Browse Topic: Engines
Drain and Fill plugs used on engines, transmissions, transfer cases and front and rear drive axles for class 5 – 8 vehicles.
Helicopter tail shake constitutes a significant limitation to both passenger comfort and aircraft stability. Under powered descent conditions, elevated Angle of Attack (AoA) cause flow separation around the rotor hub and engine cowling, leading to the development of an unsteady wake dominated by large-scale turbulent structures. To support the helicopter tail shake phenomenon investigation, a dedicated Particle Image Velocimetry (PIV) experimental setup was designed in this work, together with four aerodynamic devices aimed at mitigating tail shake. These components were then tested through a wind tunnel campaign with the PIV setup. The proposed aerodynamic components were conceived to either deflect the hub wake away from the tail empennages or to decrease the Turbulent Kinetic Energy (TKE) within the wake. To achieve these objectives, a dorsal fin, a horse-collar, and two spoiler configurations inspired by automotive applications were designed and experimentally evaluated. The devices were tested both as standalone solutions and in combined arrangements on a scaled helicopter wind tunnel model featuring a rotating hub and blade shanks. The vertical velocity component, was used as an indicator of wake deflection, and the Turbulent Kinetic Energy was used as an indicator of wake turbulence. The Horse Collar and the Large Spoiler showed a reduction in both indicators suggesting possible tail shake mitigating capabilities, and additional improvements were achieved when the two devices were deployed in combination.
Emerging technologies in the field of electrified propulsion systems offer a promising solution to reduce the dependence on fossil fuels and improve efficiency. However, the design of high-power density electric machines introduces new challenges, including limited passive cooling potential and the issue of the weight of electric motors. To address these challenges, this paper considers analysis and design methods for high torque-to-weight ratio axial flux motors. A magnetic equivalent circuit model coupled with a lumped parameter thermal network is developed for design space exploration and optimization. This inexpensive analytical model predicts the performance of a single-stator dual-rotor axial flux motor based on geometry, loading condition, and slot and pole pair combination. To enable comparisons against real-world data, the optimization study was demonstrated using the hover mission requirements from the Research Aircraft for eVTOL Enabling techNologies (RAVEN) vehicle to minimize the mass of the motor. In tandem with the analytical model, a higher-fidelity finite element model was also developed, and good agreement between predicted power and efficiency was demonstrated across a range of axial flux motor designs. The lightest weight design that satisfied the hover mission requirements was the 12 pole pair 27 slot (12PP 27S) configuration with a fixed weight of 9.28 kg. The analytic model undersized the output power of the electric motor by approximately 9% across a range of slot and pole pair combinations.
Urban Air Mobility (UAM) concepts require multidisciplinary analyses across multiple modes of operation and often involve discrete architectural differences such as propulsion type, rotor configuration, and mission context. Existing optimization and workflow frameworks support continuous design variables but provide limited mechanisms for handling discrete variants, multi-modal vehicle definitions, and vehicle management for UAM vehicles. This paper presents uam4x, an open-source Python framework that addresses these challenges through a structured problem definition representation, a plugin-based execution engine, integrated version control, and a function-based branching script mechanism for constructing analysis scenarios. The framework provides integration of existing tools including Open Vehicle Sketch Pad (OpenVSP), NASA Design and Analysis of Rotorcraft (NDARC), M4 Structures Studio (M4SS), and Intelligent Cross Section Generator (IXGEN) through unified plugin interfaces. Parameter sweeps, nested analyses, and optimization via OpenMDAO are supported within the same architecture. This paper also presents demonstrations that were created to illustrate the various capabilities and integration efforts of the framework.
Vertical Take-Off and Landing (VTOL) aircraft introduce complex monitoring challenges due to distributed propulsion, lightweight structures, and variable operating conditions. This paper presents advanced Frequency and Orders domain techniques that repurpose existing flight control, propulsion, and structural sensor data to enhance observability without additional instrumentation. By transforming vibration, acoustic, and electrical signals into frequency and order domains, the approach enables detection of harmonics, resonance, and fault signatures tied to rotor dynamics, supporting adaptive control and predictive maintenance. Beyond rotor systems, these techniques are equally effective for monitoring electric motor health, gearbox wear, bearing degradation, and structural coupling effects in composite airframes. They also provide insight into power electronics and thermal management systems by identifying spectral anomalies linked to electrical imbalance or cooling inefficiencies. Aggregated fleet data strengthens prognostic capabilities, enabling early detection of systemic issues and trend analysis. Applications include mitigating ground resonance and modal instabilities, as well as improving reliability of propulsion and structural subsystems. Integration into avionics emphasizes computational efficiency, scalability, and compliance with standards such as DO-160 [1], DO-178 [2], ARP4761 [3] and ARP4764 [4]. Simulation and bench testing confirm feasibility, demonstrating potential to enhance safety, reliability, and lifecycle cost for next-generation urban air mobility platforms.
