An integrated methodology to assess the operational and environmental performance of a conceptual regenerative helicopter

This paper aims to present an integrated multidisciplinary simulation framework, deployed for the comprehensive assessment of combined helicopter powerplant systems at mission level. Analytical evaluations of existing and conceptual regenerative engine designs are carried out in terms of operational performance and environmental impact. The proposed methodology comprises a wide-range of individual modeling theories applicable to helicopter flight dynamics, gas turbine engine performance as well as a novel, physics-based, stirred reactor model for the rapid estimation of various helicopter emissions species. The overall methodology has been deployed to conduct a preliminary trade-off study for a reference simple cycle and conceptual regenerative twin-engine light helicopter, modeled after the Airbus Helicopters Bo105 configuration, simulated under the representative mission scenarios. Extensive comparisons are carried out and presented for the aforementioned helicopters at both engine and mission level, along with general flight performance charts including the payload-range diagram. The acquired results from the design trade-off study suggest that the conceptual regenerative helicopter can offer significant improvement in the payload-range capability, while simultaneously maintaining the required airworthiness requirements. Furthermore, it has been quantified through the implementation of a representative case study that, while the regenerative configuration can enhance the mission range and payload capabilities of the helicopter, it may have a detrimental effect on the mission emissions inventory, specifically for NOx (Nitrogen Oxides). This may impose a trade-off between the fuel economy and environmental performance of the helicopter. The proposed methodology can effectively be regarded as an enabling technology for the comprehensive assessment of conventional and conceptual helicopter powerplant systems, in terms of operational performance and environmental impact as well as towards the quantification of their associated trade-offs at mission level. Ali Fakhre, Ioannis Goulos, Vassilios Pachidis School of Engineering, Energy, Power and Propulsion Division, Cranfield University, Cranfield, Bedford, MK43 0AL, UK [email protected] The Aeronautical Journal, 2015, Vol 119, Issue 1211, pp1-24 Published by Cambridge University Press. This is the Author Accepted Manuscript. This article may be used for personal use only. The final published version (version of record) is available online at 10.1017/S0001924000010253. Please refer to any applicable publisher terms of use

Stochastic axial compressor variable geometry schedule optimisation

The design of axial compressors is dictated by the maximisation of flow efficiency at on design conditions whereas at part speed the requirement for operation stability prevails. Among other stability aids, compressor variable geometry is employed to rise the surge line for the provision of an adequate surge margin. The schedule of the variable vanes is in turn typically obtained from expensive and time consuming rig tests that go through a vast combination of possible settings. The present paper explores the suitability of stochastic approaches to derive the most flow efficient schedule of an axial compressor for a minimum variable user defined value of the surge margin. A genetic algorithm has been purposely developed and its satisfactory performance validated against four representative benchmark functions. The work carries on with the necessary thorough investigation of the impact of the different genetic operators employed on the ability of the algorithm to find the global extremities in an effective and efficient manner. This deems fundamental to guarantee that the algorithm is not trapped in local extremities. The algorithm is then coupled with a compressor performance prediction tool that evaluates each individual's performance through a user defined fitness function. The most flow efficient schedule that conforms to a prescribed surge margin can be obtained thereby fast and inexpensively. Results are produced for a modern eight stage high bypass ratio compressor and compared with experimental data available to the research. The study concludes with the analysis of the existent relationship between surge margin and flow efficiency for the particular compressor under scrutiny. The study concludes with the analysis of the existent relationship between surge margin and flow efficiency for the particular compressor under scrutiny

A dynamic convergence control scheme for the solution of the radial equilibrium equation in through-flow analyses

One of the most frequently encountered numerical problems in scientific analyses is the solution of non-linear equations. Often the analysis of complex phenomena falls beyond the range of applicability of the numerical methods available in the public domain, and demands the design of dedicated algorithms that will approximate, to a specified precision, the mathematical solution of specific problems. These algorithms can be developed from scratch or through the amalgamation of existing techniques. The accurate solution of the full radial equilibrium equation (REE) in streamline curvature (SLC) through-flow analyses presents such a case. This article discusses the development, validation, and application of an 'intelligent' dynamic convergence control (DCC) algorithm for the fast, accurate, and robust numerical solution of the non-linear equations of motion for two-dimensional flow fields. The algorithm was developed to eliminate the large extent of user intervention, usually required by standard numerical methods. The DCC algorithm was integrated into a turbomachinery design and performance simulation software tool and was tested rigorously, particularly at compressor operating regimes traditionally exhibiting convergence difficulties (i.e. far off-design conditions). Typical error histories and comparisons of simulated results against experimental are presented in this article for a particular case study. For all case studies examined, it was found that the algorithm could successfully 'guide' the solution down to the specified error tolerance, at the expense of a slightly slower iteration process (compared to a conventional Newton-Raphson scheme). This hybrid DCC algorithm can also find use in many other engineering and scientific applications that require the robust solution of mathematical problems by numerical instead of analytical means

