Proceedings of the Institution of Mechanical Engineers,Part G: Journal of Aerospace Engineering

The paper deals with the conceptual design and sizing of a cabin escape system to be applied to a trans-atmospheric transportation system. At first, the role of suborbital vehicles towards the development of a hypersonic transportation system is presented. From this analysis, it has been clear that one of the key points in enhancing the public consensus is to demonstrate a higher level of safety and reliability with respect to the current space vehicles. Since the time of the Space Shuttle enterprise, the development of a proper escape system has been considered crucial to diminish the risk of loss of lives per mission, moving from space-like reliability characteristics to values closer to the aeronautical case. In particular, this paper presents the conceptual design of an escape system for a single stage vehicle aimed at parabolic flights. The proposed design methodology starts with the identification of the major requirements that will lead the design and sizing activities. Then, special attention is devoted to the identification of the required capabilities of a Cabin Escape System and to the selection of the proper subsystems able to guarantee these functionalities. Indeed, considering the high-level of complexity of such a system, during the design process, specific attention should be paid to the impact of on-board systems integration on the overall transportation system architecture and layout. At this purpose, a proper utilization of CAD models can ease the integration process allowing fast verification of mass and volume budgets as well as integrated simulation techniques could be useful. Furthermore, the possibility of exploiting this system during the different phases of the mission should be properly evaluated and, eventually, a preliminary impact risk analysis is reported.

Aircraft development programmes generally involve collaboration in engineering between different organisations, in order to develop innovative products efficiently, to involve necessary skills from the supply chain, and to spread risks and costs among the partners. The size and complexity of the programmes, the market demands and the contexts of competition all require the collaboration to be effective and efficient. Advances in information technology provide new capabilities to support collaborative design but a step change is needed to harness and coordinate this support to be effective and efficient. The extended enterprises in which the collaborative engineering activities take place span the partner organisations. Engineers wishing to cooperate are however facing security constraints. For example, technical security measures such as firewalls and proxy servers hamper smooth exchange of engineering data and seamless execution of collaborative workflows. The restrictions assist organisations in protecting their assets and in remaining compliant with legislation and regulations. From a programme technical point of view, effective and efficient collaboration in this world full of security and the resulting connectivity constraints is a major challenge. This paper describes the usefulness, necessity and challenges of collaboration between multi-disciplinary specialists in aerospace engineering. It presents the ‘Brics’ technology that supports the realisation of cross-organisation collaborative workflows. The technology supports aircraft manufacturers and their supply chains in facing the challenges and in performing analyses of innovative aircraft designs collaboratively. This technology has emerged from past research projects, and has been further developed and successfully demonstrated in the Thermal Overall Integrated Conception of Aircraft project, a research and technology development project carried out in the Seventh Framework Programme funded by the European Union. The developed technology is illustrated in the context of a multi-partner analysis and optimisation study, which has been conducted as part of a pylon design that is subject to thermal constraints.

The main goal of this paper is to investigate feasibility of using all-electric propulsion system for a mid-light business jet aircraft in the near future (20–30 years from now). The secondary goal is to assess the impact of using such system on operating costs and emission reduction. This paper presents calculations of business jet aircraft mission energy demands and compares them with batteries capabilities. Three different types of lithium batteries are investigated in terms of their energy densities projected for three different time frames. Mass of batteries that is required to provide demanded amount of energy to perform the mission is compared with the maximum mass of fuel that the baseline aircraft is able to take. On this basis, the feasibility of all-electric propulsion system is assessed. Additionally, in order to show the limitations of such system, maximum range is calculated for the mass of batteries that would potentially enable to perform the flight. Furthermore, CO₂ and NOx emission of the baseline aircraft engines are compared with the amount of gaseous pollutants which are emitted by the power plant, when energy needed to recharge batteries is being produced. Finally, the potential fuel cost reduction is calculated based on the cost of electricity that would be used to recharge batteries.

Landing gear weight calculations can be carried out using statistical or analytical methods. Statistical methods were used in the past and offered quick group weights. However, they are not capable of computing accurately the weight of landing gears, which have special geometries and performance. In this work, landing gear weight is computed using analytical methods based on parametric 3D models. The procedure established by Kraus and Wille is applied as a baseline so as to create a procedure capable of dealing with landing gear weight calculations. This method is designed to be as flexible as possible, giving the user the freedom to modify many options and parameters and integrate landing gear design into Robust Aircraft Parametric Interactive Design.

