![Cooper-Harper handling qualities rating scale [1].](https://www.researchgate.net/profile/Umberto-Saetti/publication/335135116/figure/fig2/AS:791271819993088@1565665311431/Cooper-Harper-handling-qualities-rating-scale-1_Q320.jpg)
Content uploaded by Umberto Saetti
Author content
All content in this area was uploaded by Umberto Saetti on Aug 13, 2019
Content may be subject to copyright.
A preview of the PDF is not available
Rotorcraft Flight Control Design with Alleviation of Unsteady Rotor Loads
Abstract and Figures
The objective of this research effort is to develop rotorcraft flight control laws that minimize unsteady rotor loads by acting solely through the primary flight controls (first harmonic swashplate control). As opposed to Higher-Harmonic Control, this strategy does not affect stationary (periodic trim) loads, and is therefore effective only in maneuvering flight. However, such system could be readily integrated with existing or future Automatic Flight Control Systems (AFCS). The study considers control designs for both conventional and compound configurations. Starting from a non-linear simulation model of the rotorcraft developed in FLIGHTLAB®, which includes sufficient fidelity to simulate rotor loads and vibrations, Linear Time-Periodic models (LTP) are derived via linearization. Next, the Harmonic Decomposition methodology is used to approximate the LTP systems with higher-order Linear Time-Invariant (LTI) systems. Reduced-order systems are subsequently obtained by using singular perturbation theory. By retaining the higher-harmonics of the rotor loads in the output, the reduced-order models are shown to accurately predicted the influence of the zeroth harmonics of the rigid-body and rotor flapping states on the higher-harmonics of the rotor loads. This way, previous limitations such as the reliance on non-physics-based models and curve fits to approximate rotor loads are lifted. Next, model following flight control laws are developed based on the reduced-order models. Parametric studies are performed to provide insights on how both the feed-forward and feedback paths of the model following control laws can be used to alleviate the rotor loads. Also, the impact of load alleviation on handling qualities is studied. It is shown that, for a standard helicopter configuration, load alleviation comes at the cost of a degradation in handling qualities. However, for the case of a compound rotorcraft, allocation of the control signal to the redundant control surfaces provides load alleviation without degradation in the handling qualities. The flight control laws are subsequently optimized using CONDUIT® to meet a comprehensive set of stability, handling qualities, and performance specifications for specific mission task elements while minimizing the unsteady rotor loads. Finally, since industry will not only rely on LTP systems obtained from simulation models, a novel methodology is developed to identify LTP systems from flight test data. The methodology is successfully applied to JUH-60A Black Hawk flight test data using CIFER®. The identified LTP systems capture the Nb/rev component of the rotorcraft dynamics. Further, it is shown how the higher-harmonics of the rotor states contribute to the overall rotorcraft dynamics for up to a 7%. On the other hand, the rigid-body states contribute to the overall rotorcraft dynamics almost entirely through their zeroth harmonic. Flight control design based on LTP systems identified from flight-test data could benefit the Future Vertical Lift (FVL) program. FVL is a plan to develop a new generation of military helicopters for the U.S. Army with increased capabilities in speed, range, and payload, and reduced maintenance and operational cost. Because these rotorcraft would operate at significantly higher speeds than the current helicopters, alleviation of the higher harmonic rotor loads and flight envelope protection are key elements to reduced maintenance cost. These rotorcraft are also likely to employ redundant control surfaces which, in connection with LTP-based flight control design, demonstrated outstanding effectiveness towards the alleviation of unsteady rotor loads.
Figures - uploaded by Umberto Saetti
Author content
All figure content in this area was uploaded by Umberto Saetti
Content may be subject to copyright.
... where x x x R 0 is the zeroth harmonic of the rotor state vector with the longitudinal and lateral flapping states removed, and 1c and 1s are the longitudinal and lateral flapping angles, respectively. The 10-state model will model the regressive flap mode and be accurate, at least for the UH-60 helicopter, up to about 10 rad/s, compared to approximately the 4 rad/s of the 8-state model retaining the rigid-body dynamics only [52,53]. It is worth mentioning that the regressive flap mode has a natural frequency that is significantly less than the slowest first harmonic component of the flight dynamics. ...
