Descent Flight Conditions Research Articles

Aeromechanics of a Coaxial Ultralight Rotorcraft During Turn, Climb, and Descent Flight

This work presents the aeromechanics analysis and evaluation of an ultralight counter-rotating coaxial rotorcraft using a combined numerical and experimental approach with a focus on turn, climb, and descent flight conditions. The helicopter features a 2-by-2 bladed coaxial rotor system with teeter hinges, 6.5 m in diameter and uniform airfoil sections of the linearly twisted blades. Hover and forward-flight test data are shown with varying turn rates or climb and descent velocities. The predictive capabilities of a CAMRAD II numerical model are used to extend and support the flight test data and evaluate possible flight-envelope limitations in terms of ceiling limits and maximum climb rates. The paper focuses on rotor performance, thrust, horizontal-stabilizer loads, and blade-tip clearance, with special attention on the effects from interactional rotor aerodynamics. Results in terms of rotor performance, rotor thrust, horizontal-stabilizer loads, and rotorcraft aerodynamic angles of attack resulted in satisfactory correlation between the numerical model and the measurements. Rotor–rotor interactional effects contributed to the approaching rotor-blade tips, especially during low-speed descent flights. During turn flight, required rotor power increases with greater bank angles. Maximum axial climb rate of the ultralight coaxial helicopter was at , the hover ceiling at 2180 m, and the service ceiling at 4450 m. Ceiling limitations depended on the operating conditions and reduced with greater temperature and rotorcraft weight.

The second wind-tunnel test of DLR’s multiple swashplate system

After proving its individual blade control (IBC) capabilities on a four-bladed rotor system in the wind tunnel, the DLR’s multiple swashplate system and its control software were developed further to allow IBC operation on a five-bladed rotor system. After preliminary tests in hover conditions, a second wind-tunnel test was performed in the DNW-LLF wind tunnel in late 2016. The goal of this test on a Mach-scaled model rotor was to reduce vibration, noise, and required rotor power in different flight conditions using proven IBC strategies as well as localized pitch control (LPC) on five blades. In descent flight condition, significant reductions of blade–vortex interaction noise relative to the baseline case were achieved on both sides of the rotor disk through the application of 2/rev higher harmonic control as well as using an LPC schedule. Through the application of 3/rev IBC as well as a vibration controller using a 4–6/rev multi-harmonic IBC signal, 5/rev hub loads were reduced significantly and the 5/rev vertical vibrations were nearly eliminated. In addition, during simulated high-speed-level flight with a wind speed of 76 m/s, the required rotor power was successfully reduced using a 2/rev input with an amplitude of 1 $$^\circ$$ .

The First Wind Tunnel Test of the Multiple Swashplate System: Test Procedure and Principal Results

In September 2015, the DLR completed the first wind tunnel test of its patented Multiple Swashplate System in the DNW's Large Low-Speed Facility in the Netherlands. During these tests, the potential of this new active rotor control system to effectively reduce noise, vibration, and power consumption using several individual blade control strategies, including 2/rev inputs, was successfully demonstrated on two different model rotors without using actuators in the rotating frame. For a Mach-scaled Bo105 model, rotor power reductions of up to 4% in level flight and reductions of the weighted 4/rev vibratory hub loads of up to 77% in a descent flight condition were measured, and BVI noise levels were reduced by up to 4.5 dB by the application of 3/rev higher harmonic control, confirming findings from the HART II test. For the second model rotor using a contemporary blade geometry, the maximum reductions were measured at 3%, 52%, and 3.9 dB for power, vibration, and BVI noise, respectively.

능동 탭 로터 모델링 및 BVI 소음 저감 특성 해석

Active tab is one of the promising technology for the BVI (blade-vortex interaction) noise reduction, and analysis of noise reduction performance is very important phase of the technology development. For the purpose of analysing the performance of noise reduction using active tab, CAMRAD II model for a model-scale rotor system was constructed utilizing structural design result and airfoil aerodynamic data generated by CFD (computational fluid dynamics) calculation. HHC strategy was applied to descent flight condition and air-load was calculated by CAMRAD II then variations of BVI noise was calculated by in-house program. Calculation result with respect to tab length and control phase changes showed BVI noise could be reduced by -3.3dB.

Analysis of a structural-aerodynamic fully-coupled formulation for aeroelastic response of rotorcraft

This paper deals with a computational aeroelastic tool aimed at the analysis of the response of rotary wings in arbitrary steady motion. It has been developed by coupling a nonlinear beam model for blades structural dynamics with a potential-flow boundary integral equation solver for the prediction of unsteady aerodynamic loads around three-dimensional, lifting bodies. The Galerkin method is used for the spatial integration of the resulting differential aeroelastic system, whereas the periodic blade response is determined by a harmonic balance approach. This aeroelastic model yields a unified approach for aeroelastic response and blade pressure prediction, that may conveniently be used for aeroacoustic purposes. It is able to examine configurations where blade–vortex interactions occur. Numerical results show the capability of the aeroelastic tool to evaluate blade response and vibratory hub loads for a helicopter main rotor in level and descent flight conditions, and examine the efficiency and robustness of the different numerical solution algorithms that may be applied in the developed aeroelastic solver. Comparisons among aeroelastic predictions based on different aerodynamic models are also presented.