Hydrogen-electric vertical takeoff and landing (H2eVTOL) (or fuel cell-electric VTOL) aircraft technologies are poised to emerge in the next coming decades and start operating from existing heliports and new vertiports. This paper assesses how key H2eVTOL design features interact with the ground infrastructure and how facility designers can address H2eVTOL specific facility requirements–especially the supply of hydrogen to the aircraft. Vertiport design should maximize compatibility are important to facilitate the accommodation of hydrogen technologies, minimize the need for extensive capital investments, and promote safety and operational efficiency. Considerations should be given to factors such as general aircraft configuration, electric and hybrid propulsion systems, and refueling infrastructure. The definition of notional aircraft concepts representing the evolution of critical VTOL aircraft over the next coming decades can help aviation facility planners and designers understand the type of vehicles they need to account for and also evaluate the future hydrogen demand. The lack of aviation-specific standards, especially when it comes to fire prevention, might adversely impact vertiports. Strategies are proposed for mitigating the effects of hydrogen operations at space-constrained facilities.
This paper, explores the design and sizing of a planetary gear-based electronic continuously variable transmission (ECVT) for implementation of a parallel gas-electric hybrid helicopter propulsion system. The ECVT consists of a differential planetary gear transmission (PGT) and an electric motor/generator (MG) unit. The ECVT enables power-flow between engine, motor and helicopter main rotor. The parallel arrangement enables the main rotor speed to varied continuously based on the MG speed while the engine speed can remain constant. The performance benefits enabled by the main rotor speed variation capability are offset by the added weight penalties introduced by the ECVT system. By considering factors such a as gear tooth bending and contact stress, bearing loads, required motor torque, planetary gear kinematics and pitch-line velocity constraints, this paper conducts a minimum mass design study for several PGT / ECVT arrangements. Here, three different single stage PGT/ECVT arrangements are compared along with an improved two stage ECVT. The three single stage ECVT configurations can be summarized as; I) Sun-Engine / Carrier-Motor / Ring-Out, II) Sun-Engine / Ring-Motor / Carrier-Out, and III) Carrier-Engine / Sun-Motor / Ring-Out. Of these three types, it was found that type III was significantly lighter in weight compared with types I and II since type III would have the highest relative motor speed. When sized for a 3000 Hp engine-side power input at 6000 rpm, the minimum mass design for type III was on the order of 100 lbs compared to 400 lbs and 700 lbs respectively for types I and II. Despite the seemingly obvious advantage of design type III, it's drawback is that it is effectively a speed increasing stage with respect to the engine. To address this, a two-stage ECVT with compound planetary arrangement of Type III and II was designed which achieved an overall minimum weight of 219 lbs at the 3000 Hp level while providing 1:0.351 gear reduction form engine to output. The analysis tools developed and sizing results flowing from this study will provide a baseline for evaluating performance benefits and weight penalties introduced by parallel hybrid drive-systems for rotorcraft applications.
The purpose of the NATO Next Generation Rotorcraft Capability (NGRC) Support Partnership funded Novel Powerplant concept study was to identify, analyze, and compare novel powerplant concepts that could fulfill the NGRC need in a solution-agnostic approach. The outcome of the study provided NSPA and the NGRC participating nations with increased knowledge and understanding of the powerplant domain to inform assessment of future NGRC platforms. This study modeled four aircraft configurations to derive propulsion sizing requirements and compared propulsion configurations for each. The propulsion system configurations considered included three levels of conventional gas turbine technology (In-service GT, 2025 GT, and 2035 GT), hybrid electric (battery), hybrid hydrogen fuel cell, and hydrogen combustion. The results of the study considered both quantitative and qualitative evaluations. The quantitative analysis determined aircraft and propulsion system sizing to align with the expected NGRC need at different max cruise speeds, and compared performance along with recurring, non-recurring, and operational costs for each. The qualitative analysis investigated the impacts of wider technical, capability, and supportability considerations for each propulsion concept. Assuming a notional NGRC entry into service target date of 2035, this study shows that, when considering both quantitative analysis and a qualitative review, a modern gas turbine is the most reasonable solution space to deliver the performance and multi-mission capability expected.
