Maneuver Loads Research Articles

Wavelet-based confirmatory factor analysis: monitoring of damage accumulation factors

Under consideration is further development of the wavelet-based confirmatory factor analysis intended for monitoring of factors responsible for evolution of technical and other systems. According to the proposed approach, the samples of coefficients resulted from discrete wavelet transform of initial parameter time series under study and responsible for different observation periods are considered as values of observed variables in the subsequent confirmatory factor analysis to reveal time history of factor influences and estimate factor interaction. Identification of free factor model parameters is carried out by a novel direct (noniterative) procedure, which is an alternative to traditional local iterative optimization procedures based on the maximum likelihood criteria. A technique for estimating significance of factor model components is discussed. Application of the approach to the analysis of aircraft damage accumulation is given. Analysis of the variance components factor model representing influence of maneuvering load factors occurrences and climatic conditions of basing on aircraft damage accumulation rate revealed differences between national flying training schools to yield statistically significant effect on the process of aircraft damage accumulation under repeated loads.

Structural Analysis of Ageing Pilatus P-3 Engine Mount

The Pilatus Trainer P-3 full scale fatigue test was performed in 1960 to validate a service life of 2’500 FH for the Swiss Air Force. The test showed minor damages after 5’000 FH which were all addressed in modifications for the whole fleet. During the military service life, no cracks were discovered. In 1975, a structural integrity study was done to assess the fatigue life of the fleet based on the Air Force usage. The Swiss fleet was cleared to 3’000 FH and 12’000 flights due to the less severe usage compared to the full scale fatigue test. The Swiss Air Force retired the P-3 fleet in 1995. The P-3 were sold as ̈oldtimers̈ to private people. In Switzerland, 17 aircraft received a civil registration by FOCA. The P-3 had between 3’000 and 3’400 FH during military usage. The current fleet leader has accumulated 4’270 FH. The P-3 full scale fatigue test demonstrated no fatale structural failure at 5’000 FH and 10’000 flights. Due to the high number of FH compared to the test, further investigations are needed to ensure the continuing safe operation of the civil registered P-3 airplanes. First, the available Nz spectrum from the military usage were analyzed and compared to the full scale fatigue test spectrum. Second, two Swiss operators were interviewed to understand the civil usage. The civil usage seems to be heavily depending on the operators, so a conservative engineering approach is necessary to ensure the structural integrity. An assessment was done with collecting all the data from the military usage and reviewing the available information. Some structural critical areas were identified after the first review. Most concern showed the engine mount which was never tested, and only limited static analysis was done in 1958 with no fatigue calculation. The engine mount is a tube structure with welded connection of steel material AISI / SAE 4130. A detailed FE model using ANSYS Workbench Platform was developed and 6 load cases based on EASA CS23 certification standard were applied. The mesh size was reduced to 1mm at the critical areas to model the welded structure. The static analysis was based on FKM criteria. The horizontal spin maneuver exceeds the tensile strength at two local positions. Based on the spectrum information from the test and the Nz exceedance data for the Swiss Air Force usage, a Rainflow Counting Matrix was developed. With the corresponding results for the Nz values, the spectrum could be converted to stresses at the critical locations based on principal stresses of the sub model FE data. For fatigue investigations the flying steady maneuver loads are contributing to the fatigue damage. In the first approach MIL-HDK-5J stress life data was used for max min stress level to see if the fatigue life is in the area of durability life of 107 cycles. Further analysis was done using strain life fatigue analysis which showed at two locations a much shorter life. Inspections at every 100 FH were scheduled to ensure the safe operation under civil usage.

