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Article
Publication date: 13 June 2019

Hanan Lu, Qiushi Li, Tianyu Pan and Ramesh Agarwal

For an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or…

Abstract

Purpose

For an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions.

Design/methodology/approach

This paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition.

Findings

After the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL.

Originality/value

The originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 29 no. 11
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 6 November 2018

Pericles Panagiotou, Efstratios Giannakis, Georgios Savaidis and Kyros Yakinthos

The purpose of this paper is to present the preliminary design of a medium altitude long endurance (MALE) unmanned aerial vehicle (UAV), focusing on the interaction…

Abstract

Purpose

The purpose of this paper is to present the preliminary design of a medium altitude long endurance (MALE) unmanned aerial vehicle (UAV), focusing on the interaction between the aerodynamic and the structural design studies.

Design/methodology/approach

The classic layout theory was used, adjusted for the needs of unmanned aircraft, including aerodynamic calculations, presizing methods and CFD, to estimate key aerodynamic and stability coefficients. Considering the structural aspects, a combination of layout, finite element methods and custom parameterized design tools were used, allowing automatic reshapes of the skin and the internal structural parts, which are mainly made of composite materials. Interaction loops were defined between the aforementioned studies to optimize the performance of the aerial vehicle, maximize the aerodynamic efficiency and reduce the structural weight.

Findings

The complete design procedure of a UAV is shown, starting from the final stages of conceptual design, up to the point where the detail design and mechanical drawings initiated.

Practical implications

This paper presents a complete view of a design study of a MALE UAV, which was successfully constructed and flight-tested.

Originality/value

This study presents a complete, synergetic approach between the configuration layout, aerodynamic and structural aspects of a MALE UAV.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 7
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 21 March 2019

Huan Zhao and Zhenghong Gao

The high probability of the occurrence of separation bubbles or shocks and early transition to turbulence on surfaces of airfoil makes it very difficult to design…

Abstract

Purpose

The high probability of the occurrence of separation bubbles or shocks and early transition to turbulence on surfaces of airfoil makes it very difficult to design high-lift and high-speed Natural-Laminar-Flow (NLF) airfoil for high-altitude long-endurance unmanned air vehicles. To resolve this issue, a framework of uncertainty-based design optimization (UBDO) is developed based on an adjusted polynomial chaos expansion (PCE) method.

Design/methodology/approach

The γ ̄Re-θt transition model combined with the shear stress transport k-ω turbulence model is used to predict the laminar-turbulent transition. The particle swarm optimization algorithm and PCE are integrated to search for the optimal NLF airfoil. Using proposed UBDO framework, the aforementioned problem has been regularized to achieve the optimal airfoil with a tradeoff of aerodynamic performances under fully turbulent and free transition conditions. The tradeoff is to make sure its good performance when early transition to turbulence on surfaces of NLF airfoil happens.

Findings

The results indicate that UBDO of NLF airfoil considering Mach number and lift coefficient uncertainty under free transition condition shows a significant deterioration when complicated flight conditions lead to early transition to turbulence. Meanwhile, UBDO of NLF airfoil with a tradeoff of performances under both fully turbulent and free transition conditions holds robust and reliable aerodynamic performance under complicated flight conditions.

Originality/value

In this work, the authors build an effective uncertainty-based design framework based on an adjusted PCE method and apply the framework to design two high-performance NLF airfoils. One of the two NLF airfoils considers Mach number and lift coefficient uncertainty under free transition condition, and the other considers uncertainties both under fully turbulent and free transition conditions. The results show that robust design of NLF airfoil should simultaneously consider Mach number, lift coefficient (angle of attack) and transition location uncertainty.

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Article
Publication date: 2 October 2017

M. Tahani, M. Masdari and M. Kazemi

This paper aims to analyze the influence of the changings in geometrical parameters on the aerodynamic performance of the control canard projectiles.

Abstract

Purpose

This paper aims to analyze the influence of the changings in geometrical parameters on the aerodynamic performance of the control canard projectiles.

Design/methodology/approach

Because of the mentioned point, the range of projectiles increment has a considerable importance, and the design algorithm of a control canard projectile was first written. Then, were studied the effects of canard geometric parameters such as aspect ratio, taper ratio and deflectable nose on lift to drag coefficient ratio, static margin based on the slender body theory and cross section flow.

