Single Main Rotor Research Articles

Unsteady Loads of a Scaled Rotor on and Near the Landing Deck of the NATO Generic Destroyer with Concurrent Stereo PIV and Surface Pressure Measurements

Rotorcraft shipboard operations are risky and demand high piloting skills. To gain further understanding and familiarize pilots with these complicated scenarios, various methods of computational simulations have been used in the past decades. Yet, some simulation methods may not capture a realistic response of the rotorcraft due to simplified modeling of the interactional aerodynamics in order to achieve real-time capability for pilot training. Thus, improvements to these simulations are required, and experimental data that unveil the interactional aerodynamics and dynamics between the rotor and ship are needed for computational validation efforts. In this study, an extensive experimental investigation of the ship???rotor dynamic interface problem was conducted to gain a general understanding of the interactional aerodynamics between a 1:100 wind-tunnel-scale NATO Generic Destroyer and a representative single main rotor, operating both stationary and dynamically moving in space in the vicinity of the landing deck. Data obtained through simultaneous measurements of rotor hub loads, ship deck surface pressures, and stereoscopic particle image velocimetry flow fields gave valuable insight into the highly coupled aerodynamic phenomena. Results showed that the rotor hub loads exhibited a high dependency on both the wind direction and also the position of the rotor relative to the landing deck. A bifurcation in the regions of high unsteady thrust was observed under certain wind conditions due to the ship airwake, impacting different portions of the rotor disk. The inflow angles across the rotor disk were estimated and assessed through flow field measurement, revealing how the ship airwake altered the rotor inflow and wake structure that contributed to the unique change in rotor loads under various hovering conditions.

Performance Improvement of Winged Compound Helicopter Due to Differential Flap Deflections

A lift-offset system for a single-rotor-type compound helicopter is proposed to improve the aerodynamic performance in high-speed flight. The proposed system utilizes the differential flap deflections on the fixed wings to produce a rolling moment, which is counteracted by the single main rotor, causing the rotor to operate in a lift-offset state. The performance of the proposed system is evaluated through numerical simulations. At first, the effect of lift offset with regard to lift-share ratio on the rotor performance is investigated. Then, the impact of lift offset due to the differential flaps on the overall effective lift-to-drag ratio is studied. The results show that the lift offset significantly improves the rotor performance and the overall effective lift-to-drag ratios, especially at larger rotor lift-share ratios. The overall effective lift-to-drag ratio increases by 10% due to the differential flaps compared to the zero-flap deflections. It is concluded that the lift offset due to the differential flaps achieves more efficient cruising flight for a single-rotor-type compound helicopter.

Effects of Ambient Conditions on Helicopter Harmonic Noise Radiation: Theory and Experiment

The effects of ambient atmospheric conditions, air temperature, and density on rotor harmonic noise radiation are characterized using theoretical models and experimental measurements of helicopter noise collected at three different test sites at elevations ranging from sea level to 7000 ftabove sea level. Significant changes in the thickness, loading, and blade–vortex interaction noise levels and radiation directions are observed across the different test sites for an AS350 helicopter ying at the same indicated airspeed and gross weight. However, the radiated noise is shown to scale with ambient pressure when the flight condition of the helicopter is defined in nondimensional terms. Although the effective tip Mach number is identi ed as the primary governing parameter for thickness noise, the nondimensional weight coefficient also impacts lower harmonic loading noise levels, which contribute strongly to low-frequency harmonic noise radiation both in and out of the plane of the horizon. Strategies for maintaining the same nondimensional rotor operating condition under different ambient conditions are developed using an analytical model of single main rotor helicopter trim and confirmed using a CAMRAD II model of the AS350 helicopter. The ability of the Fundamental Rotorcraft Acoustics Modeling from Experiments (FRAME) technique to generalize noise measurements made under one set of ambient conditions to make accurate noise predictions under other ambient conditions is also validated.

