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Article
Publication date: 19 October 2018

Cezary Gorniak, Zdobyslaw Jan Goraj and Bartosz Olszanski

The purpose of this research is a preliminary selection of wing section, which would be the best suited for PW-100 – a MALE class UAV of 600 kg weight. PW-100 will be used…

Abstract

Purpose

The purpose of this research is a preliminary selection of wing section, which would be the best suited for PW-100 – a MALE class UAV of 600 kg weight. PW-100 will be used as a testing platform in different institutions such as research institutes, industry research centers or universities of technology (phase 1) to enable the in-flight testing of various on-board systems (mobile radars, thermovision sensors, chemical sensors, antennas, teledetection systems and others). Untypical layout of PW-100 resulted from the plans of further development of this configuration for a military application.

Design/methodology/approach

Important role in the research described in this paper plays the selection of main wing section to fulfil the preliminary requirements regarding maximum lift coefficient, minimum drag, aerodynamic efficiency etc. Two different wing sections (R1082 and SA19) were tested in wind tunnel, both with flaps deflected at the range of 0°-30°. Experimental measurements were performed in the low turbulence wind tunnel with closed test section of 45 cm × 35 cm. Numerical simulations of the flow around the wing sections were performed using MSES code. Boundary conditions were assumed basing on the typical mission of PW-100 for flight altitude around 9,000 m, speed of 110 km/h what results in Re = 956,000.

Findings

Lift coefficients obtained from both experimental and numerical methods for single slatted airfoil SA19 are much higher than that of get for Ronch R1082 airfoil. PW-100 aircraft with SA19 airfoils will be able to be trimmed and fly at any altitude up to 9,000 m and with an arbitrary weight up to 600 kg. Aerodynamic characteristics of SA19 remain smoother and more predictable than that of R1082 airfoil. The very promising properties of SA19 airfoil are well known to the authors since the beginning of last decade when PW team worked together with IAI team on CAPECON project and now it was fully confirmed by this research.

Practical implications

It was confirmed that selection of the proper wing section for the special mission performed by UAV is of the highest importance decision to be taken at the preliminary design phase. Because there is a limited access to the base of technical parameters in many different UAVs classes and the classical analysis of trends cannot be fully applied, the wing section analysis, either experimental or numerical, must be performed. The situation is much worse than in the case of manned aircrafts because most of the modern UAVs are made of carbon or glass fiber, and therefore, there is no chance for analysis of trends.

Originality/value

This paper presents a very efficient method of assessing the influence of wing section on aircraft performance adopted for MALE class UAV, especially in an early stage of preliminary design process. The assessment is built mainly on three requirements: Maximum 2D lift coefficient for take-off configuration with flap deflected on 20 degrees should be greater than 2.4. Endurance factor CL1.5/CD for loitering conditions (Ma = 0.5 and CD0 = 0.008) should be greater than 110. The relative wing section thickness should be greater/equal than 19 per cent (it is required for high volume fuel tank located in the wings).

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 2
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 18 December 2018

Adam Tomaszewski and Zdobyslaw Jan Goraj

The purpose of this paper is to present an approach to a polar graph measurement by a flight testing technique and to propose a baseline research method for future tests…

Abstract

Purpose

The purpose of this paper is to present an approach to a polar graph measurement by a flight testing technique and to propose a baseline research method for future tests of UAV polar graphs. The method presented can be used to demonstrate a conceptual and preliminary design process using a scaled, unmanned configuration. This shows how results of experimental flight tests using a scaled flying airframe may be described and analysed before manufacturing the full scale aircraft.

Design/methodology/approach

During the research, the flight tests were conducted for two aerodynamic configurations of a small UAV. This allowed the investigation of the influence of winglets and classic vertical stabilizers on the platform stability, performance and therefore polar graphs of a small unmanned aircraft.

Findings

A methodology of flight tests for the assessment of a small UAV’s polar graph has been proposed, performed and assessed. Two aerodynamic configurations were tested, and it was found that directional stability had a large influence on the UAV’s performance. A correlation between the speed and inclination of the altitude graph was found – i.e. the higher the flight speed, the steeper the altitude graph (higher descent speed, steeper flight path angle). This could be considered as a basic verification that the recorded data have a physical sense.

Practical implications

The polar graph and therefore glide ratio of the aircraft is a major factor for determining its performance and power required for flight. Using the right flight test procedure can speed-up the process of measuring glide ratio, making it easier, faster, robust, more effective and accurate in future research of novel, especially unorthodox configurations. This paper also can be useful for the proper selection of requirements and preliminary design parameters for making the design process more economically effective.

