Search results

Relative to in-service aero-engines, the bypass ratio of future civil architectures may increase further. If traditional design rules are applied to these new configurations and the housing components are scaled, then it is expected that the overall weight, nacelle drag and the effects of aircraft integration will increase. For this reason, the next generation of civil turbofan engines may use compact nacelles to maximise the benefits from the new engine cycles. The purpose of this paper is to present a multi-level design and optimisation process for future civil aero-engines.

An initial set of multi-point, multi-objective optimisations for axisymmetric configurations are carried out to identify the trade-off between intake and fancowl bulk parameters of highlight radius and nacelle length on nacelle drag. Having identified the likely optimal part of the design space, a set of computationally expensive optimisations for three-dimensional non-axisymmetric configurations is performed. The process includes cruise- and windmilling-type operating conditions to ensure aerodynamic robustness of the downselected configurations.

Relative to a conventional aero-engine nacelle, the developed process yielded a compact aero-engine configuration with mid-cruise drag reduction of approximately 1.6% of the nominal standard net thrust.

The multi-point, multi-objective optimisation is carried out with a mixture of regression and classification functions to ensure aerodynamic robustness of the downselected configurations. The developed computational approach enables the optimisation of future civil aero-engine nacelles that target a reduction of the overall fuel consumption.

Aerodynamic shape optimisation is complex because of the high dimensionality of the problem, the associated non-linearity and its large computational cost. These three aspects have an impact on the overall time of the design process. To overcome these challenges, this paper aims to develop a method for transonic aerodynamic design with dimensionality reduction and multifidelity techniques.

The developed methodology is used for the optimisation of an installed civil ultra-high bypass ratio aero-engine nacelle. As such, the effects of airframe-engine integration are considered during the optimisation routine. The active subspace method is applied to reduce the dimensionality of the problem from 32 to 2 design variables with a database compiled with Euler computational fluid dynamics (CFD) calculations. In the reduced dimensional space, a co-Kriging model is built to combine Euler lower-fidelity and Reynolds-averaged Navier stokes higher-fidelity CFD evaluations.

Relative to a baseline aero-engine nacelle derived from an isolated optimisation process, the proposed method yielded a non-axisymmetric nacelle configuration with an increment in net vehicle force of 0.65% of the nominal standard net thrust.

This work investigates the viability of CFD optimisation through a combination of dimensionality reduction and multifidelity method and demonstrates that the developed methodology enables the optimisation of complex aerodynamic problems.

Helicopter rotor blades are usually aerodynamically limited by the severe conditions present in every revolution: strong shock wave boundary layer interaction on the advancing side and dynamic stall on the retreating side. Therefore, different flow control strategies might be applied to improve the aerodynamic performance.

The present research is focussed on the application of passive rod vortex generators (RVGs) to control the flow separation induced by strong shock waves on helicopter rotor blades. The formation and development in time of the streamwise vortices are also investigated for a channel flow.

The proposed RVGs are able to generate streamwise vortices as strong as the well-known air-jet vortex generators. It has been demonstrated a faster vortex formation for the rod type. Therefore, this flow control device is preferred for applications in which a quick vortex formation is required. Besides, RVGs were implemented on helicopters rotor blades improving their aerodynamic performance (ratio thrust/power consumption).

A new type of vortex generator (rod) has been investigated in several configurations (channel flow and rotor blades).

Aerodynamic shape optimisation is a complex problem usually governed by transonic non-linear aerodynamics, a high dimensional design space and high computational cost. Consequently, the use of a numerical simulation approach can become prohibitive for some applications. This paper aims to propose a computationally efficient multi-fidelity method for the optimisation of two-dimensional axisymmetric aero-engine nacelles.

The nacelle optimisation approach combines a gradient-free algorithm with a multi-fidelity surrogate model. Machine learning based on artificial neural networks (ANN) is used as the modelling technique because of its ability to handle non-linear behaviour. The multi-fidelity method combines Reynolds-averaged Navier Stokes and Euler CFD calculations as high- and low-fidelity, respectively.

