Search results

The purpose of this paper is to discuss published research in rotorcraft which has taken place in India during the last ten years. The helicopter research is divided into the following parts: health monitoring, smart rotor, design optimization, control, helicopter rotor dynamics, active control of structural response (ACSR) and helicopter design and development. Aspects of health monitoring and smart rotor are discussed in detail. Further work needed and areas for international collaboration are pointed out.

The archival journal papers on helicopter engineering published from India are obtained from databases and are studied and discussed. The contribution of the basic research to the state‐of‐the‐art in helicopter engineering science is brought out.

It is found that strong research capabilities have developed in rotor system health and usage monitoring, rotor blade design optimization, ACSR, composite rotor blades and smart rotor development. Furthermore, rotorcraft modeling and analysis aspects are highly developed with considerable manpower available and being generated in these areas.

Two helicopter projects leading to the “advanced light helicopter” and “light combat helicopter” have been completed by Hindustan Aeronautics Ltd These helicopter programs have benefited from the basic research and also provide platforms for further basic research and deeper industry academic collaborations. The development of well‐trained helicopter engineers is also attractive for international helicopter design and manufacturing companies. The basic research done needs to be further developed for practical and commercial applications.

This is the first comprehensive research on rotorcraft research in India, an important emerging market, manufacturing and sourcing destination for the industry.

This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.

Aeroelastic analysis of helicopter rotor with elastic blades based on finite element method in space and time and capable of considering higher harmonic control inputs is carried out. Moderate deflection and coriolis nonlinearities are included in the analysis. The rotor aerodynamics are represented using free wake and unsteady aerodynamic models.

Good correlation between analysis and HART‐II wind tunnel test data is obtained for blade natural frequencies across a range of rotating speeds. The basic physics of the blade mode shapes are also well captured. In particular, the fundamental flap, lag and torsion modes compare very well. The blade response compares well with HART‐II result and other high‐fidelity aeroelastic code predictions for flap and torsion mode. For the lead‐lag response, the present analysis prediction is somewhat better than other aeroelastic analyses.

Predicted blade response trend with higher harmonic pitch control agreed well with the wind tunnel test data, but usually contained a constant offset in the mean values of lead‐lag and elastic torsion response. Improvements in the modeling of the aerodynamic environment around the rotor can help reduce this gap between the experimental and numerical results.

Correlation of predicted aeroelastic response with wind tunnel test data is a vital step towards validating any helicopter aeroelastic analysis. Such efforts lend confidence in using the numerical analysis to understand the actual physical behavior of the helicopter system. Also, validated numerical analyses can take the place of time‐consuming and expensive wind tunnel tests during the initial stage of the design process.

While the basic physics appears to be well captured by the aeroelastic analysis, there is need for improvement in the aerodynamic modeling which appears to be the source of the gap between numerical predictions and HART‐II wind tunnel experiments.

The purpose of this study is to obtain optimum locations, peak deflection and chord of the twin trailing-edge flaps and optimum torsional stiffness of the helicopter rotor blade to minimize the vibration in the rotor hub with minimum requirement of flap control power.

Kriging metamodel with three-level five variable orthogonal array-based data points is used to decouple the optimization problem and actual aeroelastic analysis.

Some very good design solutions are obtained using this model. The best design point in minimizing vibration gives about 81 per cent reduction in the hub vibration with a penalization of increased flap power requirement, at normal cruise speed of rotor-craft flight.

One of the major challenges in the helicopters is the high vibration level in comparison with fixed wing aircraft. The reduction in vibration level in the helicopter improves passenger and crew comfort and reduces maintenance cost.

This paper presents design optimization of the helicopter rotor blade combining five design variables, such as the locations of twin trailing-edge flaps, peak deflection and flap chord and torsional stiffness of the rotor. Also, this study uses kriging metamodel to decouple the complex aeroelastic analysis and optimization problem.

To investigate the use of centre of gravity location on reducing cyclic pitch control for helicopter UAV's (unmanned air vehicles) and MAV's (micro air vehicles). Low cyclic pitch is a necessity to implement the swashplateless rotor concept using trailing edge flaps or active twist using current generation low authority piezoceramic actuators.

An aeroelastic analysis of the helicopter rotor with elastic blades is used to perform parametric and sensitivity studies of the effects of longitudinal and lateral center of gravity (cg) movements on the main rotor cyclic pitch. An optimization approach is then used to find cg locations which reduce the cyclic pitch at a given forward speed.

