Browse Topic: Failure analysis
RPM-controlled hexacopters offer mechanical simplicity and inherent redundancy, but are unable to re-trim under all failure cases in forward flight. This paper investigates the use of reverse-enabled rotors as a means of expanding the attainable trim envelope and improving fault tolerance in RPM-controlled hexacopters. Isolated rotor experiments are conducted to characterize thrust and torque behavior under forward and reverse rotation, providing validation data for aerodynamic modeling. A blade-element-based model implemented in the Rensselaer Multicopter Analysis Code (RMAC) is then used to perform comprehensive trim analyses for a 1200-lb-class hexacopter in hover and in cruise at the best-range speed of 65 kts. Post-failure trim solutions are evaluated for four configurations, including edge-first and vertex-first orientations with different rotor spin directions. Results show that enabling reverse rotation allows trim recovery for all single-rotor failure cases in cruise, including aft-rotor failures that are not trimmable with conventional RPM-controlled rotors. A systematic comparison of peak rotor torque, peak rotor power, and total aircraft power reveals that failure severity is governed primarily by yaw moment deficits arising from the combined loss of hub torque and aerodynamic drag. Among the configurations examined, the edge-first configuration with a counter-clockwise spinning rotor 1 exhibits the lowest rotor torque and rotor power requirements, post-failure.
This study presents a comprehensive analysis of single-rotor failure tolerance for a classical octocopter configuration, examining both hover and forward flight at the best range speed. Using a state-of-the-art eVTOL comprehensive analysis to retrim the octocopter post-failure, the redistribution of rotor thrust, torque, and power following individual rotor failures was quantified, along with resulting aircraft-level power penalties. In hover, orthogonal rotors to the failed rotor provide primary lift compensation, the opposing rotor operates mostly unchanged, and the four opposite spinning rotors primarily provide pitch/roll moment compensation. This results in a total aircraft level power increase of approximately 10.4%, roughly half that of comparable hexacopters. In forward flight, at best range cruise speed, load redistributions were again calculated for various individual rotor failures. In the worst case, a maximum individual rotor torque increase of 62% and power increase of 108% was observed, while total aircraft power requirements increased between 7-12%. These results demonstrate the fault-tolerant capabilities of octocopters and provide practical guidance for propulsion system sizing, energy management, and failure-case assessment on classical octocopters.
As per Committee/Henry E. Harschburger recommendations
This study examines the ability of a large (1200 lb gross weight) hexacopter with collective pitch controlled rotors to tolerate single motor failure. The hexacopter is considered in various orientations, and the vehicle is trimmed with one motor inoperative (OMI). Unlike RPM-controlled hexacopters, which were trimmable but uncontrollable in hover, and were untrimmable in cruise with an aft-rotor failure; with pitch-control the hexacopter is controllable in hover as well as trimmable for failure of any rotor in cruise (including an aft rotor failure). The study examines how pitch controls, and thrust are redistributed amongst the operational rotors, post-failure, for the different hexacopter orientations. For each case, the maximum thrust and torque increases on any individual rotor, and the total power increase, post-failure is examined. It is found that the hardest to trim cases are those where the hub torque and the hub drag induced yaw moment of the failed rotor add, and fault compensation for these cases usually comes at a high cost in terms of torque penalty and power requirement (necessitating use of more powerful motors and batteries). The results from the study indicated that operating the pitch-controlled hexacopter in an edge-first configuration with a clockwise spinning rotor 1 is the best choice from a fault tolerance perspective.
Electric aviation is advancing rapidly, with aircraft from manufacturers like Joby and Archer well on their way to certification, aircraft electrification will continue and begin to apply to larger aircraft. To support larger electrified rotorcraft, rotors will need to grow if disc-loading and hover efficiency are to be maintained. A consequence of this is the need to reduce rotor speed to maintain an acceptable acoustic signature, especially for operation in urban environments. Most current applications utilize radial flux motors, sometimes with a reduction gearbox. Gearboxes can improve overall propulsion system power density by enabling higher motor speeds but are generally not preferred as they introduce additional potential failure modes and maintenance schedules. In this paper a holistic approach is used to understand the trade-offs between rotor and motor and their consequences on propulsion system power density.