Ratios of low- and high-fidelity training samples to degrees of freedom of n_LF/n_DOFs = 50 and n_HF/n_DOFs = 12.5 provided a surrogate model with a root mean squared error less than 5% and a similar convergence to the optimal design space when compared with the equivalent CFD-in-the-loop optimisation. Similar nacelle geometries and aerodynamic flow topologies were obtained for down-selected designs with a reduction of 92% in the computational cost. This highlights the potential benefits of this multi-fidelity approach for aerodynamic optimisation within a preliminary design stage.

The application of a multi-fidelity technique based on ANN to the aerodynamic shape optimisation problem of isolated nacelles is the key novelty of this work. The multi-fidelity aspect of the method advances current practices based on single-fidelity surrogate models and offers further reductions in computational cost to meet industrial design timescales. Additionally, guidelines in terms of low- and high-fidelity sample sizes relative to the number of design variables have been established.

The decrease in specific thrust achieved by Ultra-High Bypass Ratio (UHBPR) aero-engines allows for a reduction in specific fuel consumption. However, the typical associated larger fan size might increase the nacelle drag, weight and the detrimental interference effects with the airframe. Consequently, the benefits from the new UHBPR aero-engine cycle may be eroded. This paper aims to evaluate the potential improvement in the aerodynamic performance of compact nacelles for installed aero-engine configuration.

Drooped and scarfed non-axisymmetric compact and conventional nacelle designs were down selected from a multi-point CFD-based optimisation. These were computationally assessed at a set of installation positions on a contemporary wide-body, twin-engine transonic aircraft. Both cruise and off-design conditions were evaluated. A thrust and drag accounting method was applied to evaluate different aircraft, powerplant and nacelle performance metrics.

The aircraft with the compact nacelle configuration installed at a typical installation position provided a reduction in aircraft cruise fuel consumption of 0.44% relative to the conventional architecture. However, at the same installation position, the compact design exhibits a large flow separation at windmilling conditions that is translated into an overall aircraft drag penalty of approximately 5.6% of the standard cruise net thrust. Additionally, the interference effects of a compact nacelle are more sensitive to deviations in mass flow capture ratio (MFCR) from the nominal windmilling diversion condition.

This work provides a comprehensive analysis of not only the performance but also the aerodynamics at an aircraft level of compact nacelles compared to conventional configurations for a range of installations positions at cruise. Additionally, the engine-airframe integration aerodynamics is assessed at an off-design windmilling condition which constitutes a key novelty of this paper.

This paper aims to present the results of experimental and numerical research on heat transfer distribution under the impinging jets at various distances from the wall and high jet velocity. This work is a part of the INNOLOT Program financed by National Centre for Research and Development.

The air jets flow out from the common pipe and impinge on a surface which is cooled by them, and in this way, all together create a model of external cooling system of low-pressure gas turbine casing. Measurements were carried out for the arrangement of 26 in-line jets with orifice diameter of 0.9 mm. Heat transfer distribution was investigated for various Reynolds and Mach numbers. The cooled wall, made of transparent PMMA, was covered with a heater foil on which a layer of self-adhesive liquid crystal foil was placed. The jet-to-wall distance was set to h = from 4.5 to 6 d.

The influence of various Reynolds and Mach numbers on cooled flat plate and jet-to-wall distance in terms of heat transfer effectiveness is presented. Experimental results used for the computational fluid dynamics (CFD) model development, validation and comparison with numerical results are presented.

Impinging air jets is a commonly used technique to cool advanced turbines elements, as it produces large convection enhancing the local heat transfer, which is a critical issue in the development of aircraft engines.

The achieved results present experimental investigations carried out to study the heat transfer distribution between the orthogonally impinging jets from long round pipe and flat plate. Reynolds number based on the jet orifice exit conditions was varied between 2,500 and 4,000; meanwhile, for such Re, the flow velocity in jets was particularly very high, changing from M = 0.56 to M = 0.77. Such flow conditions combination, i.e. the low Reynolds number and very high flow velocity cannot be found in the existing literature.