Pitch Link Loads Research Articles

LTI/LQE: A Physics-Based Approach for Online Estimation of Dynamic Loads in the Rotating Frame

Accurate and online load monitoring of vital components located in the rotor system is important for not only inferring usage and estimating fatigue in those components but also for developing load alleviation/limiting control schemes. In the literature, on-board models for online (i.e., real-time) estimation of rotor component loads using available in-flight aircraft fixed-frame measurements are developed using neural networks, a statistical-based approach, or a combination of both. These models are aircraft-specific, do not capture the higher harmonic dynamics of the rotor system vibratory loads, and their synthesis requires obtaining a lot of data. To remedy these issues, this paper introduces a purely physics-based approach to online estimation of rotor system loads. The proposed approach can be applied to any rotary-wing vehicle, does not rely on a large amount of data, and captures the higher harmonic dynamics of the rotor system vibratory loads. The developed model entitled LTI/LQE is synthesized using a linear time invariant (LTI) model of helicopter-coupled body-rotor-inflow dynamics and a linear quadratic estimator (LQE), that is designed to correct the LTI model state response using fixed system measurements. The estimation fidelity of the LTI/LQE model is evaluated in simulation using a high-fidelity nonlinear model of a generic helicopter for online prediction of rotor blade pitch link loads arising from vehicle maneuvers. Results obtained using the LTI/LQE model revealed an interesting finding. It was found that the N/rev (where N is the number of blades) fixed system load measurements have information that can be leveraged to make an inference about the N/rev dynamic loads in the rotating system.

Rotor Power Savings and Pitch-Link Load Reductions with Spanwise-Varying Active Camber Morphing

The effect of spanwise-varying active-camber morphing on helicopter rotor power and pitch-link loads was investigated computationally for advance ratios ranging from hovering to high-speed forward flight (μ=0.35) and was compared to results obtained with spanwise uniform camber actuation. Different actuation strategies were applied, such as constant (0P) deflection and superimposed harmonic actuation based on a 0P deflection combined with 1P (one per revolution) and 2P harmonics. A comprehensive rotor aeromechanics model of an isolated Bo-105 rotor, including elastic blade modeling and free vortex wake aerodynamics, was used for this analysis. Two different active-camber section sizes were investigated, one ranging from 0.22R to 0.9R and one from 0.5R to 0.9R. The greatest benefits from active-camber actuation were found in hover and high-speed flight. It was possible to reduce both the rotor power consumption and the peak-to-peak pitch-link loads. A relevant contribution of the spanwise variability to these power gains was obtained in hovering flight. Toward higher flight speeds, the effect of spanwise variable camber morphing became less important in terms of rotor power savings. However, it allowed a notable reduction of the actuation amplitude far outboard on the rotor blade.

Air and Structural Loads Analysis of a 5-Ton Class Rotorcraft in a Pull-Up Maneuver Using CFD/CSD Coupled Approach

The air and structural loads of a 5-ton class light helicopter (LH) rotor in a 2.24 g pull-up maneuver are investigated using a coupling between the computational structural dynamics (CSD) and computational fluid dynamics (CFD) methods. The LH rotor is characterized by a five-bladed system with elastomeric bearings and inter-bladed dampers. The periodic trim solution along with the converged CFD/CSD delta airloads obtained in steady-level flight (advance ratio of 0.287) are used to perform the transient CSD maneuver analysis. The resulting vehicle attitude angles and velocity profiles of the aircraft are then prescribed in the quasi-static (QS) CFD maneuver analysis. It is demonstrated that the present QS approach provides an effective means for the maneuver loads’ analysis. The important flow behaviors such as BVI (blade–vortex interaction)-induced oscillations and the negative pitching moment peaks met in maneuver flight are captured nicely with the proposed method. Either the vortex trajectories or the surface pressure distributions are examined to identify the sources of the oscillations. A loose CFD/CSD coupling (LC) is used to predict the blade elastic motions, structural moments, and pitch link loads at the specified maneuver revolution of the rotor and also to correlate these with the transient CSD-based predictions. A reasonable correlation is obtained. The LC results show more pronounced 5P (five per revolution) oscillations on the structural response than those of the CSD-based methods.