... Because in Eq. (24) the periodicity is limited to frequencies of one per forcing cycle, one would think it is sufficient to retain up to the first harmonic in the harmonic decomposition of state vector in Eq. (25). However, as pointed out in [53], harmonics up to one more that of the periodicity must be retained to capture all of the harmonic behaviour. Because of the complexity introduced by the extra harmonics in the calculations compared to the previous case, the high-order LTI model as well as its eigenvalues are computed numerically rather than analytically. ...
... where lat and lon are the main rotor lateral and longitudinal cyclic inputs, col is the main rotor collective input, and ped is the tail rotor collective, or pedal, input. The bare-airframe dynamics are augmented with a multi-loop nonlinear dynamic inversion (NDI) control law largely based on [51,53,57,[61][62][63][64][65][66]. The schematics of the (44b) ...
This article investigates vibrational stabilization effects in rotorcraft flight dynamics. This study is motivated by recent results in flapping-wing flight, which showed that the time-varying aerodynamic and inertial loads due to the insect wing periodic motion induce a vibrational stabilization mechanism in hover. The dynamics of flapping-wing flyers and rotary- wing vehicles are both described by time-periodic systems so vibrational stabilization mechanisms can also have an effect on stability characteristics of rotary-wing vehicles. The article extends the use of the harmonic decomposition method to vibrational stability analysis of rotorcraft. Two cases are considered: vibrational stability due blade imbalance at hover, and vibrational stability due to number-of-blades-per-rotor-revolution (Nb/rev) in high-speed forward flight. Results show that while vibrations induced by rotor blade imbalance do not stabilize the hovering dynamics of a helicopter, these vibrations can still have a significant effect on the hovering dynamics. Rotor blade imbalance results in a symmetric effect on the roll and pitch axes, in that it tends to decrease the frequency of the subsidence modes of the hovering cubic, while the unstable oscillatory modes tend to increase in frequency and decrease in damping (destabilizing effect). On the other hand, the yaw/ heave dynamics are relatively unaffected compared to the lateral and longitudinal axes. Moreover, Nb/rev rotor loads in forward flight are shown to reduce the damping of the coupled roll/pitch oscillation mode.
... The 10state model will model the regressive flap mode and be accurate, at least for the UH-60 helicopter, up to about 10 rad/s, compared to approximately the 4 rad/s of the 8-state model retaining the rigid-body dynamics only (Refs. 43,44). Both methods are used in this paper. ...
This paper investigates vibrational stabilization effects in rotorcraft flight dynamics. This study is motivated by the fact that eigenvalues of the rotorcraft flight dynamics identified from flight test often differ from those computed with physics-based simulations, and that some commonly observed mismatches may be ascribed to vibrational stability effects due to rotor blade imbalance or other periodic disturbance on the rotorcraft. Starting from a simple example involving an inverted pendulum, the paper demonstrates the use of the harmonic decomposition method for the study of vibrational stabilization effects. The concept is then extended to analyze the effect of blade imbalance on the flight dynamics of a helicopter. Additionally, vibrational stablization of a slung load in forward flight is investigated using small-amplitude and disturbances on an active cargo hook. Results show that while vibrations induced by rotor blade imbalance do not stabilize the hovering dynamics of a helicopter, these vibrations still have a significant effect on the hovering dynamics. Rotor blade imbalance results in a symmetric effect on the roll and pitch axes, in that it tends to decrease the frequency of the subsidence modes of the hovering cubic, while the unstable oscillatory modes tend to increase in frequency and decrease in damping. On the other hand, the yaw/heave dynamics are relatively unaffected compared to the lateral and longitudinal axes. Moreover, small-amplitude oscillations of an active cargo hook were shown to significantly decrease the amplitude of the limit cycle oscillation of a slung load in forward flight and to stabilize its dynamics. This constitutes a practical solution to the semi-active control of a suspended load.