The vertical flight industry is on its way to a transformative era, with autonomous technologies set to alter aerial vehicle operations. While it seems certain that fully autonomous helicopters will eventually be deployed for a variety of missions, some high-stakes situations—like medical evacuations (MEDEVAC)—will for the foreseeable future demand human participation in the form of Emergency Medical Care-giving Crew. This study describes the testbed built to run and investigate hypothetical future situations in which a helicopter is autonomously piloted while a human medic with no aviation training, subjected to aviation and medical emergencies, manages patient care onboard. A total of 22 participants, with emergency medical technician certification, nursing or a medical board certification, were invited to run and evaluate the use of AI pilot (AP) in different scenarios of medical evacuation under the following emergencies: medical, empty fuel tank, pressure sensor miscalibration, and engine failure. A comprehensive evaluation of both objective and subjective performance metrics revealed that novice medical professionals could effectively execute medical evacuation operations in conjunction with an AI pilot, even during unforeseen circumstances. The analysis of response times unveiled distinct perspectives on how medics perceive and manage various emergency situations when an AP functions as a collaborative and effective team member.
Gearbox casing cracks in helicopters would be critical impacting the aircraft's reliability and operation safety directly. The Defense Science and Technology Group (DSTG) HUMS2025 gearbox casing failure data set was the unexpected result of a test stand operation. The gearbox undergoes high cycle (> 400 acquisitions) under high torque (100% and 125% nominal torque) conditions. We hypothesized that the any cracking would be due to the planet/ring gear interaction. A condition indicator (CI) would be sensitive to a crack feature and this would be sensitive to change in gearbox torque. This paper explores the development of both a cyclo-stationary based CI (frequency-domain) and a time synchronous average CI (time-domain). The trend shows that proposed methods can help to detect localized defects in gearbox casing at an early stage and trend as the crack propagates before catastrophic failure occurs.
Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work develops structural fatigue life prediction capability in state-of-the-art emerging progressive damage failure analysis tool CDMat developed at the University of Texas Arlington Advanced Materials and Structures Lab. While JIntegral, implemented in CDMat, appears as the most objective and rigorous approach to predict delamination growth-based fatigue life of composite structures, the key material properties of the J-Integral fatigue model have not been measured with the adequate accuracy. This work addressees a fundamental challenge of eliminating the established and routine assumptions and developed a methodology to determine the key material properties meeting the material input data requirements for the JIntegral based structural fatigue life prediction models. This work generated input data for fatigue crack growth propagation, including optimized input data parameters for the in-situ cohesive law and modeling as-manufactured specimen conditions. In addition, a methodology to account for and characterize the effects of fiber bridging in static tests is presented. It uses a simple standard unidirectional panel for testing, before a component is manufactured, fabricated from the same batch of prepreg to qualify the effects of fiber bridging. Also, this work attempted to determine a minimum conservative initial crack size to streamline the fatigue crack propagation prediction. Fatigue predictions have been demonstrated on a representative composite skin–hat stiffener sub-component section and compared with tests.
Dufour Aerospace designs and manufactures an automated tilt-wing aircraft for critical cargo delivery missions. Emphasizing operational efficiency, the platform integrates path generation and tracking techniques tailored for the unique dynamics of tilt-wing flight and builds upon the existing lower level control. While there exist a myriad of methods for high-level aircraft automation ranging from PID to MPC, they often require a trade-off between complexity and the capability to handle non-linear dynamics of the system they are controlling. Hence, a lightweight, deterministic geometric path generation approach using clothoid-based transitions between three waypoints and a robust SO(3)- based path tracking controller adapted for tilt-wing dynamics are presented. Additionally, a high-level automation framework is introduced that includes failure mode handling for GNSS loss and communication breakdowns. This system ensures mission continuity and operational safety while supporting flexible mission planning. The methods are validated through extensive flight testing on both small and large-scale aircraft. The latter prove the scalability, safety, and reliability of the presented solution for tilt-wing aircraft automation and enhance the aircraft's capabilities in real-world emergency response and complex operational scenarios.