Electrification could improve full-size rotorcraft performance by reducing peak turbine power demand, reducing transmission system weight and complexity, and reducing operating costs. Integrating electric machines with mechanical powertrains requires careful consideration of the system-level weight and efficiency impacts. This paper presents an optimization framework for evaluating parallel hybrid powertrain configurations using Geometric Programming (GP). Both retrofit and clean-sheet vehicle designs are considered. The results show that high-speed electric motors integrated into a parallel hybrid configuration using batteries can reduce the sized gas turbine power, enabling more efficient engine operation at lower power levels. For retrofit designs, with a fixed vehicle gross weight, adding batteries and motors reduces usable fuel, decreasing mission capability. Clean-sheet designs offer additional flexibility to re-size the vehicle and rotor, resulting in energy savings for an equivalent design mission.
This study presents the design, modeling, and simulation of an Adaptive Speed Gearbox (ASG) with integrated electric variator for the UH-60A Black Hawk helicopter. The proposed drivetrain architecture enables main rotor speed variation independently of turbine speed, addressing operational demands for enhanced efficiency, noise reduction, and performance flexibility. A comprehensive aero-thermal model of the turboshaft engine, a dynamic drivetrain model, and a variable-speed control strategy were developed and validated. The control approach employs a two-degree-of-freedom structure combining nullspace-based feedforward torque allocation and modal-weighted LQR feedback for vibration suppression. A similarity theory-based scaling method was employed to design a demonstrator gearbox, facilitating experimental validation under representative conditions. The results demonstrate the feasibility of the ASG concept and establish a foundation for future experimental investigations and subsequent technology maturation towards higher TRLs.
Electric vertical take-off and landing vehicles are proposed as a viable solution for urban air mobility due to their potential for reducing carbon emissions, noise, and operational costs. However, the shift towards electrified aircraft introduces new thermal management issues due to the excess heat generated by electric motors and power electronics. This heat is challenging to dissipate during the mission, resulting in transient motor temperatures, especially during high-power mission segments. In addition, electrified aircraft also encounter design challenges associated with the fixed weight of electric motors and batteries. To address these challenges, this work presents a multifidelity framework for performing shape optimization of an electric motor subject to performance, geometric, and thermal transient constraints. A preliminary sizing of the electric motor is performed using a low fidelity Fourier series model. Next, the sizing is refined by utilizing a coupled electromagnetic-thermal finite element model of the the motor physics. To demonstrate the framework, a numerical optimization is performed for a hover hold mission with a relevant scale eVTOL vehicle, reducing the structural mass of the motor by 2.7kg while ensuring power and thermal constraints are satisfied.
A regulated hybrid-electric power sharing architecture was developed and tested for VTOL applications. In this architecture, there are two power supply branches and one load. The first branch draws power from an engine-generator, and it has additional components of an AC-DC rectifier, a DC-DC buck converter, and a power diode. The second branch draws power from a battery, and it has additional components of a solid-state relay, a DC-DC boost converter, and a power diode. Any specified ratio of battery-to-engine power can be achieved with this architecture. Testing on the full range of power share ratios was conducted at a low load power of 300W. The key conclusions are that: (1) regulated power sharing is feasible between an AC supply and a DC battery, including the extremes of all engine and no battery to all battery and no engine, (2) a specified power share ratio can be achieved both in steady-state and transient conditions, and (3) there is a delay in achieving a specified power share ratio, caused not by the power plant, but rather by the change in RPM of the rotor.
A propeller driven rotor uses small electric motors and propellers attached to the rotor blade to spin the main rotor. Recent propeller driven rotor hover test campaigns suffered propeller failures at relatively low main rotor rotational speeds. The dynamics of spinning a fast propeller at the end of a spinning main rotor blade were the suspected cause of the propeller blade failure. An experiment using the 10 ft diameter vacuum chamber was designed to isolate and measure the propeller flapping motion of an articulated propeller blade from inertial loads. A Coriolis coupling exists between the propeller and the main rotor, resulting in large 20° sinusoidal propeller flapping motions. The vacuum chamber experiment also demonstrated that for the propeller/rotor speed ranges tested, increasing the propeller or the main rotor speed resulted in larger propeller flapping motion. An analytical model was developed to study the coupled propeller flapping motion due to the main rotor rotation. The predicted solutions for the analytical model capture the vacuum chamber experimental results well. Key insights were obtained by analytically solving the propeller flapping equation of motion assuming periodic coefficients (β0, β1c, β1s). This analytical solution method showed that the β1s term was associated with the Coriolis coupling between the main rotor and the propeller, whereas the β0, β1c terms result from the location of the propeller forward of the rotor 1/4 chord location.