Vacuum and Hover Tests of a Dihedral–Anhedral Tip Composite Rotor

This paper presents test data from vacuum and hover tests of a 2.8-ft-radius dihedral–anhedral tip composite rotor. The paper describes the blades, their fabrication, properties, instrumentation, the test conditions, and the data acquired. The blades were Mach-scaled to a generic but representative modern rotor. Vacuum chamber tests measured rotating frequencies and strains. Hover tests measured performance, blade loads, pitch-link loads, and strains under steady and cyclic loading conditions. Three-dimensional finite element structural models were developed to ensure completeness and consistency of property definition. The three-dimensional analysis was also used for a preliminary assessment of the test data. The test data revealed that the dihedral–anhedral tip influences the torsional and higher frequencies of a rotor blade significantly. The oscillatory blade loads show patterns consistent with the vertical center-of-gravity offset introduced by the tip. The surface strains reveal interesting higher-harmonic patterns of loading particularly near the dihedral junction.

Performance and Loads of a Lift Offset Rotor, Part II: Prediction Validations with Measurements

The predictions of an upgraded UMARC comprehensive analysis are compared to experimental lift offset rotor results. The experiments cover a range of collective pitch angles (o) from 2° to 10°, advance ratios (μ) from 0.21 to 0.53, and lift offset from 0% to 20%. The experimental model rotors are from a system of coaxial hingeless rotors, with two blades each, and a first flap frequency of approximately 1.6/rev. The simulation is compared with isolated rotor performance and controls with lift offset, loads, and pitch link forces. Increasing efficiency with increasing lift offset, the impact of lift offset on different loads, and the dependence of pitch link loads on pitch bearing damping are identified in the experiment and correlated with the simulation.

High-Fidelity Aeromechanical Analysis of Coaxial Mars Helicopter

A high-fidelity coupled computational fluid dynamics and comprehensive analysis solver is developed for helicopter flight on Mars. The objectives are accurate prediction of flight loads and fundamental understanding of Martian aeromechanics. Performance, structural loads, pitch link loads, rotor wake, and interrotor blade separation are studied for hingeless and flap articulated coaxial rotors. In the absence of any validation data, pressure chamber tests are carried out at Mars-like conditions on a single rotor in hover. The data provided an elementary assessment of rotor behavior and a baseline for model validation. Predictions of the high-fidelity solver are also compared with a lower-fidelity lifting-line comprehensive analysis. It is generally concluded that lower-fidelity tools are not adequate for capturing the complex flow phenomena encountered on Mars. A motion and loads comparison of two fundamental hub types revealed that 1) an articulated hub has similar interrotor blade separation as hingeless on Mars, 2) a hingeless hub experiences only marginally higher (6–7%) flap bending moments, and 3) an articulated hub can lower steady pitch link loads (by an order of magnitude) at a marginal expense of oscillatory loads, 15% higher.

Effect of active camber morphing on rotor performance and control loads

A computational investigation of a helicopter main rotor with an active-camber morphing mechanism was conducted to identify the capability to simultaneously save rotor power and reduce pitch-link loads using as low as possible camber deflection magnitudes. Comprehensive rotor aeromechanics analysis with elastic blade modeling and a free vortex wake for the aerodynamics model was used to ensure computational efficiency. The investigation was based on a full-scale Bo 105 helicopter main rotor in level flight condition at μ=0.3 and CT/σ=0.089. In addition to a variation of the radial position and length of the active-camber section, the capabilities and system behavior were investigated in an extensive parametric study using practical actuation inputs. The same computational framework was used to obtain optimal control inputs that led to best performance in terms of power savings using two-per-rev (2P) individual blade control (IBC) via pitch-link inputs and 2P active-twist control. The relative potential of the three active mechanisms and the aerodynamic phenomenology was compared. All of the investigated active-rotor mechanisms contributed to rotor power savings via a more uniform distribution of thrust over the rotor disk. Both IBC and active twist yielded maximum performance improvement of about 1.8% over the baseline in terms of power reduction. With active camber, simultaneous rotor power reduction of 3.8% and a peak-to-peak pitch-link load reduction of 20% were obtained using a nonharmonic camber deflection deployment schedule with a half-peak-to-peak magnitude of 2∘. The global maximum of power savings observed with active camber was 4.4%, which resulted in an increase of peak-to-peak pitch-link loads by 35% and a required half-peak-to-peak camber deflection magnitude of 3.6∘.