Originality/value

This paper presents a very efficient method of assessing the design parameters of UAVs, especially the polar graph, in an early stage of the design process. Aircraft designers and producers have been widely performing flight testing for years. However, these procedures and practical customs are usually not wide spread and very often are treated as the company’s “know how”. Results presented in this paper are original, relatively easily be repeated and checked. They may be used either by professionals, highly motivated individuals and representatives of small companies or also by ambitious amateurs.

Details

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

Keywords

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Article
Publication date: 8 October 2018

Witold Artur Klimczyk and Zdobyslaw Jan Goraj

The purpose of this paper is to present a method for analysis and optimization of morphing wing. Moreover, a numerical advantage of morphing airfoil wing, typically…

Abstract

Purpose

The purpose of this paper is to present a method for analysis and optimization of morphing wing. Moreover, a numerical advantage of morphing airfoil wing, typically assessed in simplified two-dimensional analysis is found using higher fidelity methods.

Design/methodology/approach

Because of multi-point nature of morphing wing optimization, an approach for optimization by analysis is presented. Starting from naïve parametrization, multi-fidelity aerodynamic data are used to construct response surface model. From the model, many significant information are extracted related to parameters effect on objective; hence, design sensitivity and, ultimately, optimal solution can be found.

Findings

The method was tested on benchmark problem, with some easy-to-predict results. All of them were confirmed, along with additional information on morphing trailing edge wings. It was found that wing with morphing trailing edge has around 10 per cent lower drag for the same lift requirement when compared to conventional design.

Practical implications

It is demonstrated that providing a smooth surface on wing gives substantial improvement in multi-purpose aircrafts. Details on how this is achieved are described. The metodology and results presented in current paper can be used in further development of morphing wing.

Originality/value

Most of literature describing morphing airfoil design, optimization or calculations, performs only 2D analysis. Furthermore, the comparison is often based on low-fidelity aerodynamic models. This paper uses 3D, multi-fidelity aerodynamic models. The results confirm that this approach reveals information unavailable with simplified models.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 3
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 6 November 2018

Kamila Kustron, Vaclav Horak, Radek Doubrava and Zdobyslaw Jan Goraj

The risk of hail-impact occurrence that can decrease local strength property must be taken into account in the design of primary airframe structures in aviation, energy…

Abstract

Purpose

The risk of hail-impact occurrence that can decrease local strength property must be taken into account in the design of primary airframe structures in aviation, energy and space industries. Because of the high-speed of hail impact in operation, it can affect the load carrying capacity. Testing all impact scenarios onto real structure is expensive and impractical. The purpose of this paper is to present a cost-effective hybrid testing regime including experimental tests and FEM-based simulations for airframe parts that are locally exposed to the impacting hail in flight.

Design/methodology/approach

Tested samples (specimens) are flat panels of laminated and sandwich carbon/epoxy composites that are used in designing lightweight new airframes. The presented numerical simulations provide a cost effective and convenient tool for investigating the hail impact scenarios in the design process. The smoothed particle hydrodynamics (SPH) technique was selected for the simulation of projectiles. The most commonly used shape of projectiles in hail impact tests is the ice ball with a defined diameter. The proposed simulation technique was verified and validated in tests on flat composite panels (specimens).

Findings

Integration of the numerical analyses with high-speed impact tests of hail onto flat laminated and sandwich composite shells has been presented, and a developed simulation model for impact results assessment was obtained.

Originality/value

The tested coupons (specimens) are flat panels as representative of structural design deployed in real aircraft structures. These numerical simulations provide a cost effective and convenient tool for hail impact scenarios in the design process.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 3
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 2 October 2018

Ewa Marcinkiewicz, Zdobyslaw Jan Goraj and Marcin Figat

The purpose of this paper is to describe an integrated approach to spin analysis based on 6-DOF (degrees of freedom) fully nonlinear equations of motion and a…

Abstract

Purpose

The purpose of this paper is to describe an integrated approach to spin analysis based on 6-DOF (degrees of freedom) fully nonlinear equations of motion and a three-dimensional multigrid Euler method used to specify a flow model. Another purpose of this study is to investigate military trainer performance during a developed phase of a deliberately executed spin, and to predict an aircraft tendency while entering a spin and its response to control surface deflections needed for recovery.

Design/methodology/approach

To assess spin properties, the calculations of aerodynamic characteristics were performed through an angle-of-attack range of −30 degrees to +50 degrees and a sideslip-angle range of −30 degrees to +30 degrees. Then, dynamic equations of motion of a rigid aircraft together with aerodynamic loads being premised on stability derivatives concept were numerically integrated. Finally, the examination of light turboprop dynamic behaviour in post-stalling conditions was carried out.

Findings

The computational method used to evaluate spin was positively verified by comparing it with the experimental outcome. Moreover, the Euler code-based approach to lay down aerodynamics could be considered as reliable to provide high angles-of-attack characteristics. Conclusions incorporate the results of a comparative analysis focusing especially on comprehensive assessment of output data quality in relation to flight tests.