This paper presents handling qualities (HQs) research findings for electrical Vertical Take-off and Landing vehicles. Testing in the Vertical Motion Simulator (VMS) investigated handling qualities of vehicle configurations having a degraded powertrain. Powertrain components, including batteries and electric motors, can degrade as the vehicle is flown. This paper investigates the impact of low battery charge and high motor temperature degradations on the pilot's ability to execute precise maneuvers. Pilot comments and ratings that were collected from four rotorcraft test pilots in VMS testing are used to quantify the effects that powertrain degradations had on the HQs of the vehicle.
Developed in the frame of the European Clean Sky 2 program, the RACER High Speed Helicopter Demonstrator of Airbus performed its maiden flight on April 25th, 2024. In the continuity of the previous high-speed demonstrator X3 (1st flight in 2010) the RACER is a 7/8t (15000 / 18000 lb) class compound helicopter powered by two SHE Aneto-1X engines, including a wing and two propellers. The tail rotor is removed as the two propellers control the yaw axis by differential thrust. At flight 07, with its initial default settings, it reached a true airspeed of 227 kts in level flight, exceeding its objective of 220 kts.
This paper presents experimental research aimed at developing novel low lubrication methods for rotorcraft and jet engines, focusing on sustaining minimal lubrication to prevent catastrophic bearing failure during loss of lubrication (LoL) events or to increase fuel consumption performance on once-through, fuel-oil bearing lubrication engines. Utilizing two high-speed bearing test rigs simulating low and high thrust class engine conditions, the study establishes lower bounds for oil flow rates necessary to maintain thermal stability and prevent thermal runaway in hybrid ball bearings. These findings inform the design of the Zulu Pod (ZPod), a passively driven, self-contained oil delivery system that uses engine compressor bleed air to precisely meter lubricant flow. Engine test stand results demonstrate that replacing traditional fuel-oil lubrication with the ZPod system reduces thrust specific fuel consumption (TSFC) by an average of 7%, with up to 11% savings, without compromising engine thrust or bearing health. The ZPod offers a simplified, efficient alternative to fuel-lubricated systems by eliminating fuel diversion for lubrication, enhancing fuel efficiency, and maintaining bearing performance in attritable or single-use engines. Additionally, the study highlights the potential of minimal lubrication supplied by the ZPod to extend operational life during LoL scenarios, enabling safer aircraft recovery. Future work will focus on extending testing to higher thrust classes and optimizing ZPod designs for broader applications.
The oil cooling fan of a Main Gearbox (MGB) is a mechanically-driven component whose purpose is to force an air flow through an air cooled oil cooler; its performance is crucial in ensuring that the MGB oil temperature does not exceed a predefined threshold, set to alert the crew in case of an abnormal situation. The design and the certification of a cooling fan is a process involving several steps and multiple disciplines; mechanical design, aerodynamic analysis, dedicated tests carried out both on rigs and at aircraft level need to be exploited as complementary tools to assess the correct aero-mechanical behavior of the system. The aerodynamic assessment is associated to performance, measured in terms of MGB oil temperature: considering a comparison between two cooling fans, one outperforms the other if the resultant MGB oil temperature is lower, keeping the same boundary conditions (engine torque, wind speed, ambient temperature, etc.). The correct mechanical behavior is instead associated to multiple requirements that need to be satisfied: target fatigue life, high and low temperature limits, blade loss containment, maximum speed, etc. The process connecting the very first blank page to the certification of the part is usually linear and well defined, and the experience gathered by Leonardo Helicopters over the years has allowed to follow this path easily in most cases. This paper discusses an intriguing situation: the goal of the activity herein discussed was to replace an obsolete oil cooling fan with a new one. The latter was expected to outperform the first both from the aerodynamic point of view (i.e. lower MGB oil temperature) and from the mechanical point of view (i.e. longer life). The preliminary CFD analyses and the rig tests proved that the new fan was able to outperform the legacy one in regards of the aerodynamic requirement; nevertheless, during the last steps of verification carried out at aircraft level, the installation of the new fan resulted in a higher MGB oil temperature. This triggered dedicated troubleshooting: a sequence of investigative tests performed by means of dedicated rigs, targeted at properly identifying the aerodynamic operative point of the cooling fan and at understanding the reason behind the unexpected outcomes obtained. The relevance of this contribution does not lie in the description of the certification process per se, rather in the nonlinear dynamics characterizing the test sequence once the flight activities pointed out an unpredicted behavior.