Steady and Unsteady Hover Test of a Lift Compounding Rotorcraft

Compound rotorcraft can be a viable option to achieve moderately high speeds of 240 knots. However, there is no systematic wind tunnel test data in the public domain to understand the complex interactional aeromechanics involved. This work is part of a test campaign to conduct multiple high-speed wind tunnel tests of various compound configurations. The compound configuration consists of a single main rotor, a single wing on the retreating side for lift augmentation, and a single rear propeller for propulsive augmentation. In this paper, hover test results are presented. Hover tests were conducted for various configurations and at different tip Mach numbers with the maximum value of 0.5. Rotor and wing performances are measured separately using individual balances. Rotor blades are equipped with strain gauges for measuring flap, lag, and torsional moments at radial stations. The wing location is below the main rotor, and the measured wing download is only 2.8% of the rotor thrust. The wing provides a partial ground effect, making the rotor more efficient and almost compensates for the download with less than 1% reduction in net figure of merit. Blade structural loads correlate satisfactorily between the analysis and measurement except for lag bending moment. This discrepancy is a large 2/rev lag bending moment measured in the test, which needs further investigation. Propeller hover test was conducted up to 2500 revolutions per minute (tip Mach number of 0.25) at a single collective. Prediction correlates well and captures the low-Reynolds-number effect for lower revolutions per minute.

Analytical and Computational Study of Hover Performance of Novel Dissimilar Coaxial Rotor

This paper studies a novel dissimilar coaxial rotor concept with significantly reduced rotor–rotor aerodynamic interaction during hovering flight. The proposed concept’s performance is systematically compared with regular coaxial and conventional single rotor configurations using 1) analytical and 2) blade element momentum theory (BEMT)-based analysis for hover flight condition. The hover performance among different rotor configurations is compared using a baseline main rotor combined with various antitorque mechanisms. Through this study it is established that the dissimilar coaxial rotor configuration shows improved performance for torque equilibrium cases when compared to both conventional and regular coaxial configurations for moderate- to high-thrust conditions with power reduction of 16–37% when compared to a conventional single main rotor–tail rotor configuration and 14–17% of power saving when compared to a regular coaxial rotor during hovering flight with rotor thrust in the range of . The analytical method from momentum theory and BEMT results correlate well for the induced and profile torque balanced conditions. The proposed concept appears to be an efficient alternative to the conventional single main rotor with tail rotor and regular coaxial rotor for hovering flight conditions within the scope of the analysis.

Wind Tunnel Test on a Slowed Mach-Scaled Hingeless Rotor with Lift Compounding

A single main rotor helicopter's maximum forward speed is limited due to the compressibility effects on the advancing side and reverse flow and dynamic stall on the retreating side. Compound helicopters can address these issues with a slowed rotor and lift compounding. There is a scarcity of test data on compound helicopters, and the present research focuses on a systematic wind tunnel test on lift compounding. Slowing down the rotor increases the advance ratio and, hence, the reverse flow region, which does not produce much lift. The lift is augmented with a wing on the retreating side. A hingeless rotor hub helps to balance the rolling moment with lift offset. Wind tunnel tests were carried out on this configuration up to advance ratios of 0.7 at two different wing incidence angles. Rotor performance, controls, blade structural loads, and hub vibratory loads were measured and compared with in-house comprehensive analysis, UMARC. A comparison between different wing incidences at constant total lift provided many insights into the lift compounding. It increased the vehicle efficiency and reduced peak-to-peak lag bending moment and in-plane 4/rev hub vibratory loads. The only trade-off was steady rotor hub loads and rolling moment at the wing root carried by the fuselage.

Helicopter Safe Landing Trajectory after Main Rotor Actuator Failures

Main rotor actuator failure leads to catastrophic accidents for single main rotor helicopters. This paper focuses on safe landing trajectories after an actuator is locked in place by the remaining actuators, without introducing other control inputs. A general swashplate geometry is described, and new reconfiguration solutions for the control mixer are presented. The safe landing trajectories are obtained by formulating a nonlinear optimal control problem based on a nonlinear helicopter dynamic model and geometry constraints due to actuator failure. Safe landing trajectory results are shown with various initial forward velocities of all actuator failure cases. The safe initial speed boundaries are also explored by employing speed sweeps.

Dynamic Stall Modeling Using Viscous Vortex Particle Method for Coaxial Rotors

Dynamic stall is an important source of vibrations on a rotor at high advance ratios. The periodic flow separation and reattachment during dynamic stall generates large unsteady loads. In this study, the flow separation is modeled as the shedding of concentrated vorticity from the leading edge of the airfoil. The viscous vortex particle method is used to calculate the evolution of the rotor wake. Blade loads are calculated using a reduced order model obtained from computational fluid dynamics, and dynamic stall loads are calculated using the ONERA dynamic stall model. Results are presented for single main rotor and coaxial rotors at advance ratios of μ = 0.3–0.4. The separated wake modifies the angle of attack distribution on the rotor and hence impacts the hub loads. The results indicate that the separated wake modifies the vibratory hub loads by 5–10% for a single main rotor at μ = 0.3. The vibratory hub loads for the coaxial rotor are modified by 10–20% at μ = 0.4 with the inclusion of the separated wake. The upper and lower rotor tip path planes are tilted such that the blade and wake interaction is greater on the retreating side of the upper rotor and decreased on the advancing side.