This article presents the development and implementation of active noise abatement flight control laws utilizing redundant control allocation in rotorcraft. The study starts by finding a periodic equilibrium of coupled rotorcraft flight dynamics and acoustics using a modified harmonic balance trim solution method. Next, the rotorcraft nonlinear time-periodic dynamics are linearized about its periodic equilibrium and transformed into an equivalent higher-order linear time-invariant system in harmonic decomposition form. Composite aeroacoustic measures are formulated as outputs of this system. To enable faster runtime and tractable control design, a reduced 8-state model representative of the coupled rigid-body dynamics and acoustics is obtained through residualization. An Explicit Model Following (EMF) control law with pseudo-inverse allocation to redundant control surfaces -- an active horizontal stabilizer in this study -- is synthesized using the 8-state model. The article further assesses the impact of noise alleviation control on handling qualities of conventional and compound rotorcraft in forward flight. The results demonstrate that redundant control allocation is an effective method for active reduction of unsteady rotor noise, with increasing effectiveness during aggressive maneuvers. Future-generation rotorcraft, particularly Future Vertical Lift (FVL) configurations featuring high control redundancy and aggressive maneuvering capabilities, can significantly benefit from noise abatement through redundant control allocation. Furthermore, the article highlights that while implementing noise abatement control in conventional helicopters can degrade handling qualities, the use of redundant control allocation in compound configurations offers noise reduction without compromising handling qualities.
The present study considers two notional rotorcraft models: a conventional utility helicopter, representative of an H-60, and a wing-only compound utility rotorcraft, representative of an H-60 with with a wing similar to the X-49A wing. An Explicit Model Following (EMF) control scheme is designed to achieve stability and desired Rate Command / Attitude Hold (RCAH) response around the roll, pitch and yaw axes, while alleviating vibratory loads through both feed-forward and feedback compensation. The harmonic decomposition methodology is extended to enable optimization of primary flight control laws that mitigate vibratory loads. Specifically, Linear Time Periodic (LTP) systems representative of the periodic rotorcraft dynamics are approximated by Linear Time Invariant (LTI) models, which are then reduced and used in LQR design to constrain the harmonics of the vibratory loads. The LQR gains are incorporated in the EMF scheme for feedback compensation. One innovative approach is the addition of rotor state feedback to standard rigid body state feedback. A Pseudo Inverse (PI) strategy is incorporated into the EMF scheme for redundant control allocation. Finally, simulation results with and without load alleviation are compared and the impact of PI feed-forward and rotor state feedback compensation on handling qualities is assessed in terms of ADS-33E specifications.
Linearized time-periodic models are extracted from a high-fidelity comprehensive nonlinear helicopter rotor model at a trim flight condition. Subsequently, a Fourier expansion-based model reduction method is used to obtain linearized timeinvariantmodels fromthe time-periodic models. These linearized models predict the effects of on-blade control systems such as active flaps on the rotor hub loads and hence are an important component required for vibration control and flight control interaction studies. The linearized models are verified against the nonlinear model at low-speed descending and cruise flight conditions. Inclusion of aerodynamic states corresponding to the unsteady aerodynamic model reduces the error in the peak-to-peak hubload amplitude predictions from over 80% to less than 10%. The time-periodic and time-invariant models are in excellent agreement. The linear time-invariant (LTI) models are also verified for closed-loop vibration reduction performance fidelity using active flaps activated through the higher harmonic controller. The flap deflection histories and the vibratory loads predicted using the LTI model and nonlinear model agree very well when the flap deflection is limited to 2?. The errors between the two models increase with the flap deflection amplitude.