As per certification requirements, for a large rotorcraft that does not meet the Category A requirements, the Height-Velocity (HV) avoid region must be determined in total power failure condition. The development of a digital twin representative of the real rotorcraft behaviour allows to reduce flight testing hours and to increase flight tests safety, especially in such critical conditions, thus decreasing risks and costs. In this work, an extensive simulation activity has been carried out to generate HV charts for a medium twin-engine helicopter in case of loss of both engines. An in-house software that emulates pilot logics has been exploited, coupled with a Flightlab model representative of the rotorcraft and validated against flight data. Manoeuvres performed after a dual engine failure were simulated starting from an all engine operative hover out of ground effect (HOGE) and in ground effect (HIGE) or level flight condition until landing, in a grid of heights and velocities and for different weights and altitudes combinations. Sensitivity analyses on the parameters that affect the manoeuvres the most have been performed and the effects of the assumptions and requirements at touchdown on the HV avoid region investigated.
ABSTRACT The use of active inceptor systems allows for control of the aircraft even during mechanical failures within the control inceptor. For the specific case of isometric failure, whereby the inceptor 'freezes' in position, a virtual force displacement model is used to continue to provide control input. Testing on DLR's experimental helicopter (ACT/FHS) has shown the potential to encounter pilot-induced oscillation (PIO) tendencies when flying using this mode. This paper presents results from a simulation campaign undertaken to determine whether PIOs could be exposed through this use of control and/or the resultant severity. The results show that control limiters cause severe PIOs during the isometric failure. Unacceptable failure characteristics were found for six different vehicle configurations and PIOs were exposed by all four pilots. PIO incipience was predicted through the use of offline tools. In future, it is recommended that specific PIO investigations are undertaken during the evaluation of active inceptor failure modes.
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.
This paper deals with the influence of engine failure during hover on the wiring harness mass of electrical Vertical Take-Off and Landing (eVTOL) aircraft. It starts by presenting possible strategies which can be used to distribute the additional thrust needed during an engine failure among the remaining engines. The most efficient strategy is selected and the impact of different single engine failures on the overall thrust share, while using this strategy, is discussed. The paper proceeds by applying the selected thrust compensation strategy to the mission simulation of three common reference models, which are representative of current eVTOL aircraft configurations. This simulation is used to determine the worst flight phase for the One Engine Inoperative (OEI) condition to occur. The main purpose of the simulation is to optimize the wire sizes of the wiring harness of each configuration while satisfying different design objectives. The results of these optimizations are used to discuss the criticality of each engine failure and its influence on the wiring harness design, especially its mass. It concludes with design recommendations for the wiring harness of eVTOL aircraft.
This paper presents a framework with associated concepts to define a method of compliance for the failure rate requirements of the Army Military Airworthiness Certification Criteria (AMACC), Chapter 5, for fleet qualification and first flight. The fleet failure rate requirement is paraphrased as less than one structural failure in 20 million flight hours at 95% confidence and applies specifically to fatigue failure of primary structural elements (PSEs). This method of compliance assumes a reliability model sufficient to support an analytical failure rate that bounds the uncertainty arising from practical constraints of an aircraft qualification program and aims to optimize the competing objectives of safety of flight and operational capability. The requirements as defined verify that the aircraft system will perform as intended for a specified fleet life and flight test program duration and serve as the analytical basis for the assessment of emergent issues identified throughout the product life cycle.
ABSTRACT A flight simulation model for the UH-60 Black Hawk based on Sikorsky's GenHel model is modified to simulate a locked failure of a main rotor swashplate servo actuator and is compensated by using the stabilator as a redundant control effector. Steady state trim analysis is performed to demonstrate feasibility of trimmed flight in various conditions with different locked servo actuator positions for the forward, aft, and lateral actuators. A model-following, linear dynamic inversion controller is implemented and modified to account for locked actuator position. Post-failure, the control mixing and feed-forward control coupling terms are reconfigured to partially reallocate the control authority in the longitudinal axis from the main rotor longitudinal cyclic to a symmetric deflection of the stabilator. This is done by manipulation of only the control allocation relating pilot stick inputs to servo actuator positions, the feedback control gains and mechanical rigging between servo actuators and rotor pitch controls remain identical to the baseline controller. Flight simulation results demonstrate the ability of this reconfiguration to compensate for locked failure of the forward main rotor swashplate servo actuator, as well as the ability of the aircraft to decelerate from cruise at 120 knots to 50 knots which is less than the published safe rolling landing speed of 60 knots. A similar range of locked positions of the forward and aft actuators is demonstrated to be feasible for aircraft recovery using control of the stabilator. Feasibility of aircraft recovery for locked positions of the lateral servo actuator is also considered.