Turning maneuver caused response in an aircraft rotor-ball bearing system

This paper focuses on the nonlinear response caused by aircraft turning flight in a rotor-ball bearing system with internal clearances and Hertzian contact forces. The equations of motion are formulated by using Lagrange equation, where the aircraft turning flight is modeled as a maneuver load. Numerical analyses are carried out to detect the response of the system under maneuver load. The sub-harmonic resonance and irregular motions subjected to the maneuver load are obtained and presented by bifurcation diagrams, amplitude spectrum diagrams, Poincare maps and phase portraits. Two typical patterns that the maneuver load inducing response are found: one is the alternating of period 2 motion and period 1 motion for $$\omega =1{,}350\,\hbox {rad/s}$$ ; the other is the alternating of quasi-periodic motion and period 2 motion for $$\omega =1{,}470\,\hbox {rad/s}$$ . Moreover, mass eccentricity shows a significant effect on the sub-harmonic resonance of the rotor-ball bearing system under turning flight, which is different from that without considering maneuvering flight condition (Bai and Zhang in Int J Nonlinear Mech 50:1–10, 2013). The results obtained in this paper will contribute an understanding of the nonlinear dynamic behaviors of aircraft rotor systems in maneuvering flight.

Usage and Flight Loads Analysis of King Airs in Aerial Firefighting Missions

Flight data of Beechcraft King Airs used as aerial supervision modules and lead aircraft by the U.S. Forest Service are analyzed. The operational conditions experienced by the airplanes are studied as well as the magnitude of the vertical loads. The airplanes are flown in four different types of missions in U.S. Forest Service operation. Data from 1853 flights and seven aircraft are presented. Spectra are shown for gust and maneuver loads, separated using the two-second rule. These are given for the entire flight for various mean sea-level altitudes. In addition, the combined loads are presented and compared to existing data and Federal Aviation Regulation 23. It is shown that these aircraft are operated well within their design envelope. It is demonstrated that, although gust loads are comparable to those outlined in Federal Aviation Regulation 23, the maneuver load frequencies far exceed those of typical airframe use. Although these results do not warrant changes to Federal Aviation Regulation 23, they highlight the need for modified airframe maintenance and inspection procedures as well as the estimated airframe fatigue life.

Comparison of the Flight Loads Spectra of Two Business Jets

Abstract Operational flight loads have been analyzed from two business jets, a Global 5000 and a Global Express XRS. It is shown that both airframes were subjected to nearly the same number of ground-air-ground cycles, even though the flight times were much different. Flights have been divided into various phases, and loads and turbulence data have been categorized by altitude bands within each phase. Cumulative occurrences of incremental vertical gust load factors have been compared and shown to be comparable for the two airframes. Maneuver load factors have been shown to spread over a wider range of values for the 5000 in every phase. This has been confirmed through comparison of combined loads with those from a CRJ100 and an ERJ-145XR. Derived gust velocities, obtained from the load factors are presented in the form of exceedance spectra. These results from both aircraft are shown to agree well

Aircraft Design with Active Load Alleviation and Natural Laminar Flow

Maneuver load alleviation, gust load alleviation, and natural laminar flow are integrated into aircraft conceptual design. The concurrent design of the aircraft and its active load alleviation systems can yield significant gains. The simultaneous application of maneuver load alleviation and gust load alleviation to a short-haul aircraft leads to an 11% reduction in fuel burn and 7% reduction in direct operating cost. These savings are diminished if maneuver load alleviation and gust load alleviation are applied independently. The synergy between active load control and natural laminar flow is also explored. It is possible to invest some of the weight savings from load alleviation to achieve extensive laminar flow. The combination of load alleviation and natural laminar flow increases the fuel and cost savings to upwards of 18 and 11%, respectively. It is concluded that load alleviation can shift the transonic transport optimum toward low-sweep, natural laminar flow wings. Sensitivity studies finally demonstrate that contemporary commercial aircraft should be able to meet the control requirements for effective gust load alleviation.

Reduced-Order Aeroelastic Models for Dynamics of Maneuvering Flexible Aircraft

This paper investigates the model reduction, using balanced realizations, of the unsteady aerodynamics of maneuvering flexible aircraft. The aeroelastic response of the vehicle, which may be subject to large wing deformations at trimmed flight, is captured by coupling a displacement-based flexible-body dynamics formulation with an aerodynamic model based on the unsteady vortex lattice method. Consistent linearization of the aeroelastic problem allows the projection of the structural degrees of freedom on a few vibration modes of the unconstrained vehicle but preserves all couplings between the rigid and elastic motions and permits the vehicle flight dynamics to have arbitrarily large angular velocities. The high-order aerodynamic system, which defines the mapping between the small number of generalized coordinates and unsteady aerodynamic loads, is then reduced using the balanced truncation method. Numerical studies on a representative high-altitude, long-endurance aircraft show a very substantial reduction in model size, by up to three orders of magnitude, that leads to model orders (and computational cost) similar to those in conventional frequency-based methods but with higher modeling fidelity to compute maneuver and gust loads. Closed-loop results for a cantilever wing finally demonstrate the application of this approach in the synthesis of a robust stability augmentation system.

