Rotorcraft Comprehensive Analysis System Research Articles

Correlation of High-Speed Tiltrotor Stability Predictions with Test Data and Parametric Study

High-speed stability of tiltrotor was studied. The University of Maryland’s Maryland Tiltrotor Rig (MTR) was chosen for the analysis due to availability of properties and test data, and its interesting high-stability behavior observed in the Glenn L. Martin wind tunnel in August 2022. A Rotorcraft Comprehensive Analysis System (RCAS) model of the MTR gimbaled hub was built in addition to the University of Maryland Advanced Rotorcraft Code-II (UMARC-II) model from previous work. The objective is threefold: i) validate RCAS tiltrotor stability predictions, ii) shed light on the high-stability behavior of the MTR, and iii) find ways to lower the instability speed of the MTR for future wind tunnel tests. Trim collective for freewheeling and stability predictions were compared with wind tunnel test data up to 200 knots. RCAS and UMARC-II predictions showed good agreement with each other and the test data. Predictions show that MTR is stable up to 215 knots (490-knots full-scale flight) although the wing is only 18% thick (current technology is 23%). A parametric study was carried out. The impact of wing stiffness, pitch-flap coupling (δ3 angle), lag stiffness, blade chord, number of blades, pylon mass, pylon center of gravity (c.g.), pylon location, and rotor speed was studied. MTR’s pylon c.g. is unconventionally behind the wing elastic axis. It was found that this significantly improved stability. This behavior is not specific to MTR; full-scale aircraft stability can also be improved by moving the pylon c.g. backward if wing beam is the least stable mode. A combination of forward pylon c.g., reduced rotor speed, and increased blade chord reduced the instability speed by more than 55 knots to near 160 knots, helping researchers obtain high-quality test data in the upcoming Glenn L. Martin wind tunnel tests.

Experimental and Computational Investigation of Rotor Noise in Hover

The noise magnitude and directivity of a four-bladed 1.108-m-radius rotor in hover were measured at a tip Mach number of 0.42, representative of an electric vertical takeoff and landing vehicle rotor. The tonal noise was filtered from the acoustic spectra, allowing for analysis of the separate contributions of tonal and broadband noise. Accompanying computations of the same rotor and operating conditions were performed using the rotorcraft comprehensive analysis system (RCAS) with a viscous vortex particle method (VVPM) for aerodynamic calculations and PSU-WOPWOP with the semi-empirical Brooks, Pope, and Marcolini (BPM) broadband model for predicting acoustics. Measured broadband and tonal noise contributed similar magnitudes to the overall sound pressure level, with broadband always several decibels greater. This is in contrast to conventional helicopter rotors, which are dominated by tonal noise. Simulated trends of tonal noise matched the measurements, though the higher harmonic components of the acoustic spectra were underpredicted by up to 30 dB. The broadband noise was underpredicted by 10 dB at moderate rotor collectives, as the measured noise contained noise from laminar and turbulent boundary layers, while the model predicted predominantly laminar boundary layer–vortex shedding noise. Comparing the time histories of the measured and simulated rotor thrust revealed that the RCAS-VVPM method does not capture the amplitude of the 4/rev component of the thrust and lacks the higher frequency content present in the measured thrust, indicating a possible source of discrepancy in the tonal acoustic predictions. The results of this study show that methods to predict rotor tonal and broadband noise without the use of lengthy computational fluid dynamics computations must be further developed.