Linearly variable chord-extension morphing for helicopter rotor blades

A new morphing concept called linearly variable chord extension was studied for its effectiveness in improving the efficiency of a helicopter rotor. Apart from chord extension itself, an additional feature which is deflection of the extended part of the chord resulting in an effective camber and additional twist to the airfoil, is also studied for its effect on rotor efficiency improvement. Trim analyses were carried out for various chord-extended rotors for hover as well as various forward flight velocities using DLR’s in-house comprehensive analysis code S4. Chord extension of up to 100% and chord-extension–deflection of up to 15° were considered. Results show that the linearly variable chord-extension concept is effective in reducing power requirement in both hover and forward flight. Deflection of the extended chord also helps reduce power requirement in hover, especially at higher blade loadings. However, the root torsional moments and hence, the pitch-link loads are seen to increase substantially for the morphed rotors.

Vibratory load predictions of a high-advance-ratio coaxial rotor system validated by wind tunnel tests

To investigate the aeromechanics of coaxial counter-rotating lift-offset rotor systems, a comprehensive analysis model of a laboratory-scale torque-balanced rotor designed for high-advance-ratio forward flight was developed. Measured blade and control system geometries and structural properties were input to the model. Lower-order aerodynamics modeling with a free-vortex wake method was used. While previous analytical studies on this coaxial rotor test rig have focused on performance and control requirements, in this current work, vibratory hub and pitch link loads, the influence of rotor–rotor phasing and the effects on blade deflections and tip clearance were investigated. The analysis was validated by wind-tunnel tests at advance ratios of 0.21–0.52 and for a lift offset varying from zero to 25%. Coaxial rotor performance, pitch link loads, unsteady thrust and rolling moments correlated well with the measurements. Pitching and rolling moment 2/rev and 4/rev harmonics correlated well for all lift offsets and advance ratios, whereas the vibratory torque was significantly overpredicted. The correct trends for varying lift offset and advance ratio were predicted in drag, side force, and thrust harmonics. Corresponding magnitudes were also predicted well, although an underprediction of the side force 4/rev harmonics was observed. Good correlation was found for the predicted blade tip clearance between the rotors over the entire range of lift offset and rotor–rotor phase angles, showing that advance ratio had little effect and judicial use of rotor phasing can increase the critical tip clearance.

Blade Passage Loads and Deformation of a Coaxial Rotor System in Hover

In a closely spaced, coaxial counter-rotating (CCR) rotor, the passage of upper and lower rotor blades over each other results in transient loads and blade deformation. This paper describes the measurement and numerical modeling of these transient loads and blade deformations in a single-bladed, 2-m-diam, rigid CCR rotor system in hover. Three-dimensional rotor blade deformation was measured using a time-resolved digital image correlation technique, simultaneously with the rotor thrust and pitch link loads. A numerical model of the rotor was developed using the comprehensive analysis CAMRAD II. This model was validated by measurements on an isolated single rotor and was used to explore the aerodynamic nature of the transient blade passage loads in the CCR rotor system. Rotating-frame natural frequencies were measured, and mode shapes of the first and second flap mode as well as the first lag mode of the rotor blade were extracted using the Complexity Pursuit algorithm. The 2-per-revolution tip displacement of the coaxial lower rotor due to the blade passages was found to be approximately 6% of the mean tip displacement. Vibratory thrust loads of the coaxial lower rotor were nearly 10% of the mean rotor thrust, whereas the vibratory pitch link loads were approximately 30% of the mean value. The numerical model predicted the vibratory hub and pitch link loads as well as out-of-plane deformations satisfactorily within the measurement uncertainties. With the validated analysis, additional numerical studies on the aerodynamic angles of attack, the inflow velocities from rotor–rotor interaction, and sectional lift distributions over the upper and lower coaxial rotor disks provided further insight into the sources of transient loads due to blade passage.