F-35 Full Scale Durability Modeling and Test

The F-35 Joint Strike Fighter program includes three aircraft variants, one of which has been designed and built according to US Air Force requirements, and the other two of which have been designed and built according to US Navy requirements. For all three variants, a system design and development (SDD) configuration aircraft is being subjected to a full-scale durability (FSD) test. In each case, the complete airframe is being subjected to two lifetimes of severe design spectrum loading, with maneuver, catapults/arrestments (carrier variant only) and buffet loads applied as separate, alternating 1000 flight hour blocks during the major test sequence. For the airframe tests, the buffet loads are applied quasi-statically; for the separate vertical tail component tests, they are applied dynamically. In addition, tests of doors and attachments (local tests) are conducted when the full airframe test is down for inspections (as required, for example, between the first and second lifetimes). In this paper, we describe the manner in which the airframe tests were designed, including fatigue spectrum development and test adequacy analyses. In addition, we provide a summary of the test findings to date, along with a description of the analytical simulation for a typical finding. The paper includes an analysis vs test correlation summary that provides an indication of the validity of the fatigue crack initiation (FCI) and fatigue crack growth (FCG) analysis methods used to design the aircraft.

Physics-Based Modeling of Maneuver Loads for Rotor and Hub Design

A hybrid Navier–Stokes free-wake computational-fluid-dynamics methodology, coupled with a multibody dynamics analysis code, has been applied to the UH-60A rotor to study the loads developed during a severe diving turn maneuver (flight counter 11680) characterized by a mean advance ratio of 0.388 and a mean load factor of 1.48. The helicopter undergoing flight counter 11680 experiences the most severe pitch-link loads and root torsional moments recorded in the NASA/Army UH-60A Airloads Program. The feasibility of a quasi-steady loosely coupled approach for predicting severe maneuver loads was explored for predicting critical aerodynamic and aeroelastic phenomena. The structural model is first validated through measured airloads, which helps in decoupling the physics of structural dynamics and aerodynamics. The blade aeromechanical loads, push-rod loads, and harmonic content of blade structural loads for a rotor revolution characterized by peak load factor (revolution 12) have been examined in detail. The prediction of advancing blade-stall phenomena unique to severe maneuvers and the extensive occurrence of stall are examined using contours plots of sectional airloads and through comparisons of surface pressure values.

Operational Assessment of Heavy Air Tankers

Flight data collected by various P2V and P3A air tankers in actual operation over several seasons have been used to construct statistical information on aircraft usage and maneuver load spectra in this role. The data used for this report consist of 5316 flight files, collected during the 2008 and 2009 fire seasons. Flights have been divided into multiple phases, separating the segments when retardant is dropped from others. Basic flight parameters such as airspeed, altitude, flight duration and distance, bank, and pitch angles are shown in statistical form. Vertical accelerations are used to construct diagrams for comparison with the manufacturers’ data. Airframe cycle times are shown to be approximately 45 min. Several coincident flight events are shown and compared with operational limits when available. The altitudes, the normal accelerations, and the airspeeds have not revealed any operation outside of the established boundaries. Relatively large dive and roll angles were prominently present during all but the drop phases. Nonetheless, they are shown to remain within the limits stated by the manufacturer or imposed by the Federal Aviation Administration type certificate data sheets. Cumulative occurrences of maneuver loads have been examined and compared with data from other sources when available. It is shown that the cumulative occurrences of maneuver load factor varies by flight phase, with the largest values associated with the drop and exit phases. For the overall flight, maneuver loads were shown to occur at lower frequencies than those stated in Military specification MIL-A-8866C(AS).