The AW609 tiltrotor features a unique high-mounted wing with rotatable nacelles positioned at the wing tips, it is capable of operating both in airplane and vertical flight mode. To achieve suited protection of the occupants during emergency landing, the wing - which is particularly stiff in order to sustain the heavy weights at the tips, where rotors, engines and transmissions are positioned - implements a controlled failure mechanism at root, so that during emergency landings it breaks and unloads the fuselage of the weight of wingbox and nacelles, thus avoiding catastrophic collapse. As the effectiveness of such mechanism was never demonstrated under impact conditions, certification agencies requested an empirical validation through experimental testing. The test was carried out July 2022 at Polytechnic of Milan, Italy; the present work details the Test activity, from its preliminary phases to the Test Day, to the analyses of its outcomes.
A state-of-the-art emerging progressive damage failure analysis tool CDMat has been successfully applied to multiple material systems on open-hole tension and compression, and double shear bearing laminate coupons under static and fatigue loading including simulation to ultimate failure. CDMat also successfully demonstrated component-level strength/fatigue analysis under the Air Force Composite Airframe Life Extension (CALE) and the Fail-Safe Technologies for Bonded and Unitized Composite Structures (FASTBUCs) Programs. Building on the success of CDMat an integrated software solution for certification and sustainment of rotorcraft primary composite structures is being developed. A method and an algorithm for fatigue crack growth simulation in laminated structures are proposed to improve the accuracy of CDMat fatigue predictions. The method is based on using cohesive material model, tracking material points at the crack front, and calculating the pointwise energy release rate employing the J-integral. The algorithm was implemented as a set of user material subroutines developed within the framework of explicit finite element formulation for ABAQUS. The effectiveness of the method is demonstrated on several examples of Mode I and II fatigue crack growth.
The airframe digital twin analysis framework developed at the National Research of Canada is being transposed to safe life applications for rotorcraft components. A probabilistic safe life prediction approach, consisting of uncertain material property data and uncertain load spectra is used to calculate risk assessment metrics, such as the cumulative probability of failure, the hazard rate, and the average hazard rate as a function of time. A demonstration of this approach is presented for a CH-146 Griffon component, for which the uncertain loads are estimated from a model developed through machine learning. This preliminary assessment shows the feasibility of using digital twin concepts as a viable alternative to traditional deterministic life predictions, with the potential to reduce maintenance costs and increase aircraft availability.
This paper examines the Handling Quality Rating (HQR) of the Model-Based Pilot Controller (MBPC) in failure scenarios within the Automatic Flight Control System (AFCS). The MBPC aims to automate the testing of malfunctions in the AFCS of the T625 Gökbey platform. It is constructed using optimal control and estimation theory, with the cost function representing human characteristics determined by weighting matrices. The optimal values of weighting matrices that minimize the cost function are achieved via Genetic Algorithm. This algorithm utilized to systematically minimize user-defined cost functions tailored to optimize performance for selected maneuvers within the scope of ADS33E-PRF, considering user-defined constraints. Time-domain metric performance is provided for two maneuvers: vertical maneuver and hovering turn. The HQRs of the MBPC evaluated according to Power Frequency and Inceptor Peak Power-Phase (IPPP) metrics. The MBPC satisfies the ADS33 desired performance criteria in case of a failure scenario in AFCS throughout the maneuver, achieving satisfactory handling quality results.