Compound Helicopter X3 in High-Speed Flight: Correlation of Simulation and Flight Test

This work presents the correlation of simulation results and flight-test data for a high-speed (V = 220 kt), high advance ratio (μ > 0.5) flight of the compound helicopter X3. The simulation tool chain consists of state-of-the-art coupling between the computational fluid dynamics (CFD) code FLOWer and the comprehensive analysis tool HOST. By applying a freeflight trim procedure, the experimental flight state is accurately represented in the simulation. The deviations of most trim controls is below 1°, and the maximum deviation is less than 1.4°. The analysis of the high-fidelity CFD results illustrates key features of the flow physics at this high advance ratio, such as wake interactions, reverse flow, and advancing side loading. The correlation of rotor dynamics data between simulation and flight test is favorable. Good accordance is demonstrated for flap bending moments, torsion moments, and pitch link loads. In contrast, the correlation is weaker for the chord bending moments for which it is shown that the interblade damper and drive train model mostly determine the structural loads.

Simulation of Dynamic Stall on an Elastic Rotor in High-Speed Turn Flight

A highly loaded, high-speed turn flight of Airbus Helicopters' Bluecopter demonstrator helicopter is simulated to investigate dynamic stall using a loose computational fluid dynamics/structural dynamics (CFD/CSD) coupling of the flow solver FLOWer and the rotorcraft comprehensive code CAMRAD II. The rotor aerodynamics is computed using a high-fidelity delayed detached-eddy simulation (DDES). A three-degree-of-freedom trim of an isolated rotor is performed, yielding main-rotor control angles that agree well with the flight-test measurements. The flow field in this flight condition is found to be highly unsteady and complex, featuring massively separated flow, blade–vortex interaction, multiple dynamic-stall events, and shock-induced separation. The computed pitch-link loads are compared to flight-test measurements. This shows that all CFD/CSD cases underpredict the amplitudes of the flight test and yield phase shifts. However, overall trends agree reasonably. Also, varying the computational setup reveals that the shear stress transport–DDES turbulence model performs better than Spalart–Allmaras–DDES, that the consideration of the rotor hub and fuselage improves the agreement with flight-test data, and that the elastic twist plays only a minor role in the dynamic-stall events.

Rotor performance enhancement via localized pitch control and its effects on hub vibrations and pitch link loads

This paper presents an approach to enhance rotor performance and reduce vibration levels at the rotor hub by introducing local blade pitch variations at specific azimuthal locations. Two different variants of these “localized pitch control” (LPC) functions are presented and their effects both on rotor performance and vibrational loads are studied using parametric variation and subsequent numerical optimization of the respective LPC signals. The results are compared with those obtained from a conventional approach using 2/rev higher harmonic control (HHC). Numerical calculations were performed using the DLR’s comprehensive rotor code S4 with a modified Beddoes inflow-model simulating an isolated hingeless rotor at an advance ratio of µ = 0.382. In those simulations power reductions of up to 7.61 and 3.36% could be reached using two sets of different model rotor blades. Both for power reduction and simultaneous reduction of power and 4/rev vibrational loads, LPC performed slightly better than conventional 2/rev HHC but also resulted in higher actuation demands which have to be taken into account for real-world applications of LPC.

Consideration of structural constraints in passive rotor blade design for improved performance

ABSTRACTThis design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral and the radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending and torsion) and the high-g manoeuvre pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1.0% in hover, 10.0% at μ = 0.30 and 17.0% at μ = 0.40, relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0.30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre. Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement.

Consideration of structural constraints in passive rotor blade design for improved performance

AbstractThis design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral, and radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending, and torsion), and the high g manoeuvring pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1·0% in hover, 10·0% at μ = 0·30, and 17·0% at μ = 0·40 relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0·30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement.