Constant-excitation caused response in a class of parametrically excited systems with two degrees of freedom

This paper focuses on the response in parametrically excited systems caused by constant excitation. Taking a maneuvering cracked rotor system as an example, we formulate the vibration equations with two degrees of freedom, in which the breathing of the crack constitutes parametric excitation, and the maneuver load of the maneuvering rotor is simplified as a constant excitation, and it is supposed that the rotor system is balanced without the consideration of eccentricity. By solving the equations with harmonic balance method, each order of harmonic components related with the rotating speed and the constant excitation is derived to analyze the corresponding resonance of the system. Results show that the constant excitation plays a decisive role in the parametrically excited primary and super-harmonic resonances of the system that agrees with the gravity dominance in common cracked rotor systems without maneuver load. And the stronger the constant excitation, the greater the resonances. Moreover, the orientation of the constant excitation makes a great impact on the parametrically excited primary resonance, but does not have a significant effect on the parametrically excited super-harmonic resonances. Results implies that constant excitation may increase the parametrically excited super-harmonic resonances of the cracked rotor systems, which is disadvantageous to the operating of the system. From another point of view, however, constant excitation can be used for early detection of crack faults in rotor systems.

Analysis of 1/2 sub-harmonic resonance in a maneuvering rotor system

This paper focuses on the 1/2 sub-harmonic resonance of an aircraft’s rotor system under hovering flight that can be modeled as a maneuver load G in the equations of motion. The effect on the rotor system is analyzed by using theoretical methods. It is shown that the sub-harmonic resonance may occur due to maneuvering flight conditions. The larger the eccentricity E and the maneuver load G, the greater the sub-harmonic resonance. The effects of nonlinear stiffness, damping of the system, maneuver load, and eccentricity on the sub-harmonic resonance region in parameter planes are also investigated. Bifurcation diagrams of the analytical solutions are in good agreement with that of the numerical simulation solutions. These results will contribute to the understanding of the nonlinear dynamic behaviors of maneuvering rotor systems.

Damping and Inertia Coefficients for Two End Sealed Squeeze Film Dampers With a Central Groove: Measurements and Predictions

Aircraft engine rotors, invariably supported on rolling element bearings with little damping, are particularly sensitive to rotor imbalance and sudden maneuver loads. Most engines incorporate squeeze film dampers (SFDs) as a means to dissipate mechanical energy from rotor motions and to ensure system stability. The paper experimentally quantifies the dynamic forced performance of two end sealed SFDs with dimensions and an operating envelope akin to those in actual jet engine applications. The current experimental results complement and extend prior research conducted with open ends SFDs (San Andrés, 2012, “Damping and Inertia Coefficients for Two Open Ends Squeeze Film Dampers With a Central Groove: Measurements and Predictions,” ASME J. Eng. Gas Turbines Power, 134, p. 102506). In the tests, two journals make for two SFD configurations, both with a diameter D = 127 mm and nominal radial film clearance c = 0.127 mm. One short length damper has film lands with extent L = 12.7 mm, while the other has 25.4 mm ( = 2L) land lengths. A central groove of length LG = L and depth at ¾ L separates the film lands. A light viscosity lubricant is supplied into the central groove via three orifices, 120 deg apart, and then flows through the film lands whose ends are sealed with tight piston rings. The oil pushes through the piston rings to discharge at ambient pressure. In the tests, a static load device pulls the damper structure to increasing eccentricities (maximum 0.38c) and external shakers exert single-frequency loads 50–250 Hz, inducing circular orbits with amplitudes equaling ∼5% of the film clearance. The lubricant feed and groove pressures and flow rates through the top and bottom film lands are recorded to determine the flow resistances through the film lands and the end seals. Measured dynamic pressures in the central groove are as large as those in the film lands, thus demonstrating a strong flow interaction, further intensified by the piston ring end seals which are effective in preventing side leakage. Dynamic pressures and reaction loads are substantially higher than those recorded with the open ends dampers. Comparisons to test results for two identical damper configurations but open ended (San Andrés, 2012, “Damping and Inertia Coefficients for Two Open Ends Squeeze Film Dampers With a Central Groove: Measurements and Predictions,” ASME J. Eng. Gas Turbines Power, 134, p. 102506) demonstrate at least a threefold increase in direct damping coefficients and no less than a double increment in added mass coefficients. Predictions from a physics-based model that includes the central groove, the lubricant feed holes, and the end seals' flow conductances are in agreement with the test results for the short length damper. For the long damper, the predicted damping coefficients are in good agreement with the measurements, while the added masses are under-predicted by ∼25%.