Higher harmonic control (HHC) has been shown to be capable of reducing vibration and noise in rotorcraft. The majority of the previous work has concentrated on the HHC itself, and not the interaction between the automatic flight control system (AFCS) and HHC. Limited studies have been done on HHC/AFCS interaction using a linear time invariant (LTI) model and offline inputs including the use of dynamic crossfeeds to improve HHC performance while maneuvering. Most recently, results on HHC/AFCS interaction using nonlinear piloted simulations have been obtained. It has been observed that the nominal T-matrix controller approach, while effective in steady state, does not reduce vibration during a maneuver. Increasing the feedback gain can slightly improve RMS and settling time but magnifies peak response. The aim of the current work is to optimize HHC controller performance in a realistic piloted maneuver, and validate the design on a nonlinear simulation model. Achieving this objective required the following accomplishments: 1) Extraction, reduction, and validation of a linear time invariant model from a nonlinear model. 2) Baseline higher harmonic controller design and validation based on the linear model applied to the nonlinear model. 3) Optimization of the higher harmonic controller design to improve performance during a maneuver while maintaining stability. 4) Fitting ideal crossfeeds with a low-order approximation. 5) Implementation of crossfeeds to further improve vibration controller performance during a maneuver.
Reducing the theoretical methods of flight control to design practice, Practical Methods for Aircraft and Rotorcraft Flight Control Design: An Optimization-Based Approach compiles the authors’ extensive experience and lessons learned into a single comprehensive resource for both academics and working flight control engineers.
Several methods for analysis of linear time periodic (LTP) systems have successfully been demonstrated using harmonic decompositions. Onemethod recently examined is to create a linear time invariant (LTI) model approximation by expansion of the LTP system states into various harmonic state representations, and formulating corresponding LTI models. Although this method has shown success, it relies on a second-order formulation of the original LTP system. This second-order formulation can prove problematic for degrees of freedom not explicitly represented in second-order form. Specifically, difficulties arise when performing the harmonic decomposition of body and inflow states as well as interpretation of LTI velocities. Instead this paper presents a more generalized LTI formulation using a first-order formulation for harmonic decomposition. The new first-order approach is evaluated for a UH-60A rotorcraft model and is used to show the significance of particular harmonic terms, specifically that the coupling of harmonic components of body and inflow states with the rotor states makes a significant contribution to LTI model fidelity in the prediction of vibratory hub loads.
Formulation of linear time invariant (LTI) models of a nonlinear system about a periodic equilibrium using the harmonic domain representation of LTI model states is well established in the literature. This paper presents a computationally efficient scheme for implementation of a previously developed method for extraction of linear time invariant (LTI) models from a helicopter nonlinear model in forward flight. The fidelity of the extracted LTI models is evaluated using response comparisons between the extracted LTI models and the nonlinear model in both time and frequency domains. For time domain evaluations, individual blade control (IBC) inputs that have been tried in the literature for vibration and noise control studies are used. For frequency domain evaluations, frequency sweep inputs are used to obtain frequency responses of fixed system hub loads to a single blade IBC input. The evaluation results demonstrate the fidelity of the extracted LTI models, and thus, establish the validity of the LTI model extraction process for its use in integrated flight and rotor control studies.
In a traditional single main rotor helicopter, only 4 controls arc available - lateral cyclic, longitudinal cyclic, main rotor collective pitch and tail rotor collective pitch. As such, redundancy is not available as all four controls arc needed to maneuver the aircraft in the four control axes - pitch, roll, yaw and heave. In contrast, high speed rotorcraft configurations (such as tilt-rotors and compound helicopters) have multiple redundant control effectors. The strategy for applying these control effectors, and in which proportion, is called 'control allocation1. This paper provides an evaluation and comparison of different control allocation methods, applied to a medium lift tilt rotor. The control allocation methods are evaluated based on a comprehensive set of time and frequency domain metrics, as well as robustness criteria. In addition to evaluation of standard control allocation methods, several control allocation methods arc further developed and evaluated for reducing flap motion of the rotors. A piloted fixed base simulation is performed, and pilot comments and data are presented. The paper concludes with discussion of the results and trade-offs of the most effective control allocation techniques.