Performance and Acoustics of a Stacked Rotor with Differential Collective Pitch

Performance and acoustics of a 1.108-m radius stacked rotor in hover with an axial spacing between the upper and lower rotors of 0.73 chords were measured. The differential collective pitch of the rotor system, de ned as half the difference between the upper rotor pitch and lower rotor pitch was varied along with the azimuthal spacing between the upper and lower blade sets. Coupled comprehensive analysis and acoustics simulations were conducted using the Rotorcraft Comprehensive Analysis System, the acoustic solver PSU-WOPWOP, and the Brooks, Pope, and Marcolini method implemented in ANOPP2. Increased sensitivity to azimuthal spacing was measured for large, negative differential collectives, resulting in a 76.5% change in total rotor thrust and a 46% change in total rotor power over 11.25 ° change in azimuthal spacing. The trends in individual and total system thrust were predicted with good accuracy, but the total power was overpredicted. Overall sound pressure level decreased by 3.5 dB at – 2° differential collective and 90° azimuthal spacing due to reduced broadband levels, indicating a possible quiet mode of operation for this type of rotor. By assessing the highest performance and lowest noise, an optimum rotor con guration was observed at an azimuthal spacing of – 45° and differential collective of 2°. Acoustic predictions captured tonal and broadband noise trends in most conditions. Conditions where broadband noise was underpredicted could be a result of vortex interactions not modeled by the broadband simulations.

Variations in Thrust Sharing for Torque-balanced Lift-offset Coaxial Rotors

This paper presents numerical calculations, on coaxial rotor systems, that show the variations of rotor thrust share with lift offset, system thrust, and advance ratio. The calculations are based on a coupled analysis between the U. S. Army's Rotorcraft Comprehensive Analysis System (RCAS) and a viscous vortex particle method. The level of agreement between calculations and experimental data is good in hover and mixed in edgewise flight. In hover and zero lift offset–a condition in which the upper rotor typically produces more thrust than the lower rotor–changing system thrust does not significantly alter the rotor thrust share. For advance ratios up to approximately 0.25, independently increasing lift offset increases the lower rotor thrust; this is an aeroelastic phenomenon that is eliminated with rigid blades. Independently increasing advance ratio leads to three distinct regions with different thrust sharing behaviors; these behaviors are governed by changes to the longitudinal skew of the upper rotor wake and its proximity to the lower rotor.

Unified Modeling of the TiltRotor Aeroelastic Stability Testbed for Proprotor Whirl Flutter

The present paper studies whirl flutter stability of the TiltRotor Aeroelastic Stability Testbed (TRAST) using the comprehensive analysis code Rotorcraft Comprehensive Analysis System (RCAS). The TRAST wing and pylon are systematically developed in RCAS as beams based on a NASTRAN model starting with a spar and more sophisticated models are incrementally built up by adding components one at a time. Natural frequencies are compared with those of the NASTRAN model at each step. The elastic wing–pylon beam model is then combined with an elastic gimbaled rotor model, and whirl flutter analysis is carried out. Comparisons are made between uniform inflow and dynamic inflow for aeroelastic stability. Increased flutter speed with dynamic inflow is observed against uniform inflow. Wing and pylon aerodynamics are also added to the model, and their effects on aeroelastic stability are examined. A few key design parameters [pitch spring stiffness, rotor rotational speed, and pitch–flap coupling ()] are varied, and their effects on whirl flutter speed are examined. In particular for TRAST, the pitch spring stiffness and pitch–flap coupling have an important influence on whirl flutter speed.

Experimental and computational investigation of stacked rotor performance in hover

Experiments were performed on a 1.108 m diameter coaxial, co-rotating rotor, also called a stacked rotor, in hover. The performance of the stacked rotor was measured at various azimuthal and axial spacings. The performance of the rotor was also analyzed using the Rotorcraft Comprehensive Analysis System (RCAS) employing the viscous vortex particle method (VVPM) as well as computational fluid dynamics (CFD). Both the experimental results and the simulations showed that as the azimuthal spacing changed from negative angles (lower blades leading) to positive angles (upper blades leading), the thrust on the upper rotor increased by ∼50% and the thrust on the lower rotor decreased by a similar amount. The total thrust decreases by as much as 10% as the index angle reached zero. These changes in individual and total rotor thrust are explained by the interference velocity due to the bound circulation of the lifting blades. This property, combined with advances in electric motor control technology, can be useful as a novel method of aircraft control by in-flight adjustments of the azimuthal spacing. A 17.1% change in total thrust was observed with an azimuthal spacing change of 22.5∘ at constant collective. While RCAS and CFD predicted the rate of change with azimuthal angle well, the total change in thrust was underpredicted (11.6% change in thrust predicted by CFD and 6.3% predicted by VVPM). Configurations were determined which improved rotor efficiency, measured by CT/CP, by 11.6% compared to a single, four-bladed rotor. VVPM and CFD simulations were used to understand the physical mechanisms responsible for changes in thrust and efficiency including the blades proximity to tip vorticies.