It is found that the longitudinal cyclic pitch and lateral cyclic pitch can be driven to zero at a given forward speed by shifting the cg forward and to the port side, respectively. There also exist pairs of numbers for the longitudinal and lateral cg locations which drive both the cyclic pitch components to zero at a given forward speed. Based on these results, a compromise optimal cg location is obtained such that the cyclic pitch is bounded within ±5° for a BO105 helicopter rotor.

The reduction in the cyclic pitch due to helicopter cg location is found to significantly reduce the maximum magnitudes of the control angles in flight, facilitating the swashplateless rotor concept. In addition, the existence of cg locations which drive the cyclic pitches to zero allows for the use of active cg movement as a way to replace the cyclic pitch control for helicopter MAV's.

The purpose of this paper is to develop an integrated rotorcraft design and virtual manufacturing framework. The framework consists of two major sub‐frameworks which are e‐design and virtual manufacturing frameworks. This paper aims to describe the process of generating a specific framework for helicopter design and manufacturing in general, and a method for main rotor blade design.

The e‐design process integrates a pre‐conceptual, conceptual and preliminary design phases and includes many high accuracy physics‐based analysis tools and in‐house codes. The development of analysis programs and integration of flow data are discussed under the e‐design process. The virtual manufacturing process discusses physical three‐dimensional (3D) prototypes using rapid prototyping, virtual process simulation model development using Delmia Quest, virtual machine tool simulation and process‐based cost model. Vehicle geometry is modelled parametrically in computer‐aided 3D interactive application (CATIA) V5 to enable integration between the e‐design and virtual manufacturing processes, and then saved in Enovia SmartTeam which is commercial software for product data management (PDM). Data saved in Enovia SmartTeam are used as a database for the virtual manufacturing process.

The integration framework was constructed by using Model Center software. A multi‐disciplinary design optimization loop for rotor blade considering manufacturing factors is discussed to demonstrate the robustness and efficiency of the framework.

The manufacturing (practical factors) could be considered at an early stage of the rotor blades design.

The gap between theoretical (engineering design: aerodynamic, structural, dynamic, design, etc.) and practical aspects (manufacturing) is bridged through integrated product/process development framework. The modern concurrent engineering approach is addressed for helicopter rotor blade design throughout the case study.

The purpose of this paper is to evaluate the prediction capability of comprehensive structural dynamics (CSD) analysis codes for the higher harmonic control aeroacoustic rotor test (HART) II data.

A nonlinear flexible multibody dynamics analysis code DYMORE, as well as the comprehensive analytical model of rotorcraft aerodynamics and dynamics (CAMRAD) II, are used to perform the task. The predicted results on rotating free vibration analysis, airloads, blade elastic motions, and structural moments are correlated with the measured data for the baseline, minimum noise, and minimum vibration cases.

The DYMORE analysis results with a free wake model show a good performance in capturing blade vortex interaction peaks in the prediction of section normal forces but apparently with a phase shift problem. The high‐frequency behavior in the airloads signal does not affect much on the aeroelastic response and structural moments of the rotor.

The present approach uses two separate CSD codes to systematically validate the HART II data. The accuracy of each code on structural dynamic aspects of HART II rotor is assessed using a consistent set of inputs. The effects of blade tip deflections on the interaction of blades and their trailed vortices leading to a reduced noise emission are also investigated.

The purpose of this paper is to determine dependencies between a rotor-blade shape and a rotor performance as well as to search for optimal shapes of blades dedicated for helicopter main and tail rotors.

The research is conducted based on computational methodology, using the parametric-design approach. The developed parametric model takes into account several typical blade-shape parameters. The rotor aerodynamic characteristics are evaluated using the unsteady Reynolds-averaged Navier–Stokes solver. Flow effects caused by rotating blades are modelled based on both simplified approach and truly 3D simulations.

The computational studies have shown that the helicopter-rotor performance may be significantly improved even through relatively simple aerodynamic redesigning of its blades. The research results confirm high potential of the developed methodology of rotor-blade optimisation. Developed families of helicopter-rotor-blade airfoils are competitive compared to the best airfoils cited in literature. The finally designed rotors, compared to the baselines, for the same driving power, are characterised by 5 and 32% higher thrust, in case of main and tail rotor, respectively.

The developed and implemented methodology of parametric design and optimisation of helicopter-rotor blades may be used in future studies on performance improvement of rotorcraft rotors. Some of presented results concern the redesigning of main and tail rotors of existing helicopters. These results may be used directly in modernisation processes of these helicopters.

The presented study is original in relation to the developed methodology of optimisation of helicopter-rotor blades, families of modern helicopter airfoils and innovative solutions in rotor-blade-design area.