Originality/value

The conducted calculations take into account aerodynamic and flight dynamic interaction of an aerobatic-category turboprop in spin conditions. A number of manoeuvres considering different aircraft configurations were simulated. The computational outcomes were subsequently compared to the results of in-flight tests and the collected data were thoroughly analysed to draw final conclusions.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 3
Type: Research Article
ISSN: 1748-8842

Keywords

Content available
Article
Publication date: 13 November 2018

Professor Zdobyslaw Jan Goraj

Abstract

Details

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

Content available
Article
Publication date: 4 September 2017

Zdobyslaw Jan Goraj

Abstract

Details

Aircraft Engineering and Aerospace Technology, vol. 89 no. 5
Type: Research Article
ISSN: 1748-8842

Content available
Article
Publication date: 13 March 2019

Zdobyslaw Jan Goraj

Abstract

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 3
Type: Research Article
ISSN: 1748-8842

Content available
Article
Publication date: 22 March 2021

Mariusz Kowalski, Zdobyslaw Jan Goraj and Bartłomiej Goliszek

The purpose of this paper is to present the result of calculations that were performed to estimate the structural weight of the passenger aircraft using novel…

Abstract

Purpose

The purpose of this paper is to present the result of calculations that were performed to estimate the structural weight of the passenger aircraft using novel technological solution. Mass penalty resulting from the installation of the fuselage boundary layer ingestion device was needed in the CENTRELINE project to be able to estimate the real benefits of the applied technology.

Design/methodology/approach

This paper focusses on the finite element analysis (FEA) of the fuselage and wing primary load-carrying structures. Masses obtained in these analyses were used as an input for the total structural mass calculation based on semi-empirical equations.

Findings

Combining FEA with semi-empirical equations makes it possible to estimate the mass of structures at an early technology readiness level and gives the possibility of obtaining more accurate results than those obtained using only empirical formulas. The applied methodology allows estimating the mass in case of using unusual structural solutions, which are not covered by formulas available in the literature.

Practical implications

Accurate structural mass estimation is possible at an earlier design stage of the project based on the presented methodology, which allows for easier and less costly changes in designed aircrafts.

Originality/value

The presented methodology is an original method of mass estimation based on a two-track approach. The analytical formulas available in the literature have worked well for aeroplanes of conventional design, but thanks to the connection with FEA presented in this paper, it is possible to estimate the structure mass of aeroplanes using unconventional technological solutions.

Details

Aircraft Engineering and Aerospace Technology, vol. ahead-of-print no. ahead-of-print
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 25 January 2019

Zdobyslaw Jan Goraj, Mariusz Kowalski and Bartlomiej Goliszek

This paper aims to present the results of calculations that checked how the longerons and frames arrangement affects the stiffness of a conventional structure. The paper…

Abstract

Purpose

This paper aims to present the results of calculations that checked how the longerons and frames arrangement affects the stiffness of a conventional structure. The paper focuses only on first stage of research – analysis of small displacement. Main goal was to compare different structures under static loads. These results are also compared with the results obtained for a geodetic structure fuselage model of the same dimensions subjected to the same internal and external loads.

Design/methodology/approach

The finite element method analysis was carried out for a section of the fuselage with a diameter of 6.3 m and a length equal to 10 m. A conventional and lattice structure – known as geodetic – was used.

Findings

Finite element analyses of the fuselage model with conventional and geodetic structures showed that with comparable stiffness, the weight of the geodetic fuselage is almost 20 per cent lower than that of the conventional one.

Research limitations/implications

This analysis is limited to small displacements, as the linear version of finite element method was used. Research and articles planned for the future will focus on nonlinear finite element method (FEM) analysis such as buckling, structure stability and limit cycles.

Practical implications

The increasing maturity of composite structures manufacturing technology offers great opportunities for aircraft designers. The use of carbon fibers with advanced resin systems and application of the geodetic fuselage concept gives the opportunity to obtain advanced structures with excellent mechanical properties and low weight.

Originality/value

This paper presents very efficient method of assessing and comparison of the stiffness and weight of geodetic and conventional fuselage structure. Geodetic fuselage design in combination with advanced composite materials yields an additional fuselage weight reduction of approximately 10 per cent. The additional weight reduction is achieved by reducing the number of rivets needed for joining the elements. A fuselage with a geodetic structure compared to the classic fuselage with the same outer diameter has a larger inner diameter, which gives a larger usable space in the cabin. The approach applied in this paper consisting in analyzing of main parameters of geodetic structure (hoop ribs, helical ribs and angle between the helical ribs) on fuselage stiffness and weight is original.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 6
Type: Research Article
ISSN: 1748-8842

Keywords

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