Findings

The code results show that aspect ratio increment, results in an increase in lift-to-drag ratio of the missile, but increase in canard taper ratio results in increasing of lift-to-drag ratio at 1° angle of attack, while during increasing the canard taper ratio up to 0.67 at 4° angle of attack, lift to drag first reaches to maximum and then decreases. Also, static margin decreases with canard taper ratio and aspect ratio increment. The developed results for this type of missile were compared with same experimental and computational fluid dynamic (CFD) results and appreciated agreement with other results at angles of attack between 0° and 6°.

Practical implications

To design a control canard missile, the effect of each geometric parameter of canard needs to be estimated. For this purpose, the suitable algorithm is used. In this paper, the effects of canard geometric parameters, such as aspect ratio, taper ratio and deflectable nose on lift-to-drag coefficient ratio and static margin, were studied with help of the slender body theory and cross-section flow.

Originality/value

The contribution of this paper is to predict the aerodynamic characteristics for the control canard missile. In this study, the effect of the design parameter on aerodynamic characteristics can be estimated, and the effect of geometrical characteristics has been analyzed with a suitable algorithm. Also, the best lift-to-drag coefficient for the NASA Tandem Control Missile at Mach 1.75 was selected at various angles of attack. The developed results for this type of missile were compared with same experimental and CFD results.

Details

Aircraft Engineering and Aerospace Technology, vol. 89 no. 6
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 2 January 2019

Wienczyslaw Stalewski and Wieslaw Zalewski

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…

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Practical implications

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.

Originality/value

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.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 5
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 1 June 2006

M. Vázquez, A. Dervieux and B. Koobus

To propose an integrated algorithm for aerodynamic shape optimization of aircraft wings under the effect of aeroelastic deformations at supersonic regime.

Abstract

Purpose

To propose an integrated algorithm for aerodynamic shape optimization of aircraft wings under the effect of aeroelastic deformations at supersonic regime.

Design/methodology/approach

A methodology is proposed in which a high‐fidelity aeroelastic analyser and an aerodynamic optimizer are loosely coupled. The shape optimizer is based on a “CAD‐free” approach and an exact gradient method with a single adjoint state. The global iterative process yields optimal shapes in the at‐rest condition (i.e. with the aeroelastic deformations substracted).

Findings

The methodology was tested under different conditions, taking into account a combined optimization goal: to reduce the sonic boom production, while preserving the aerodynamic performances of flexible wings. The objective function model contains both aerodynamic parameters and an acoustic term based on the sonic boom downwards emission.

Practical implications

This paper proposes a shape optimization methodology developed by researchers but aiming at the final strategic goal of creating tools that can be really integrated in design processes.

Originality/value

The paper presents an original loosely coupled method for the shape optimization of flexible wings in which recent and modern techniques are used at different levels of the global algorithm: the aerodynamic optimizer, the aeroelastic analyser, the shape parametrization and the objective function model.

Details

Engineering Computations, vol. 23 no. 4
Type: Research Article
ISSN: 0264-4401

Keywords

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Article
Publication date: 1 July 2021

Zi Kan, Daochun Li, Shiwei Zhao, Jinwu Xiang and Enlai Sha

This paper aims to assess the aeroacoustic and aerodynamic performance of a morphing airfoil with a flexible trailing edge (FTE). The objective is to make a comparison of…

Abstract

Purpose

This paper aims to assess the aeroacoustic and aerodynamic performance of a morphing airfoil with a flexible trailing edge (FTE). The objective is to make a comparison of the aerodynamic noise characteristics between the conventional airfoil with a flap and morphing airfoil and analyse the noise reduction mechanisms of the morphing airfoil.

Design/methodology/approach

The computational fluid dynamic method was used to calculate the aerodynamic coefficients of morphing airfoil and the Ffowcs-Williams and Hawking’s acoustic analogy methods were performed to predict the far-field noise of different airfoils.

Findings

Results show that compared with the conventional airfoil, the morphing airfoil can generate higher lift and lower noise, but a greater drag. Additionally, the noise caused by the one-unit lift of the morphing airfoil is significantly lower than that of the conventional airfoil. For the morphing airfoil, the shedding vortex in the trailing edge was the main noise resource. As the angle of attack (AoA) increases, the overall sound pressure level of the morphing airfoil increases significantly. With the increase of the trailing edge deflection angle, the amplitude and the period of sound pressure of the morning airfoil fluctuation increase.

Practical implications

Presented results could be very useful during designing the morphing airfoil with FTE, which has significant advantages in aerodynamic efficiency and aeroacoustic performance.

Originality/value

This paper presents the aerodynamic and aeroacoustic characteristics of the morphing airfoil. The effect of trailing edge deflection angle and AoA on morphing airfoil was investigated. In the future, using a morphing airfoil instead of a traditional flap can reduce the aircraft`s fuel consumption and noise pollution.