Fuel consumption of rotorcrafts and potentiality for hybrid electric power systems

The paper proposes a simulation approach to evaluate the power required by a rotorcraft in standard flight missions and in emergency landing maneuvers, and the corresponding fuel consumption, in order to compare the feasibility and potential fuel savings for different hybrid power systems. More in detail, three options are analyzed, namely electrification of the tail rotor, fully hybrid electric propulsion and electric emergency landing. Weight penalty and potential fuel saving for the proposed hybridization schemes are evaluated for an Agusta-Westland A109 twin engine helicopter model. Nonetheless the discussed methods of analysis have general validity for single main rotor helicopter configurations.Two different scenarios are considered in this investigation: current technologies for batteries and motors and improved electrical components, with performance projections as of 2040. According to this analysis, electrification of the tail rotor and parallel hybridization are feasible with available technology, whereas a fully electrical power system for emergency landing could be developed only in the future. Finally, a parallel hybrid electric power system is sized according to the analysis of power request over four different missions. Fuel savings are evaluated for different energy management strategies. According to the results of this investigation, the parallel hybrid electric power system with present-day and future technologies can save fuel up to 5% and 12%, respectively, with an appropriate energy management strategy.

Parametric analysis of a hybrid power system for rotorcraft emergency landing sequence

The use of a hybrid powertrain for a conventional single main rotor helicopter is investigated, with the objective of assessing its feasibility and its potential impact on improving safety, especially for single-engine rotorcraft. The study is focused on the characteristics of the powertrain and required battery pack. It is based on a simple analysis of power required in forward flight and the estimate of the total energy required for a powered landing manoeuvre after thermal engine failure. Current technologies are considered as well as expected improvements, especially as far as energy density and power density of the battery are concerned. The latter analysis is based on current trends for battery and motors technologies, in order to determine the technological breakthrough limit.

Effects of rotor blade-tip geometry on helicopter trim and control response

ABSTRACTFor performance improvement and noise reduction, swept and anhedral tips have been incorporated in advanced-geometry rotor blades. While there are aerodynamic benefits to these advanced tip geometries, they come at the cost of complicated structural design and weight penalties. The effect of these tip shapes on loads, vibration and aeroelastic response are also unclear. In this study, a comprehensive helicopter aeroelastic analysis which includes rotor-fuselage coupling shall be described and the analysis results for rotor blades with straight tip, tip sweep and tip anhedral shall be presented and compared. The helicopter modelled is a conventional one with a hingeless single main rotor and single tail rotor. The blade undergoes flap, lag, torsion and axial deformations. Tip sweep, pretwist, precone, predroop, torque offset and root offset are included in the model. Aerodynamic model includes Peters-He dynamic wake theory for inflow and the modified ONERA dynamic stall theory for airloads calculations. The complete 6-dof nonlinear equilibrium equations of the fuselage are solved for analysing any general flight condition. Response to pilot control inputs is determined by integrating the full set of nonlinear equations of motion with respect to time. The effects of tip sweep and tip anhedral on structural dynamics, trim characteristics and vehicle response to pilot inputs are presented. It is shown that for blades with tip sweep and tip anhedral/dihedral, the 1/rev harmonics of the root loads reduce while the 4/rev harmonics of the hub loads increase in magnitude. Tip dihedral is shown to induce a reversal of yaw rate for lateral and longitudinal cyclic input.