On the Influence of Inflow Model Selection for Time-Domain Tiltrotor Aeroelastic Analysis

The methods of using the viscous vortex particle method, dynamic inflow, and uniform inflow to conduct whirl-flutter stability analysis are evaluated on a four-bladed, soft-inplane tiltrotor model using the Rotorcraft Comprehensive Analysis System. For the first time, coupled transient simulations between comprehensive analysis and a vortex particle method inflow model are used to predict whirl-flutter stability. Resolution studies are performed for both spatial and temporal resolution in the transient solution. Stability in transient analysis is noted to be influenced by both. As the particle resolution is refined, a reduction in simulation time-step size must also be performed. An azimuthal time step size of 0.3 deg is used to consider a range of particle resolutions to understand the influence on whirl-flutter stability predictions. Comparisons are made between uniform inflow, dynamic inflow, and the vortex particle method with respect to prediction capabilities when compared to wing beam-bending frequency and damping experimental data. Challenges in assessing the most accurate inflow model are noted due to uncertainty in experimental data; however, a consistent trend of increasing damping with additional levels of fidelity in the inflow model is observed. Excellent correlation is observed between the dynamic inflow predictions and the vortex particle method predictions in which the wing is not part of the inflow model, indicating that the dynamic inflow model is adequate for capturing damping due to the induced velocity on the rotor disk. Additional damping is noted in the full vortex particle method model, with the wing included, which is attributed to either an interactional aerodynamic effect between the rotor and the wing or a more accurate representation of the unsteady loading on the wing due to induced velocities.

Evaluation of Finite-State Dynamic Inflow for Rotors

In an effort to evaluate the accuracy of the finite-state dynamic inflow of Peters and He (“Finite State Induced Flow Models Part II: Three-Dimensional Rotor Disk,” Journal of Aircraft, Vol. 32, No. 2, 1995, pp. 323–333) for use in rotor design, comparisons between dynamic inflow-based calculations and measured test data are presented for a wide variety of single rotor configurations and operating conditions. The quantities compared include rotor performance parameters, blade flap bending moments, blade airloads, and blade pitch angles. The dynamic inflow calculations are performed using the U.S. Army’s Rotorcraft Comprehensive Analysis System. Rotor performance calculations are mostly accurate prior to stall, except for rotor torque calculations from an individual blade pitch control phase sweep. Blade airloads calculations do exhibit the correct dominant features, including those from blade tip vortices, but the calculated amplitudes of the wake-induced loading are significantly less than the test data; this leads to significantly lower amplitudes in the flap bending moment response and is problematic for calculating half peak-to-peak values in conditions with significant blade–wake interactions. Blade collective and cyclic pitch angles are reasonably accurate in edgewise flight.

Tiltrotor Conversion Maneuver Analysis with RCAS

Aeromechanics analysis is performed using the comprehensive analysis code RCAS (Rotorcraft Comprehensive Analysis System) to study the transient conversion maneuver of a tiltrotor. The analytical model is based on the XV-15 research tiltrotor aircraft in size and dynamic characteristics. A generic (not representative of XV-15) tiltrotor control system is developed to simulate a conversion maneuver. Hover and cruise performance of the present XV-15 analytical model is validated against available test data. The conversion calculation begins with a trim analysis at hover, which is followed by the conversion maneuver. During the maneuver analysis, the pilot control model is activated to fly the aircraft following a desired airspeed profile and minimum altitude change. Time histories of vehicle dynamics, rotor controls, rotor flapping, rotor performance, and blade structural loads are investigated for various transient conversion maneuvers. The aircraft acceleration during the transient maneuver has a significant influence on the rotor performance and loads.