ABSTRACT The US Army is seeking improvements in the fuel efficiency of their military vehicles.. They have initiated a number of R&D projects aimed at advancing the state-of-the-art of powertrain efficiency including demonstration in a laboratory environment. This effort will set a benchmark for the vehicle integrators, allowing them to improve future vehicle offerings. The SAIC, AVL, Badenoch, QinetiQ and Ker-Train Research team offered powertrain solutions from 7 Tons to 40 Tons that achieved the goal of 44% thermal efficiency and the stringent flexible fuel and emissions requirements. In each of these offerings the team was able to identify modifications to existing engines that allowed dramatic improvements in the thermal efficiency. These efficiency improvements were achieved through a combination of techniques, combustion cycle adjustments using in-cylinder pressure monitoring and precise control of fuel injector timing, and turbo-compounding. For the R&D project, the fuel injector timing will be controlled using commercial engine development hardware and software. The high speed hardware emulates the engine control module but allows the developer to finely tune the fuel injection to maximize the 50% Maximum Fuel Burn point (MFB50) with only limited NOx production. This will be accomplished using a variety of fuels and maintaining the output power to within 2% of the engine’s nominal rating. This paper will describe the fundamental diesel combustion process that must be controlled and techniques for usable power extraction from the waste exhaust gases to provide this performance. It will describe the engine development tools that enable these controls changes to be realized within a vehicle development cycle and retain the baseline engine maturity.
WHY DO WE NEED SIMULATIONS? This paper is intended to provide a broad presentation of the simulation techniques focusing on transmission testing touching a bit on power train testing. Often, we do not have the engine or vehicle to run live proving ground tests on the transmission. By simulating the vehicle and engine, we reduce the overall development time of a new transmission design. For HEV transmissions, the battery may not be available. However, the customer may want to run durability tests on the HEV motor and/or the electronic control module for the HEV motor. What-if scenarios that were created using software simulators can be verified on the test stand using the real transmission. NVH applications may prefer to use an electric motor for engine simulation to reduce the engine noise level in the test cell so transmission noise is more easily discernable.
The Joint Tactical Aerial Resupply Vehicle (JTARV) project is supporting the expansion of the Army's unmanned aerial reconnaissance capability by working to obtain high-quality test data that is scarce for group 2 and group 3 UAVs to validate physics-based models. This paper will evaluate the system identification results from transient testing at a wind tunnel speed of 29.2 kts (15 m/s) of the commercial off the shelf T-motor '28x9.2' rotor which is used on the TRV-80 platform, a group 3 UAV. Rotor angles of attack of -90° to 0°, which represent climb to edgewise flight respectively, were tested in the wind tunnel. Chirp Pulse-Width Modulation (PWM) sweep inputs that varied the ΔPWM amplitude by 50, 100, and 150 (approximately 3%, 6%, and 10% of the steady state PWM) were analyzed to derive an electric motor model that can be used in a physics-based simultion. Additionally, it was found that the identified coefficients of thrust and torque were higher in transient condtions versus steady for most angle of attack test condtions. Lastly, the closed loop wind tunnel system, thrust due to ΔPWM amplitude sweep, was compared to the closed loop system of the TRV-80. A similar motor lag frequency was identified in both systems which shows the relevance of the electric motor model and rotor performance data derived from wind tunnel data preseneted herein.
Revealed in 1941, the Dirigible Helicopter or 'Koun's Craft,' was an ambitious but ill-fated fusion of convertiplane and lighter-than-air technology. This S/VTOL (Short/Vertical Take Off and Landing) concept (a veritable puzzle of diverse airplane parts) was powered by a single, tilting propeller engine and was affixed with wing mounted, helium filled enclosures for additional buoyancy. Dismissed historically as being an eccentric folly of its layman inventor, Korean-American Young Ha Koun, the development of the Dirigible Helicopter has never been thoroughly studied. This paper will examine the origins of this unique design, its creator's possible motivations for building such an aircraft, and successor convertiplane concepts that attempt to achieve the same purpose to this day.