NDARC을 이용한 헬리콥터 개념설계 및 성능해석 검증 연구

본 연구에서는 단일로터-꼬리로터 형상의 일반 헬리콥터에 대한 개념설계 및 성능해석 기법을 정립하고자 회전익 항공기 개념설계 및 성능해석 코드인 NDARC을 이용하여 UH-60A 헬리콥터에 대한 개념설계 및 성능해석을 수행하였다. 적절히 가정한 임무 형상을 이용하여 UH-60A 헬리콥터의 개념설계를 수행한 뒤 설계 목표 값과의 비교를 통하여 UH-60A 헬리콥터의 형상 및 중량을 적절히 설계할 수 있음을 확인하였다. 더불어, 본 연구로부터 설계된 UH-60A 헬리콥터 모델을 이용하여 다양한 비행 조건에서의 성능해석을 수행한 뒤, 본 연구의 해석 결과를 UH-60A의 성능 시험 및 선행 연구의 해석 결과와 비교하여 본 연구 결과의 타당성을 검증하였다. 그 결과, 단일로터-꼬리로터 형상의 일반 회전익 항공기의 개념설계 및 성능해석 기법을 적절히 정립하였음을 확인하였다. A validation study is conducted for the conceptual design and performance analysis of UH-60A Black Hawk in order to establish the conceptual design and performance analysis techniques for conventional helicopters using a single main rotor and a tail rotor. As a tool for conceptual design and analysis, NDARC(NASA Design and Analysis of Rotorcraft) is used for the present study. The conceptual design for UH-60A is successfully validated as compared with the target values. Then, various performance analyses in hover and forward flight are conducted for the UH-60A model obtained from the present design work, and they are compared well with the wind tunnel test, flight test, and previous analyses using various analysis tools. Through this validation work, the conceptual design and performance analysis techniques for the conventional helicopter are appropriately established.

Influences of aerodynamic layout of single main rotor helicopter on flight performance

Influences of aerodynamic layout of single main rotor helicopter on flight performance

Model Predictive Control Architecture for Rotorcraft Inverse Simulation

A novel inverse simulation scheme is proposed for applications to rotorcraft dynamic models. The algorithm adopts an architecture that closely resembles that of a model predictive control scheme, where the controlled plant is represented by a high-order helicopter model. A fast solution of the inverse simulation step is obtained on the basis of a lower-order, simplified model. The resulting control action is then propagated forward in time using the more complex one. The algorithm compensates for discrepancies between the models by updating initial conditions for the inverse simulation step and introducing a simple guidance scheme in the definition of the tracked output variables. The proposed approach allows for the assessment of handling quality potential on the basis of the most sophisticated model, while keeping model complexity to a minimum for the computationally more demanding inverse simulation algorithm. The reported results, for an articulated blade, single main rotor helicopter model, demonstrate the validity of the approach.

Adaptive Failure Compensation of a Single Main-Rotor Helicopter

This paper develops an adaptive control strategy for failure compensation of a single main-rotor helicopter. The adaptive controller addresses the problem of instability due to failure in the main rotor or tail rotor assemblies, or unexpected variations in thrust or yaw due to damage to the rotor blades. An abstracted model of the helicopter is used for the development of the adaptive control laws. The effects of instability due to damage to the helicopter are modeled. A set of adaptive controllers are developed to stabilize the helicopter's states under failure. Simulation results establish the efficacy of the adaptive control scheme over the conventional control laws in providing stability.

Helicopter Tail Rotor Thrust and Main Rotor Wake Coupling in Crosswind Flight

The tail rotor of a helicopter with a single main rotor configuration can experience a significant reduction in thrust when the aircraft operates in crosswind flight. Brown’s vorticity transport model has been used to simulate a main rotor and tail rotor system translating at a sideslip angle that causes the tail rotor to interact with the main rotor tip vortices as they propagate downstream at the lateral extremities of the wake. The tail rotor is shown to exhibit a distinct directionally dependent mode during which tail rotors that are configured so that the blades travel forward at the top of the disk develop less thrust than tail rotors with the reverse sense of rotation. The range of flight speeds over which this mode exists is shown to vary considerably with the vertical location of the tail rotor. At low flight speeds, the directionally dependent mode occurs because the tail rotor is immersed within not only the downwash from the main rotor but also the rotational flow associated with clusters of largely disorganized vorticity within the main rotor wake. At higher flight speeds, however, the tail rotor is immersed within a coherent supervortex that strongly influences the velocity field surrounding the tail rotor.

Eulerian Simulation of the Fluid Dynamics of Helicopter Brownout

A computational model is presented that can be used to simulate the development of the dust cloud that can be entrained into the air when a helicopter is operated close to the ground in desert or dusty conditions. The physics of this problem, and the associated pathological condition known as ‘brownout’ where the pilot loses situational awareness as a result of his vision being occluded by dust suspended in the flow around the helicopter, is acknowledged to be very complex. The approach advocated here involves an approximation to the full dynamics of the coupled particulate-air system. Away from the ground, the model assumes that the suspended particles remain in near equilibrium under the action of aerodynamic forces. Close to the ground, this model is replaced by an algebraic sublayer model for the saltation and entrainment process. The origin of the model in the statistical mechanics of a distribution of particles governed by aerodynamic forces allows the validity of the method to be evaluated in context by comparing the physical properties of the suspended particulates to the local properties of the flow field surrounding the helicopter. The model applies in the Eulerian frame of reference of most conventional Computational Fluid Dynamics codes and has been coupled with Brown’s Vorticity Transport Model. Verification of the predictions of the coupled model against experimental data for particulate entrainment and transport in the flow around a model rotor are encouraging. An application of the coupled model to analyzing the differences in the geometry and extent of the dust clouds that are produced by single main rotor and tandem-rotor configurations as they decelerate to land has shown that the location of the ground vortex and the size of any regions of recirculatory flow, should they exist, play a primary role in governing the extent of the dust cloud that is created by the helicopter.