This paper investigates the role of the aerodynamic torque on propeller whirl flutter stability. The generalized force due to the torque is first computed and subsequently included in the equations of motion of a rigid propeller-pylon system. Preliminary evaluations indicate that the torque modifies the real part of the backward and forward modes, providing a stabilizing effect on powered propellers. Analyses are conducted on a 3-bladed propeller driven by an electric motor. Stability predictions are obtained with a simple analytical model and validated by multibody simulations coupled with a mid-fidelity aerodynamic solver, based on a vortex particle method. Furthermore, a simple control law acting on the propeller's collective pitch and rotational speed is presented. The control variables are modified to increase the whirl flutter stability margins, without altering the trim conditions of the aircraft. Results demonstrate the effectiveness of the proposed control strategy, although propeller efficiency is reduced. In a failure scenario, the control law can be exploited to compensate for a weak mounting stiffness, allowing the aircraft to land safely.
Health and Usage Monitoring Systems installed on modern rotorcraft can be used for substantiating component life extensions leveraging actual loads and usage information, reducing uncertainty from assumed loads and usage. One challenge inhibiting fielding such extensions is lack of clearly defined methodology to robustly determine the criticality of the HUMS application. This paper proposes a semi-quantifiable methodology for determining criticality. It is applicable to both loads- and usage-based approaches for part-number-level life adjustments. The approach relies on baseline component design data and is independent of HUMS data to eliminate the possibility of a HUMS errors affecting the bounding criticality assigned to the application. It takes into account the specifics of the component, including the SN working curve, relevant component failure mode severity, and validated compensating provisions that either reduces the likelihood of the component failure mode and/or provides significant "hidden" conservatism not explicitly considered in the assigned uncompensated component failure severity level.
Rolling element bearing failures form one of rotating equipment's most critical failure modes. Vibration analysis has been successfully used for bearing fault detection and diagnostics but does not estimate the spall length of the bearing. An estimate of the spall length would provide insight into the degrading reliability of a drivetrain as the fault propagates. This would improve the timeliness of scheduling a maintenance action. In this paper, a synthetic tachometer signal is generated from the bearing fault itself. It is synchronous to the rolling element, allowing for a time-domain representation of waveform using the time-synchronous average. From this, an estimate of the length of the bearing fault can be determined.
The Eagle Flight Research Center (EFRC) at Embry-Riddle Aeronautical University (ERAU) is investigating the handling qualities of partial and full rotor failure modes of a multi-rotor vehicle testbed employing Distributed Electric Propulsion (DEP) systems intended for Advanced Air Mobility (AAM) vehicles. In order to pave the way for commercial operations, the AAM industry requires a deeper understanding of the handling characteristics and the vehicle's dynamics and controllability under rotor failure conditions. The objective of the research performed at the EFRC centered around designing and testing different thrust and moment control allocation methods for an electric Vertical Take-Off and Landing (eVTOL) vehicle, in addition to assessing their performance in both nominal and failure modes of operation. This paper focuses on analyzing the predicted handling qualities for a full-scale quadrotor testbed vehicle with RPM, collective, and cyclic blade pitch control allocation. The study uses flight-test data to develop a bare airframe dynamic model to optimize the controller gains, which are used to evaluate the predicted handling qualities though development of vehicle specific Handling Qualities Task Elements (HQTE) maneuvers. The evaluation is performed using an in-house built flight simulator together with the Cooper-Harper Rating Scale (CHRS) for Handling Qualities test evaluation. The handling qualities were evaluated and compared for different control strategies and a method for quantitative evaluation of handling qualities is presented.
Dual three-phase motors are simulated on a 1200 lb hexacopter to examine its ability to tolerate any single electronic speed controller (ESC) failure. Redundancy is added, as each motor includes two sets of stator windings, powered by electrically independent ESCs, which drive a single motor shaft. Flight control laws are tuned on the healthy aircraft to meet standard flying-qualities specifications. Power constraints are embedded in the controller to enforce both instantaneous and continuous power limits for each ESC. With the proposed dual three-phase motor, the aircraft is shown to be capable of trimming in all single ESC failure cases, including the case of aft motor failure in forward flight. Time-domain simulations demonstrate the aircraft's ability to re-establish trim and recover from any single ESC failure in hover and forward flight. Flying-qualities specifications are re-evaluated for the post-failure aircraft, showing minor, but not debilitating, degradation in several flying-qualities metrics. The post-failure aircraft's ability to follow commands without violating power constraints is verified through time-domain simulations.
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