Multidisciplinary methodology for turbine overspeed analysis

In this paper, an integrated approach to turbine overspeed analysis is presented, taking into account the secondary air system dynamics and mechanical friction in a turbine assembly following an unlocated high-pressure shaft failure. The axial load acting on the rotating turbine assembly is a governing parameter in terms of overspeed protection since it governs the level of mechanical friction which acts against the turbine acceleration due to gas torque. The axial load is dependent on both the force coming from secondary air system cavities surrounding the disc and the force on the rotor blades. It is highly affected by secondary air system dynamics because rotor movement modifies the geometry of seals and flow paths within the network. As a result, the primary parameters of interest in this study are the axial load on the turbine rotor, the friction torque between rotating and static structures and the axial position of the rotor. Following an initial review of potential damage scenarios, several cases are run to establish the effect of each damage scenario and variable parameter within the model, with comparisons being made to a baseline case in which no interactions are modelled. This allows important aspects of the secondary air system to be identified in terms of overspeed prevention, as well as guidelines on design changes in current and future networks that will be beneficial for overspeed prevention

Improvements in the rotorcraft fuel economy and environmental impact through multiple-landing mission strategy

This paper presents an integrated rotorcraft multidisciplinary simulation framework, deployed for the comprehensive assessment of combined rotorcraft–powerplant systems performance at mission level. The proposed methodology comprises a wide-range of individual modelling theories applicable to rotorcraft performance and flight dynamics, gas turbine engine performance, and estimation of gaseous emissions (i.e. nitrogen oxides, NOx). The overall methodology has been deployed to conduct a comprehensive mission level feasibility study for a twin-engine light (TEL) rotorcraft, modeled after the Airbus Helicopters Bo105 configuration operating on a multiple-landing flying (MLF) mission approach compared to rotorcraft employing a conventional flying (CF) mission approach. The results of the analyses allow mission level assessment of the both aforementioned approaches for a wide-range of useful payload (UPL) values, mission range as well as mission level outputs (e.g. fuel burn, mission time, and gaseous emissions i.e. NOx). Furthermore, evaluation of engine cycle parameters (i.e. overall pressure ratio (OPR), turbine entry temperature (TET), and engine mass flow) are also carried out with respect to both approaches. The results acquired through the parametric analyses suggest that the MLF mission approach has the potential to significantly reduce rotorcraft mission fuel burn as well as gaseous emission (i.e. NOx). It has also been established through the acquired results that rotorcraft employing the MLF mission approach requires lower engine operating power throughout the entire mission duration, and therefore operates on a relatively lower engine OPR, combustor entry temperature, mass flow, rotational speed, and the TET compared to rotorcraft employing CF mission approach. It is emphasized that such operation of the engine can potentially improve the rate at which the engine components (i.e. compressor, combustor, and turbine) may deteriorate, thus the MLF mission approach can potentially provide further benefit in terms of engine maintenance and overall engine life. Finally it has been emphasised that the mission total range is a critical parameter in determining the level of benefit that can be attained from the employment of MLF mission approach

Modelling and analysis of coupled flap-lag-torsion vibration characteristics helicopter rotor blades

This paper presents the development of a mathematical approach targeting the modelling and analysis of coupled flap-lag-torsion vibration characteristics of non-uniform continuous rotor blades. The proposed method is based on the deployment of Lagrange’s equation of motion to the three-dimensional kinematics of rotor blades. Modal properties derived from classical-beam and torsion theories are utilized as assumed deformation functions. The formulation, which is valid for hingeless, freely hinged and spring-hinged articulated rotor blades, is reduced to a set of closed-form integral expressions. Numerical predictions for mode shapes and natural frequencies are compared with experimental measurements, non-linear finite element analyses and multi-body dynamics analyses for two small-scale hingeless rotor blades. Excellent agreement is observed. The effect of different geometrical parameters on the elastic and inertial coupling is assessed. Additionally, the effect of the inclusion of gyroscopic damping is evaluated. The proposed method, which is able to estimate the first seven coupled modes of vibration in a fraction of a second, exhibits excellent numerical stability. It constitutes a computationally efficient alternative to multi-body dynamics and finite element analysis for the integration of rotor blade flexible modelling into a wider comprehensive rotorcraft tool

An improved streamline curvature-based design approach for transonic axial-flow compressor blading