Hover Performance of Corotating and Counterrotating Coaxial Rotors

Hover testing was performed on a 2-m-diameter coaxial corotating (stacked) rotor system to investigate the effect of azimuthal spacing, or index angle, on performance. Individual rotor thrust and power measurements were performed as a function of the index angles. It was found that the contributions of individual rotors to the total stacked rotor performance were remarkably dependent on the index angle. For an index angle of 10°, that is, when the lower rotor trails the upper rotor by 10°, a nearly 3.6% increase in total system power loading was observed compared to that for an index angle of –10°. Numerical calculations using the U.S. Army's Rotorcraft Comprehensive Analysis System (RCAS) were used to better understand this behavior in terms of self-induced and interference power constituents. For the coaxial, counterrotating (CCR) rotor, the upper rotor performance was close to that of an isolated two-bladed rotor. The lower rotor operating in the wake of the upper rotor showed a relatively poor performance, requiring 70% more induced power than the upper rotor. The improvement in performance for a CCR rotor results from a larger thrust share and better performance of the upper rotor as compared to the lower rotor.

Rotorcraft Comprehensive Analysis Calculations of a Coaxial Rotor with Lift Offset

This paper presents comparisons of rotor performance, blade pitch, rotor hub loads, and blade clearance between two sets of rotorcraft comprehensive analysis calculations and experimentally measured test data for a coaxial rotor system in forward flight conditions. The calculations are from the U.S. Army’s Rotorcraft Comprehensive Analysis System (RCAS) with one set calculated using a prescribed vortex wake model and another set calculated by coupling RCAS to a viscous vortex particle method. For rotor performance and blade pitch, the calculations and test data are presented as functions of lift offset for four different advance ratios and for two different test upper rotor collective pitch settings. Rotor hub loads and blade clearance are compared on a time-varying basis for test points that form a lift offset sweep, an advance ratio sweep, and a rotor-to-rotor phase sweep. The main conclusion is that both sets of calculations generally reflect the same trends as the test data, which provides confidence in the ability of RCAS at calculating the aeromechanics behaviors of lift-offset coaxial rotors.

Using Control Redundancy for Power and Vibration Reduction on a Coaxial Rotor Helicopter at High Speeds

This study uses the Rotorcraft Comprehensive Analysis System (RCAS) to examine the effect of control redundancy on power and vibratory hub loads of a lift-offset coaxial rotor helicopter operating at 230 kt cruise speed. An aircraft nose-up pitch attitude of 3° resulted in very low main rotor power (less than 10% of the total power), with the majority of the power consumption attributed to an efficient axial propeller. At this 3° pitch attitude, the rotor speed and differential lateral pitch, which are redundant controls, were parametrically varied, and low power (LP) and low vibration (LV) states identified. The LP state (80% Nr and 3° differential lateral) required 3.5% lower power than the LV state (90% Nr and 0° differential lateral), but the latter had substantially lower 3/rev vibratory hub loads. The lower power in the LP state is primarily due to reduced main rotor power on account of smaller drag on the advancing blade tip at lower rotor speeds. The rotor drag is comparable for the two states, but the LV state has larger drag contributions from the advancing side, whereas the LP state has larger contributions from the reverse flow region (accounting for 14% of the total rotor drag) due to higher pitch on the retreating side and larger reverse flow velocities. Even so, the rotor drag accounts for under 30% of the total propulsor thrust requirement, with the fuselage (and hub) drag being the dominant component. Rotor L/De values for the LP and LV states were 12.3 and 11.3, respectively.

Coupled Computational Fluid Dynamics and Comprehensive Analysis Calculations of a Gimballed Tiltrotor

In an effort to assess the suitability of the HPCMP CREATE-AV™ Helios computational tool for calculating tiltrotor aeromechanics characteristics, comparisons of rotor performance, blade pitch, and blade loads between calculations and test data are presented for the Tilt Rotor Aeroacoustic Model rotor in helicopter mode at an advance ratio of 0.15. The Helios calculations are performed by coupling the Fully Unstructured Navier–Stokes Three-Dimensional solver with the Rotorcraft Comprehensive Analysis System along with modeling the gimbal motion of the rotor. For rotor performance and blade pitch, the comparisons are with test points that form shaft angle sweeps for two different thrust levels. For blade loads, the comparisons are analyzed in detail for two test points with 3 deg forward shaft tilt and different thrust levels. The level of agreement between the calculations and the test data is very good for rotor performance, blade pitch, and blade airloads. The level of agreement is mixed for blade structural loads.