On the Integration of Real-Time Diagnosis and Prognosis for Scheduled Maintenance Optimization

Helicopters are very critical aircrafts from the point of view of fatigue loads. A structural damage could grow fast to critical size because of the wide range of manoeuvre loads. As a consequence, scheduled maintenance is a major cost during the operative life of the aircraft. The most advanced methodologies for Structural Health Monitoring and Prognostic Health Management available today are aimed to provide a real-time structural diagnosis, thus maximizing the availability of the helicopter and reducing the maintenance costs. However, on-board diagnostic systems might be gradually introduced in the current maintenance procedures, trying to minimize the risk associated to these newly developed technologies. The work presented inside this paper is about the simulation of the integration of real-time diagnosis and prognosis into a typical scheduled maintenance procedure. A diagnostic unit is capable for anomaly detection and damage quantification (in terms of crack length). It is trained with Finite Element simulated damages and tested with real experimental data. The diagnostic output is then processed in a particle filter algorithm, based on sequential importance sampling technique, aimed at refining the estimation of the structural condition as well as to update the inference on the residual useful life distribution. The coupling of real-time diagnosis with off-line measures (taken during scheduled maintenance stops) is analyzed and applied to a damage tolerant structure, trying to outline the advantages and drawbacks of the proposed approach.

Nonlinear vibration phenomenon of an aircraft rub-impact rotor system due to hovering flight

This paper focuses on the nonlinear vibration phenomenon caused by aircraft hovering flight in a rub-impact rotor system supported by two general supports with cubic stiffness. The effect of aircraft hovering flight on the rotor system is considered as a maneuver load to formulate the equations of motion, which might result in periodic response instability to the rotor system even the eccentricity is small. The dynamic responses of the system under maneuver load are presented by bifurcation diagrams and the corresponding Lyapunov exponent spectrums. Numerical analyses are carried out to detect the periodic, sub-harmonic and quasi-periodic motions of the system, which are presented by orbit diagrams, phase trajectories, Poincare maps and amplitude power spectrums. The results obtained in this paper will contribute an understanding of the nonlinear dynamic behaviors of aircraft rotor systems in maneuvering flight.

Damping and Inertia Coefficients for Two Open Ends Squeeze Film Dampers With a Central Groove: Measurements and Predictions

Aircraft engine rotors are particularly sensitive to rotor imbalance and sudden maneuver loads, since they are always supported on rolling element bearings with little damping. Most engines incorporate squeeze film dampers (SFDs) as means to dissipate mechanical energy from rotor vibrations and to ensure system stability. The paper quantifies experimentally the forced performance of a SFD comprising two parallel film lands separated by a deep central groove. Tests are conducted on two open ends SFDs, both with diameter D = 127 mm and nominal radial clearance c = 0.127 mm. One damper has film lands with length L = 12.7 mm (short length), while the other has 25.4 mm land lengths. The central groove has width L and depth 3/4 L. A light viscosity lubricant flows into the central groove via three orifices, 120 deg apart and then through the film lands to finally exit to ambient. In operation, a static loader pulls the bearing to various eccentric positions and electromagnetic shakers excite the test system with periodic loads to generate whirl orbits of specific amplitudes. A frequency domain method identifies the SFD damping and inertia force coefficients. The long damper generates six times more damping and about three times more added mass than the short length damper. The damping coefficients are sensitive to the static eccentricity (up to ∼ 0.5 c), while showing lesser dependency on the amplitude of whirl motion (up to 0.2 c). On the other hand, inertia coefficients increase mildly with static eccentricity and decrease as the amplitude of whirl motion increases. Cross-coupled force coefficients are insignificant for all imposed operating conditions on either damper. Large dynamic pressures recorded in the central groove demonstrate the groove does not isolate the adjacent squeeze film lands, but contributes to the amplification of the film lands’ reaction forces. Predictions from a novel SFD model that includes flow interactions in the central groove and feed orifices agree well with the test force coefficients for both dampers. The test data and predictions advance current knowledge and demonstrate that SFD-forced performance is tied to the lubricant feed arrangement.