The increasing demand to obtain more accurate turbomachinery blading performance in the design and analysis process has led to the development of higher fidelity flow field models. Despite extensive flow field information can be collected from threedimensional (3-D) Reynolds-averaged Navier-Stokes (RANS) numerical simulations; it comes at a high computational cost in terms of time and resources, particularly if a comprehensive design space is explored during optimization. In contrast, through-flow methods such as streamline curvature (SLC), provide a flow solution in minutes whilst offering acceptable results. Furthermore, if the SLC fidelity is improved, a more detailed component-blading study is expected. For this reason, a fully-detailed transonic flow framework was implemented and validated in an existing in-house two-dimensional (2-D) SLC compressor performance to improve the performance results fidelity in transonic conditions. The improvements consist of two sections: (1) blade-profile modelling; (2) flow field solution. The bladeprofile modelling considers a 3-D blade-element-layout method to correctly model the sweep and lean angle, which determine the shock structure. The essential part of the transonic flow framework is its solution, formed of two parts: (1) a physics-based shock-wave model to predict its structure, and associated losses; (2) and a novel choking model to define the choke level for future spanwise mass flow redistribution. To demonstrate the functionality of the full comprehensive transonicflow approach, the well-known NASA Rotor 67 compressor was used to prove that the inlet relative flow angle should be limited by the choking incidence at the required blade span locations. A 3-D RANS numerical simulation for the NASA Rotor 67 validated the transonic-flow model, showing minimum differences in the spanwise mass flow distribution for the choked off-design cases. The current improvements implemented in the 2-D SLC compressor/fan performance simulator enhance the fidelity not only in analysis mode, but also in design optimisation applications

Mission performance analysis of a conceptual coaxial rotorcraft for air taxi applications

The rotorcraft industry has recently shown a new interest in compound rotorcraft as a feasible alternative to tackle the rapid growth of civil aviation activities and associated environmental impact. Indeed, aircraft contribution to the global emissions of CO2CO2, NOxNOx, and noise are driving the development of innovative technologies and vehicles. At present, compound rotorcraft architectures are regarded by the industry as promising platforms that can potentially increase productivity at a reduced environmental cost. In order to quantify the benefits of compound rotorcraft, this paper details the performance analysis of a coaxial counter-rotating rotor configuration with a pusher propeller. A comprehensive approach targeting the assessment of the aforementioned rotorcraft design for civil applications is presented herein. The methodology developed encompasses a rotorcraft flight dynamics simulation module and an engine performance module, coupled with a gaseous emissions prediction tool for environmental impact studies. They have been integrated together to constitute a standalone performance simulation framework and verified with the performance calculations of Harrington's “rotor 1” and the Sikorsky X2TD. The method is then applied to evaluate the performance of a conceptual coaxial rotorcraft, during a notional inter-city air taxi mission, in terms of cruise altitude, speed, and range, overall mission time and environmental impact. The several trade-offs between these parameters highlight the need for an integrated optimisation process. Besides, the concept demonstrates the benefits of the compound rotorcraft architecture with a best range speed reaching 90 m/s leading to reduced response times and increase of round trips in a given time. As a consequence, operators will need fewer vehicles and heliports to cover the same areas. This outcome is highly attractive in the current growing market

Development of a streamline curvature axial-flow compressor performance simulator graphical-user-interface for design and research

The all-time interest to increase turbomachinery efficiencies and pressure ratios has led to the progression of more robust and accurate simulation methods and tools. Even though 3-D CFD analyses are highly detailed and despite the computational power nowadays, they can be costly in terms of time and resources. Conversely, 2-D SLC methods provide acceptable performance and flow field results in short times. Because of economical and practical reasons, SLC still represents the cornerstone for turbomachinery design. In the present, the knowledge demand from the academia community in the airbreathing engine field has been expanding year after year. Nevertheless, there are very few open-source turbomachinery solvers that can be accessed, where user needs to know at least the basics of the programming language syntax and familiarize with it. For these reasons, a GUI was developed for an existing in-house 2-D SLC axial-flow compressor performance code, called SOCRATES. A GUI in this context supports as a teaching mechanism to explain not only the method itself, but also the compressor aerodynamic behaviour. The SOCRATES GUI consists in the axial-flow compressor model setup, solution and visualization for geometry and results. This paper outlines the main features of the 2-D SLC GUI, and uses a two-stage fan to show the flow field parameters and compressor/fan map, showing a consistent agreement against measured data

A design approach for controlled blade-off in overspeeding turbines

Following a shaft failure or loss of load in a gas turbine engine, the turbine overspeeds due to the continuing expansion through the stage(s). The overspeed may result in hazardous conditions which have to be prevented. Several mitigation methods include the control system’s response by shutting the fuel flow, mechanical friction to reduce turbine acceleration, and blade release at a predetermined rotational speed. The release of the blades not only terminates the gas torque which accelerates the disk, but also increases the disk burst speed at reduced centrifugal load. In this manuscript, a design space exploration is presented to avoid disk burst by blade-off in a civil large turbofan engine through a parametric design of blade firtree and disk post system. The firtree design parameters used in the study are the contact angle between the blade firtree and the disk post, firtree bottom flank angle, firtree flank length and firtree thickness with respect to the disk post. The LS-DYNA finite element software was used in the simulations to generate possible failure scenarios. These were ‘disk burst’ and ‘blade-off’. Blade-off conditions manifested in two ways as a function of design parameters. The first type was blade release from serrations without disk post failure, and the second type was blade escape with disk post failure. Following the design space exploration, the effect of several design and material parameters on the design space was investigated. These parameters are; the contact friction coefficient between the blade firtree and disk post, firtree serration number, and the strain hardening behavior of the material