Optimum Design of a Compound Helicopter

A design and aeromechanics investigation was conducted for a 100,000-lb compound helicopter with a single main rotor, which is to cruise at 250 knots at 4000 ft/95 deg F condition. Performance, stability, and control analyses were conducted with the comprehensive rotorcraft analysis CAMRAD II. Wind tunnel test measurements of the performance of the H-34 and UH-1D rotors at high advance ratio were compared with calculations to assess the accuracy of the analysis for the design of a high speed helicopter. In general, good correlation was obtained when an increase of drag coefficients in the reverse flow region was implemented. An assessment of various design parameters (disk loading, blade loading, wing loading) on the performance of the compound helicopter was conducted. Lower wing loading (larger wing area) and higher blade loading (smaller blade chord) increased aircraft lift-to-drag ratio. However, disk loading has a small influence on aircraft lift-to-drag ratio. A rotor parametric study showed that most of the benefit of slowing the rotor occurred at the initial 20 to 30% reduction of the advancing blade tip Mach number. No stability issues were observed with the current design. Control derivatives did not change significantly with speed, but the did exhibit significant coupling.

Modelling the interaction of helicopter main rotor and tail rotor wakes

AbstractThe mutual interaction between the main rotor and tail rotor wakes is central to some of the most problematic dynamic phenomena experienced by helicopters. Yet achieving the ability to model the growth and propagation of helicopter rotor wakes with sufficient realism to capture the details of this interaction has been a significant challenge to rotorcraft aerodynamicists for many decades. A novel computational fluid dynamics code tailored specifically for rotorcraft applications, the vorticity transport model, has been used to simulate the interaction of the rotors of a helicopter with a single main rotor and tail rotor in both hover and low-speed quartering flight, and with the tail rotor rotating both top-forward and top-aft. The simulations indicate a significant level of unsteadiness in the performance of both main and tail rotors, especially in quartering flight, and a sensitivity to the direction of rotation of the tail rotor. Although the model thus captures behaviour that is similar to that observed in practice, the challenge still remains to integrate the information from high fidelity simulations such as these into routine calculations of the flight dynamics of helicopters.

Recent Development of Rotorcraft Configuration

In view of the flight theory of rotorcraft, breakthroughs in the field of rotorcraft configuration during recent twenty years are described in this paper. From the traditional configuration (including the single main rotor with tail rotor, tandem rotors, side by side twin rotors, coaxial rotors and crossing rotors) to the recent ones (including ducted tail rotor, non-tail rotor helicopter, hingeless rotor, bearingless rotor, tilt-rotor, tip-jet rotor, reverse velocity rotor system, the compound helicopter and rotor/wing transition helicopter), the great progress on rotorcraft configuration have been revolutionarily made. These new technologies have an effect on the rotorcraft performance in high speed and supermaneuverability and high-agility flights. In light of advancing blade transonic shock and stall of retreating blade at high angle of attack in forward flight for helicopter conventional layout, the new progress in rotorcraft configuration in recent years show the potential to exceed maximum flight speed limit for current helicopters, and even to be greater than the cruise speed of propeller aircraft. At the same time, maneuverability and stability in forward flight have been improved. With the increased rotor efficiency, overload capability and good flight performance, these new types of rotorcraft have good maneuverability and lower vibration/noise level, and avoid ground or air resonance. Owing to the introduction of the new types of rotor configuration, the reliability and maintainability as well as survivability of the rotorcraft can be improved significantly. Therefore, their service life has been prolonged. In short, the rotorcraft with new types of configuration will become the more economical, effective and rapid vehicles for air traffic transportation in the future. Keywords: Rotorcraft, configuration, ducted tail rotor, non-tail rotor helicopter, hingeless rotor, bearingless rotor, tilt-rotor, tipjet rotor, reverse velocity rotor, compound helicopter