Optimal trends in Manoeuvre Load Control at subsonic and supersonic flight points for tailless delta wing aircraft

Manoeuvre Load Alleviation (MLA) is the concept of redistribution of forces and moments on airframe through optimal actuation of control surfaces. The effect is to minimise the stresses near wing root and improve fatigue life of airframe components, such as, the frequently studied wing root fitting connecting wing to the fuselage. The configuration is a typical tailless delta wing aircraft having inboard and outboard Elevons (a combination of Elevator and Aileron) placed symmetrically on starboard and port side. The problem is formulated as minimisation of independent conflicting multiple objectives of wing root bending and twisting moments, with stability equations as equality constraints and actuator hinge moments as inequality constraints. Hybrid method is used as optimisation technique, which is evolved from the combination of heuristic based Elitist non-dominated Sorting Genetic Algorithm (NSGA-II) and calculus based Goal Programming methods. The primary objective of this paper is to bring out the structural benefits attainable for some of the combined critical manoeuvre load cases at subsonic and supersonic flight points. The combined critical manoeuvre load cases demarcate the limit to which airframe is loaded. The minimisation achieved is better for flight points at supersonic as compared to subsonic Mach numbers.

Multi-model convex design of a scheduled lateral feedforward control law for a large flexible BWB aircraft

AbstractThis paper presents the robust multi-objective design of a lateral feedforward control law for a large flexible blended wing body (BWB) type passenger aircraft. A multitude of both time- and frequency-domain design constraints are considered to achieve manoeuvre loads alleviation and the desired tracking performance. A convex multi-model feedforward control design is utilized. Scheduled and fixed control laws are obtained and numerically validated and compared on several mass cases of the aircraft model, showing high loads alleviation performance while obeying all stringent design constraints. The benefit of scheduling for a fuel mass parameter is demonstrated.

Aeroelastic optimization design for wing with maneuver load uncertainties

An aeroelastic optimization design methodology for air vehicle considering the uncertainties in maneuver load conditions is presented and applied to a structural design process of low-aspect-ratio wing. An aerodynamic load correction model is developed and used to predict the critical load conditions with the perturbations of theoretical linear aerodynamic forces and experimental aerodynamic forces from wind-tunnel test, when concerning the uncertainties in use of theoretical linear and experimental aerodynamic forces. Three objective functions of critical loads are defined. The load evaluations for three wing sections are investigated in four characteristic maneuvers, and the most critical load conditions are confirmed by using the sequential quadratic programming method. On this basis, the aeroelastic optimization design employing the genetic-gradient hybrid algorithm is conducted, in which the objective is to minimize structural mass subject to the constraints of stress, deformation and flutter speed. The resulting optimal structure is heavier than the one simply based on the theoretical linear or experimental aerodynamic forces. However, it is more robust when encountering the critical load conditions in actual flight due to the consideration of uncertainties in aerodynamic forces in the early design phase, thereby, the risk of structural redesign can be reduced.

Multi-Objective Optimization Methods Applied for Manoeuvre Load Control on Combat Aircraft wing

Manoeuvre Load Control (MLC) is a subset of Active Control Technology (ACT), deals with optimal rotation of control surfaces to effectively redistribute the forces and moments on airframe, resulting in structural benefits and performance improvements. The present formulation consists of flight control parameters as design variables; bending and twisting moment at the wing root as conflicting independent multiple objective functions, together with aircraft stability/trim equations as equality constraints and actuator hinge moments as inequality constraints. This work attempts to bridge the gap observed with computational approaches and contemporary flight test programs, by treating MLC problem as multi-objective functions. Several multi-objective optimization methods such as Goal programming method, Multi-objective Evolutionary Algorithm (MOEA) and its hybrid form, which is a combination of the above two methods are explored for their capabilities when applied to this real world problem. Another significan...