Comparison of Comprehensive Analyses Predicting Whirl Flutter Stability of the Wing and Rotor Aeroelastic Test System

Whirl flutter stability is a critical limitation for tiltrotor aircraft. This paper investigates whirl flutter predictions for the Wing and Rotor Aeroelastic Test System (WRATS) using comprehensive analysis, focusing on the comparison of the whirl flutter stability predictions between Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD) II and the Rotorcraft Comprehensive Analysis System (RCAS). The analytical models were created using a modular approach to systematically validate the modeling process of the WRATS tiltrotor. Comparison of nonrotating frequencies for blade, flexbeam, flexbeam and cuff, and blade with flexbeam and cuff shows excellent agreement between CAMRAD II and RCAS. The assembled model is then used to predict the whirl flutter stability boundary for various configurations with varying levels of fidelity. Results show near exact agreement between the two analyses for a rigid rotor and linear aerodynamics, and good to fair agreement when an elastic rotor is used. Predicted wing beam mode frequencies and damping values are also compared against the wind tunnel test data, and the frequencies are shown to be reasonably well-predicted. However, damping values, and thus stability boundaries, are not accurately predicted.

Assessing Calculated Blade Loads of the Tilt Rotor Aeroacoustic Model

In an effort to assess the suitability of the U.S. Army’s Rotorcraft Comprehensive Analysis System for calculating tiltrotor aeromechanics characteristics, comparisons of rotor performance and blade loads between calculations and test data are presented for the Tilt Rotor Aeroacoustic Model rotor in helicopter mode at an advance ratio of 0.15. For rotor performance, the comparisons are with test points that form shaft angle sweeps for two different thrust levels. These calculations include those from a free vortex wake model and those from coupling the comprehensive analysis to a viscous vortex particle method. The emphasis of this work is on the comparisons of blade loads, which are analyzed in detail for two test points with a 3 deg forward shaft tilt and different thrust levels. Blade airloads calculations include the two aforementioned sets of calculations, whereas blade structural loads calculations also include those from prescribing experimentally measured airloads. The level of agreement between calculations with the free vortex wake model and test data is generally comparable with those of helicopter rotors at similar conditions. The level of agreement improves with the viscous vortex particle method. Meanwhile, some deficiencies in structural loads calculations with prescribing measured airloads are noted.

UH-60A Rotor Performance and Loads Correlation at High Advance Ratios

The UH-60A slowed rotor wind-tunnel test data is a unique and high-quality database for the validation of rotorcraft analysis tools at high advance ratio flight conditions. The U.S. Army’s rotorcraft comprehensive analysis system, known as RCAS, incorporates several analysis options, which are evaluated for their suitability at high advance ratios. RCAS predictions of performance, controls, airloads, and structural loads are compared against high advance ratio test data. RCAS modeling options of dynamic inflow, prescribed vortex wake, yawed flow corrections, and radial drag are evaluated. The strengths and weaknesses of the analysis options are highlighted. Rotor thrust and drag predictions are in good agreement with test data, with the addition of radial drag and yawed flow models. Dynamic inflow and prescribed vortex wake performance and loads predictions provide comparable correlation; however, dynamic inflow provides the most robust convergence without numerical issues in the reverse flow region. Yawed flow corrections improve normal force correlation at inboard stations but do not improve the structural load correlation at midspan.

