurate transfer function is derivable between the "available torque of the engine" and RPM. Specifically, the main rotor and tail rotor torques are modeled accurately with a bias term plus a nonlinear term involving the product of RPM squared times the main rotor blade pitch angle raised to the three-halves power. Application of this non-linear modeling technique resulted in a simple, representative and accurate transfer function model of the open-loop plant for the entire helicopter system so that all the feedback control laws for autonomous flight purposes could be derived easily using classical control theory. This is one of the contributions of this dissertation work

Autocalibrating vision guided navigation of unmanned air vehicles via tactical monocular cameras in GPS denied environments

This thesis presents a novel robotic navigation strategy by using a conventional tactical monocular camera, proving the feasibility of using a monocular camera as the sole proximity sensing, object avoidance, mapping, and path-planning mechanism to fly and navigate small to medium scale unmanned rotary-wing aircraft in an autonomous manner. The range measurement strategy is scalable, self-calibrating, indoor-outdoor capable, and has been biologically inspired by the key adaptive mechanisms for depth perception and pattern recognition found in humans and intelligent animals (particularly bats), designed to assume operations in previously unknown, GPS-denied environments. It proposes novel electronics, aircraft, aircraft systems, systems, and procedures and algorithms that come together to form airborne systems which measure absolute ranges from a monocular camera via passive photometry, mimicking that of a human-pilot like judgement. The research is intended to bridge the gap between practical GPS coverage and precision localization and mapping problem in a small aircraft. In the context of this study, several robotic platforms, airborne and ground alike, have been developed, some of which have been integrated in real-life field trials, for experimental validation. Albeit the emphasis on miniature robotic aircraft this research has been tested and found compatible with tactical vests and helmets, and it can be used to augment the reliability of many other types of proximity sensors

Benefits assessment of active control technology and related cockpit technology for rotorcraft

Two main-rotor active control concepts, one incorporating multicyclic actuators located just below the swashplate, and the other providing for the actuators and power supplies to be located in the rotating frame are considered. Each design concept is integrated with cockpit controllers and displays appropriate to the actuation concept in each case. The benefits of applying the defined ACT/RCT concepts to rotorcraft are quantified by comparison to the baseline model 412 helicopter. These benefits include, in the case of one active control concept; (1) up to 91% reduction in 4/rev hub shears; (2) a flight safety failure rate of 1.96 x 10 to the 8th power failures per flight-hour; (3) rotating controls/rotor hub drag reduction of 40%; (4) a 9% reduction in control system weight; and (5) vibratory deicing. The related cockpit concept reduces pilot workload for critical mission segments as much as 178% visual and 25% manual

A Continuous-Time Nonlinear Observer for Estimating Structure from Motion from Omnidirectional Optic Flow

Various insect species utilize certain types of self-motion to perceive structure in their local environment, a process known as active vision. This dissertation presents the development of a continuous-time formulated observer for estimating structure from motion that emulates the biological phenomenon of active vision. In an attempt to emulate the wide-field of view of compound eyes and neurophysiology of insects, the observer utilizes an omni-directional optic flow field. Exponential stability of the observer is assured provided the persistency of excitation condition is met. Persistency of excitation is assured by altering the direction of motion sufficiently quickly. An equal convergence rate on the entire viewable area can be achieved by executing certain prototypical maneuvers. Practical implementation of the observer is accomplished both in simulation and via an actual flying quadrotor testbed vehicle. Furthermore, this dissertation presents the vehicular implementation of a complimentary navigation methodology known as wide-field integration of the optic flow field. The implementation of the developed insect-inspired navigation methodologies on physical testbed vehicles utilized in this research required the development of many subsystems that comprise a control and navigation suite, including avionics development and state sensing, model development via system identification, feedback controller design, and state estimation strategies. These requisite subsystems and their development are discussed