The advent of electric propulsion is revolutionizing the paradigm of rotorcraft design, leading to new electric Vertical Take-Off and Landing (eVTOL) aircraft. Direct drive topologies are common within these new designs, and some designers have chosen to utilize this mechanism for Primary Flight Control (PFC), effectively utilizing the aircraft engines as PFC actuators to control the speed of the rotors. This decision integrates the propulsion and flight control systems, and intrinsically couples the aircraft sizing and control. Four separate tools were exercised throughout this study to conduct a conceptual design exploration of eVTOL aircraft handling qualities. The main tasks for these tools were: 1) aircraft sizing and performance analysis, including the calculation of trim; 2) flight dynamics modeling and analysis; 3) handling qualities-centric control law optimization; and 4) electric motor sizing. Sizing of an RPM-controlled Hexacopter concept explored the dependency of aircraft size to fundamental design parameters: 1) disk loading and 2) blade loading coefficient. Increasing the design disk loading resulted in designs with high agility, but at the cost of significant growth in the design gross weight. Finally, Categories II and III pilot-induced-oscillation (PIO) are known potentially-catastrophic handling qualities deficiencies of fly-by-wire flight control systems such as those expected of eVTOL aircraft. Consideration of PIO predictive metrics in conceptual design control synthesis led to increased PIO robustness.
Wear debris monitoring and analysis is a common practice for the condition assessment of engine and transmission health. Oil debris monitoring (ODM) and electronic chip detectors (ECD) are two common methods deployed for continuous monitoring of oil wetted component health in-flight. This study evaluates the diagnostic performance of the two sensing technologies within controlled rolling element bearing (REB) fault experiments. Progressive visual inspection of the REB spall progression through failure provided a ground truth against which both systems could be compared. Quantifiable metrics of reliability, diagnostic accuracy, provided maintenance interval were defined to create a framework for condition-based maintenance (CBM) program decision making. In summary, it was found that the ODM sensor system provided earlier fault notice, but more so, vastly outperformed the ECD in reliability and avoidance of false positives.
Previous work documented the use of IVHMS data on the U.S. Army's fleet of UH-60 Black Hawk helicopters to update the fatigue lives of six specific components on the A/L and M models. This paper documents a significant expansion of the level of data applied to the usage spectrum, as well as applying it to all components on the aircraft. As a design spectrum for the yet to be fielded Improved Turbine Engine (ITE) equipped UH-60M, changes due to new engine capability needed to be addressed. The new spectrum has been developed and is being used for planning of flight testing. The spectrum along with flight test loads will be used to generate fatigue lives for the new aircraft. Once deployed for several years the spectrum will be reviewed to determine if any changes are needed. This work highlights what the Army considers to be the most significant issues when applying monitored usage to critical fatigue components, and rationale for dealing with issues such as insufficient data for various purposes.
Rotorcraft experience significant vibrations due to periodic aerodynamic forces and moments on the rotor blades and wings. Rotor torque damping is a novel vibration damping method which uses small torque perturbations from the main electric motor to reduce vibrations. The large inertial and aerodynamic rotor loading and relatively high frequency torque perturbations mean that the rotor speed changes are small, so the rotor thrust and flight control performance are not significantly affected. This paper investigates the application of electric motor torque control for damping structural vibrations of an aircraft. The structural dynamics of the aircraft are represented using a finite element model of a quad tiltrotor eVTOL. Using collocated angular rate feedback on all four rotors provides more than 10% damping in controllable modes. The RMS value of flap-wise angular rate can be reduced by 91% with less than 1.2 RPM rotor speed change in response to a 20% vertical step gust in airplane mode. For N/rev disturbance cancellation, an optimal controller is designed assuming known disturbance location and frequency for active vibration control (AVC). The transfer matrix of a single cantilevered wing is calculated and used to feed forward harmonic rotor torques. The wing undergoes aerodynamic disruptions at N/rev and the harmonic controller reduces N/rev shear force and bending moment at the root by 20% and 58% with less than 1 RPM rotor speed change, respectively.
Items per page:
50
1 – 50 of 42476