Physics-Based Hydraulic Damper Model for Rotor Structural Loads

A physics-based, fully nonlinear hydraulic damper model was developed within the framework of the Rotorcraft Comprehensive Analysis System in order to facilitate integrated damper and rotorcraft analysis and design. The hydraulic damper governing equations of motion were formulated using flow continuity equations for each hydraulic device within the damper. The developed damper model was compared with bench test data of a hydraulic damper under harmonic oscillation. It was also evaluated as an integrated rotor-damper solution against test data from the NASA–Army UH-60A Airloads Program. For these evaluations, the accuracy of blade load predictions using the newly developed damper model was compared against conventional linear and nonlinear (table lookup) damper models. From these comparisons, the physics-based damper model is able to capture the hysteresis loops of the isolated damper under harmonic oscillation and provide physical insight by modeling the damper’s component parts. When integrated into Rotorcraft Comprehensive Analysis System, the physics-based model was found to be consistent with the legacy nonlinear model for calculating blade loads for the high-speed and high-thrust conditions evaluated.

Flutter of Variations on a 5 MW Swept Wind Turbine Blade1

Previous research has shown that as wind turbine rotor designs increase in diameter they approach flutter instabilities. Recent blade designs incorporate blade sweep to lower fatigue loads and increase rotor diameter for more energy production; however, the flutter margin may be reduced with sweep. In this work, flutter speeds are predicted and compared for turbines with straight and swept blades. For this study, a wind turbine analysis tool specifically developed for blade sweep (CurveFAST) was used, in addition to a tool developed for helicopter analysis: rotorcraft comprehensive analysis system (RCAS). The results showed that CurveFAST generally predicted a higher flutter speed and RCAS predicted a lower flutter speed compared to previous research. The difference may be attributed to the unsteady aerodynamics modeling. The results for an extended radius and swept 5 MW turbine showed that the flutter speed was most sensitive to the position of the blade center of mass axis. Using the selected design parameters, it was not possible to design a swept 5 MW blade with adequate flutter margin that would produce 5% more energy with the same loads as a baseline straight rotor. Further study of swept blades is required to determine if there is a size limit above which flutter cannot be avoided.

New strategy for designing composite rotor blades with active flaps

This article presents the development of a mixed-variable optimization framework for the design of composite rotor blades with active flaps in order to maximize the control authority for vibration reduction. For the studies presented in this article, the amplitude of dynamic twist at the blade tip at 4/rev flap actuation frequency is used as the objective function, and it is shown that there is a direct correlation between the amplitude of dynamic twist and the vibration reduction authority of active flaps. The optimization framework developed includes Intelligent Cross-section Generator as the cross section and mesh generator, University of Michigan/Variational Asymptotic Beam Sectional Analysis for the cross-sectional analysis, and Rotorcraft Comprehensive Analysis System for the aeroelastic analysis of active rotor blades. The optimization problem is solved using a surrogate-based approach in combination with the efficient global optimization algorithm. The optimum results with both mixed and continuous design variables are obtained for three different spanwise locations of active flap.

Assessment of Beam and Shell Elements for Modeling Rotorcraft Blades

A geometrically exact shell element is developed within the finite-element, multibody dynamics-based Rotorcraft Comprehensive Analysis System. The shell element accommodates transverse shear deformation as well as arbitrarily large displacements and rotations. The shell element is developed using an approach that allows for compatibility with other structural elements in the Rotorcraft Comprehensive Analysis System. It is validated by comparing its predictions with benchmark problems. The two-dimensional shell and one-dimensional beam finite-element analyses are compared for three typical blade configurations of varying slenderness ratio: a swept-tip blade, a blade with discontinuous chordwise elastic axis and center-of-gravity locations, and a blade with a flex beam. The purpose is to quantify the differences between two-dimensional-shell one-dimensional-beam finite elements for modeling rotor blades. There is good agreement between the one- and two-dimensional analyses in predicting the natural frequencies of straight blades with a slenderness ratio of greater than 12.5. However, there are up to 16.5% differences in frequency predictions for a slenderness ratio of 17.5 or less and a tip sweep angle of 30 deg or higher. The differences increase with decreasing slenderness ratio and increasing tip sweep angle. For a blade with a discontinuous center-of-gravity offset, frequencies predicted using the two- and one-dimensional analyses show up to 12.5% difference and increase with the center-of-gravity offset. For a blade with discontinuous elastic axis offset, the frequencies predicted using the two analyses have good agreement. The differences are observed when both approaches calculate the structural loads (bending moments) and deformations of a blade with an elastic axis offset.