Details

Aircraft Engineering and Aerospace Technology, vol. 93 no. 5
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 17 August 2012

Yueheng Qiu, Weiguo Zhang, Xiaoxiong Liu and Pengxuan Zhao

The purpose of this paper is to present the research into fault detection and isolation (FDI) and evaluation of the reduction of performance after failures occurred in the…

Abstract

Purpose

The purpose of this paper is to present the research into fault detection and isolation (FDI) and evaluation of the reduction of performance after failures occurred in the flight control system (FCS) during its mission operation.

Design/methodology/approach

The FDI is accomplished via using the multiple models scheme which is developed based on the Extend Kalman Filter (EKF) algorithm. Towards this objective, the healthy mode of the FCS under different type of failures, including the control surfaces and structural, should be considered. It developed a bank of extended multiple models adaptive estimation (EMMAE) to detect and isolate the above mentioned failures in the FCS. In addition, the performances including the flight envelope, the voyage and endurance in cruising are proposed to reference and evaluate the process of mission, especially for UAV under failure conditions.

Findings

The contribution of this paper is to provide the information not only about the failures, but also considering whether the UAV can accomplish the task for the ground station.

Originality/value

The main contribution of this paper is in the areas of the structural and control surface faults researching, which are occurred in the mission procedures and emphasized the identification of those failures' magnitudes. The FDI scheme includes the performance evaluation, while the evaluation obtained through the extensive numerical simulations and saved in the offline database. As a consequence, it is more accurate and less computationally demanding while evaluating the performance.

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Article
Publication date: 30 September 2014

P. Laskaridis, V. Pachidis and P. Pilidis

The performance benefits of boundary layer ingestion (BLI) in the case of air vehicles powered by distributed propulsors have been documented and explored extensively by…

Abstract

Purpose

The performance benefits of boundary layer ingestion (BLI) in the case of air vehicles powered by distributed propulsors have been documented and explored extensively by numerous studies. Therefore, it is well known that increased inlet flow distortion due to BLI can dramatically reduce these benefits. In this context, a methodology that enables the assessment of different propulsion architectures, whilst accounting for these aerodynamic integration issues, is studied in this paper.

Design/methodology/approach

To calculate the effects of BLI-induced distortion, parametric and parallel compressor approaches have been implemented into the propulsion system analysis. The propulsion architectures study introduces the concept of thrust split between propulsors and main engines and also examines an alternative propulsor configuration. In the system analysis, optimum configurations are defined using thrust-specific fuel consumption as figure of merit.

Findings

For determined operating conditions, the system analysis found an optimum configuration for 65 per cent of thrust delivered by the propulsor array, which was attributed mainly to the influence of the propulsor’s intake losses. An alternative propulsor design, which used the ejector pump effect to re-energize the boundary layer, and avoiding the detrimental effects of BLI are also cited in this work.

Originality/value

To summarize, this paper contributes with a general review of the research that has been undertaken to tackle the aforementioned aerodynamic integration issues and, in this way, make viable the implementation of distributed propulsion systems with BLI.

Details

Aircraft Engineering and Aerospace Technology: An International Journal, vol. 86 no. 6
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 13 June 2019

Farid Shahmiri, Maryam Sargolzehi and Mohammad Ali Shahi Ashtiani

The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined…

Abstract

Purpose

The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on figure of merit (FM) and lift-to-drag responses developed for hover and forward flight, respectively.

Design/methodology/approach

The approach is to couple a general flight dynamic simulation code, previously validated in the time domain, with design of experiment (DOE) required for the response surface development. DOE includes I-optimality criteria to preselect the data and improve data acquisition process. Desirability approach is also implemented for a better understanding of the optimum rotor blade planform in both hover and forward flight.

Findings

The resulting system provides a systematic manner to examine the rotor blade design variables and their interactions, thus reducing the time and cost of designing rotor blades. The obtained results show that the blade taper ratio of 0.3, the point of taper initiation of about 0.64 R within a SC1095R8 airfoil satisfy the maximum FM of 0.73 and the maximum lift-to-drag ratio of about 5.5 in hover and forward flight.

Practical implications

The work shows the practical possibility to implement the proposed optimization process that can be used for the advanced rotor blade design.

Originality/value

The work presents the rapid and reliable optimization process efficiently used for designing advanced rotor blades in hover and forward flight.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 9
Type: Research Article
ISSN: 1748-8842

Keywords

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