Analysis of the Helishape Descent Database with Regards to Blade-Wake Interaction Noise

Helicopter rotor Blade Wake Interaction (BWI) noise is known to be significant during take-off and level flight. Less attention has been given to descent flight conditions, where Blade Vortex Interaction (BVI) noise is dominant. The present study investigates BWI noise in descent flight conditions by analyzing the HELISHAPE descent case database. Through signal analysis of both acoustic waveforms and spectra, the rotor azimuthal region responsible for BWI noise is localized and the dominance of BVI noise in the BWI frequency region is shown. Coherence analysis of the blade pressure data indicate significant chordwise coherence in the 3 to 4 Strouhal number range and absence of acoustic dipoles in the BWI frequency range. Comparison with analysis of a swept-back blade-tip database showed that blade-tip shape does not affect BWI noise. The findings of this study support BWI prediction models based on Amiet's theory and suggest that BWI noise can be ignored for predictions of rotor noise in descent flight conditions.

Helicopter Rotor Noise in the Merged Tip-Vortex and Blade Interaction Condition

BVI (blade vortex interaction) phenomena cause unsteady loading and noise in the helicopter rotor. It is difficult to accurately predict BVI aerodynamics and noise, because rotor wake and tip-vortices are numerically dissipated. In the present study, CFD and a time-marching free-wake coupling analysis were used to describe inflow and outflow without wake dissipation. Rotor noise was also computed using the Farassat formula, derived from the Ffowcs-Williams Hawkings equation. To validate aerodynamics and the noise prediction solver, a two-blade rotor was simulated and the results were compared with experimental data. To investigate the vortex interaction of a multi-blade rotor, a four-blade rotor was also simulated. In a slow descent flight condition, tip vortices are merged at the advancing side. The merged tip-vortex and blade interaction were simulated and are discussed in this paper.

자유후류기법에 의한 고해상도 공기력과 음향상사법을 이용한 헬리콥터 로터 블레이드-와류 상호작용 소음 예측

The BVI(blade vortex interaction) noise Prediction has been one of the most challenging acoustic analyses in helicopter aeromechanical Phenomenon. It is well known high resolution airloads data with accurate tip vortex positions are necessary for the accurate prediction of this phenomenon. The truly unsteady time-marching free-wake method, which is able to capture the tip vortices instability in hover and axial flights, is expanded with the rotor flapping motion and trim routine to predict unsteady airloads in forward and descent flights. And Farassat formulation 1-A based on the FW-H equation is applied for the noise prediction considering the blade flapping motion. Main objective of this study is to validate the newly developed prediction code. To achieve the objective, the descent flight condition of AH-1 OLS(operational loads survey) configuration is analyzed using present code. The predicted sectional thrust distribution and sectional airloads time histories show the present scheme is able to capture well the unsteady airloads caused by a parallel BVI. Finally, the predicted noise data, observed in two different positions where are 3.44 times of rotor radius far from the hub center, are quite reasonable agreements with the experimental data compared to the other analysis results.

Tracking control of trim trajectories of a blimp for ascent and descent flight manoeuvres

A blimp is a small airship that has no metal framework and collapses when deflated. It belongs to family of unmanned aerial vehicles (UAVs). In this paper we address the problem of designing tracking feedback control of an underactuated autonomous UAV. The ascent and descent flight conditions as one in which the rate of change (of magnitude) of the airship's state vector is zero and the resultant of the applied forces and moments is constant lead to trimmed flight trajectories. The subject of the tracking control is to stabilize the engine around the planned flight. Using a combined integrator backstepping approach and Lyapunov theory, the stability results are local and overcome the minimum number of actuators (inputs) with respect to the blimp's six degrees of freedom. Considering physic limits in UAVs, other trimmed flights are investigated and compared.

Oscillating air jets for implementing blade variable twist, enhancing engine and blade efficiency, and reducing drag, vibration, download and IR signature

A porous surface on an aircraft structure driven with oscillating positive and negative pressures is used as an active control device for attenuating negative aerodynamic interactions. The porous surfaces can be driven with positive and negative pressures either continuously or when predetermined flight conditions are present. The porous surfaces can be used on rotor blades to reduce BVI noise in descent flight conditions. The porous surfaces can be configured on rotor blades for affecting blade variable twist in accordance with various flight conditions, and can further be incorporated for reducing rotor hub vibrations as well. Porous surfaces placed on aerodynamic surfaces below the rotor blades of a tiltrotor aircraft can attenuate or eliminate download and fountain flow conditions. When placed on the trailing edges of a tip jet-exhaust driven rotor blade, the porous surfaces of the present invention can supplement the tip jet momentum of the exhaust to thereby reduce an amount of exhaust needed to drive the rotor blade.

ヘリコプタのブレード/渦干渉騒音の解析

Two types of prediction tools for helicopter noise have been developed under the cooperative research between Advanced Technology Institute of Commuter Helicopter Ltd. (ATIC) and National Aerospace Laboratory (NAL) . One of them is a combined method of CAMRAD II, interpolation code for blade motion and wake geometry, aerodynamic code of 3D unsteady Euler solver, and aeroacoustic code based on Ffowcs Williams and Hawkings (FW-H) formulation. The other consists of CAMRAD II, 3D unsteady Euler solver using moving overlapped grid method, and FW-H code. The acoustic waveform of Blade-Vortex Interaction (BVI) noise predicted by the former tool is in good agreement with the experimental data of 1/7-scale model AH-1 Operational Loads Survey (OLS) rotor. This method is applied to investigate the effect of blade-tip shape on the intensity of BVI noise. As a result, it is shown that anhedral and swept-forward tip shapes effectively reduce the BVI noise of OLS rotor in a descent flight condition. The predicted Effective Perceived Noise Level (EPNL) of a helicopter is also compared with the experimental data obtained by ATIC and reasonable correlation is obtained. The latter tool successfully predicts the distinct spikes in the BVI wave-form of ATIC model rotor tested in German-Dutch Wind Tunnel (DNW) . In the comparison of measured and calculated carpet noise contours, reasonable agreement is obtained. The present tools are expected to be useful for the design of low-noise helicopters in the future.