Coupled Aeroelastic Prediction of the Effects of Leading-Edge Slat on Rotor Performance

A coupled computational fluid dynamics and computational structural dynamics methodology is developed to analyze the effects of a leading-edge slat on rotor performance. The multielement intermesh connectivity was handled by implicit hole cutting overset implementation. The three-dimensional coupled computational fluid dynamics/computational structural dynamics model is extensively validated against a UH-60 flight-test condition: C9017. The solver is then used to successfully demonstrate the effectiveness of a leading-edge slat in mitigating (or eliminating) dynamic lift and moment stall on a modified UH-60A blade with a 40%-span slatted airfoil section. This results in a reduction of torsional structural loads (up to 73%) and pitch link loads (up to 62%) as compared to the baseline C9017 values. A dynamically moving slat strategy, actuating between slat positions S-1 and S-6 with 1 per revolution harmonic, is considered for stall mitigation. Further, it is shown that using slats with up to 10% higher thrust beyond the limit imposed by McHugh’s stall boundary (C9017) can be achieved. Stall mitigation due to the slat results in a reduction of torsional load up to 54% and a reduction of pitch link load up to 32% as compared to the baseline C9017 flight-test values, even for an increase in thrust of 10%.

Coupled Pitch Links for Multiharmonic Isolation Using Fluidic Circuits

Replacing stiff pitch links on rotorcraft with coupled fluidic devices has the potential to reduce the aerodynamic blade loads transmitted through the pitch links to the swashplate. Analytical models of two fluidic devices coupled with three different fluidic circuits are derived. These passive fluidlastic systems are tuned, by varying the fluid inertances and capacitances of each fluidic circuit, to reduce the transmitted pitch-link loads. The different circuit designs result in transmitted pitch-link loads reduction at up to three main rotor harmonics. The simulation results show loads reduction at the targeted out-of-phase and in-phase harmonics of up to 88% and 93%, respectively. Experimental validation of two of the fluidic circuits demonstrates loads reduction of up to 89% at the out-of-phase isolation frequencies and up to 81% at the in-phase isolation frequencies.

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.

Investigation of rotor control system loads

This paper concentrates on the aeroelasticity analysis of rotor blade and rotor control systems. A new multi-body dynamics model is established to predict both rotor pitch link loads and swashplate servo loads. Two helicopter rotors of UH-60A and SA349/2, both operating in two critical flight conditions, high-speed flight and high-thrust flight, are studied. The analysis shows good agreements with the flight test data and the calculation results using CAMRAD II. The mechanisms of rotor control loads are then analyzed in details based on the present predictions and the flight test data. In high-speed conditions, the pitch link loads are dominated by the integral of blade pitching moments, which are generated by cyclic pitch control. In high-thrust conditions, the positive pitching loads in the advancing side are caused by high collective pitch angle, and dynamic stall in the retreating side excites high-frequency responses. The swashplate servo loads are predominated by the rotor pitch link loads, and the inertia of the swashplate has significant effects on high-frequency harmonics of the servo loads.

Higher harmonic pitch link loads reduction using fluidlastic isolators

Fluidlastic isolators are proposed for higher harmonic pitch link loads reduction in helicopter rotors. An aeroelastic simulation of a soft in-plane rotor in high forward flight is conducted to investigate the dynamic characteristics of the coupled rotor and fluidlastic isolators. Using Hamilton’s principle, the system equations of motion are derived based on the generalized force formulation. The results indicate that the application of the fluidlastic isolator can reduce the 4/rev pitch link load by 98.9% in high forward flight with small variations in the other harmonic loads. The isolator has significant influence on the higher harmonic torsional rotation of the blade tip. Increasing the tuning port area ratio can significantly reduce the tuning mass with little variation of the isolation ability. Within 5% variation of the normal rotor speed, the 4/rev isolator can reduce more than 80% of the 4/rev pitch link load. The effects of isolator’s damping, forward speed, and thrust on the performance of the isolator are also studied. The ability to isolate other higher harmonic pitch link loads is also investigated.