This paper discusses the problems of designing a fuel installation for the I-31T, four seat, fully composite aircraft designed in accordance with FAR 23/CS 23. The project was carried out under the European Union project entitled ESPOSA (7^th Frame Programme)—“Efficient Systems and Propulsion for Small Aircraft”. I-31 Turbo has been created on the base frame of the I-23 “Manager” aircraft, through replacement of its original piston engine, the Lycoming O-360 A1A, with a TP-100 turbo shaft engine. Moreover, the front of the plane, from the bulkhead 1, was altered, while the old navigation and piloting equipment was substituted with a modern “glass cockpit”. These modifications required designing of a new fuel system within the engine chamber. The existing fuel system in the fuselage and wings was adapted for the new JET-A fuel. In order to confirm whether the installation in the wings and the fuselage would provide the consumption rate required by the TP-100 engine (a tripling of the fuel flow and a ten-fold increase in kinematic viscosity of fuel), calculations and simulations were carried out. The calculations were verified using the second I-23 aircraft, designated for static campaign tests. The ground tests studied the flow of the fuel from the auxiliary tank and were carried out under two different fuel temperatures (−20 ℃ and 20 ℃) and two different fuselage angles (0 and −10.7°). The project was documented in 3D in CATIA V5 program and in project documentation. To sum up, the main innovative element of the fuel system in the engine chamber is the module containing the pumps, filters, one-way valves, flow meter, among others. This concentration of parts allows a significant reduction in plane servicing time in between flights.

The aim of this paper is to present the results of prepreg nacelles design and manufacturing for I-31T aircraft. The work was part of the European Union project ‘Efficient Systems and Propulsion for Small Aircraft’ (ESPOSA). The new engine cowling design was preceded by computational fluid dynamics numerical analyses stage. Use of TP100 engine was assumed, which was installed on the I-31T aircraft in the framework of the ESPOSA project. This article presents the process of external geometry adjustment, chosen results of numerical analyses, 3D model design, manufacturing process and tests results. The new cowlings underwent ground and in-flight tests. Monitoring during the test included external and internal cowling temperatures. The collected test data were further analysed. Throughout accumulated energy in hot parts of the engine the temperature inside the engine nacelles rises. ‘Hot’ composite nacelles were used in order to withstand the impact of high temperature. In majority of modern aircraft propulsion systems, cowlings are made of composite. Currently, there is a trend to manufacture initially solution treated materials – i.e. glass and carbon fiber pre-pregs in out-of-autoclave process. Such an approach allows to accelerate the fabric layup process and achieve highly repeatable structure. Materials that are currently available in the market allow to manufacture the cowlings and simultaneous weight minimization. Owing to the applied 3D software for both design and manufacture of machining tools, it is possible to shorten the time of manufacturing a complete element. The cowlings were designed in cooperation with the NLR – Netherlands Aerospace Centre, who was a partner in the ESPOSA project and has experience in prepreg elements design. NLR was responsible for material selection and final part manufacturing. Tooling has been designed and manufactured in the Institute of Aviation in Warsaw. The assembly works were carried out by Zakłady lotnicze Margański & Mysłowski.

Conceptual and preliminary designs of future aircraft have become increasingly complex due to the enlargement of the basic criteria for evaluating emerging solutions. In the past, the basic performance characteristics of an airplane were the only selection criteria. Today, more and more emphasis is placed on factors such as impact on the environment, cost-effectiveness, or comfort of travel. The method for optimization of the key parameters of a small aircraft for use in the initial phase of a project is presented in this paper. It takes into account the requirements of aviation safety imposed by the European Union certification specifications CS-23 and requirements of aircraft competitiveness. Requirements and design assumptions were formulated based on the concept of the Small Air Transport system (SATs). The method is based on the multidisciplinary design optimization and covers the basic areas related to the design of aircraft: aerodynamics, aircraft structure, performance and expected operating costs. The objective function is defined as the value of the direct operating cost per passenger-kilometre. An evolutionary algorithm was applied to solve the optimization problem. As an example of the use of this method, the optimization of design parameters of the two classes of aircraft, 9-seater and 19-seater was carried out. The results were compared with the parameters of aircraft, which are in service. Sensitivity analysis of the objective function with respect to selected parameters of the aircraft was also made. The analysis allowed to select the most important parameters responsible for the operational costs.

This paper presents a design approach for large passenger aircraft nacelles. It describes nacelle functions, configurations, and major components. Design criteria for particular components of the nacelle are presented. Typical materials used for structure build including composites are elaborated. Certification requirements are shown. The paper presents challenges for nacelle design and build, including weight, aerodynamic smoothness, nacelle integration, and cost.

This paper presents the results of a numerical study of the aerodynamic shape of the Rocket Plane LEX. The Rocket Plane is a main part of the Modular Airplane System – MAS; a special vehicle devoted to suborbital tourist flights. The Rocket Plane was designed for subsonic and supersonic flight conditions. Therefore, the impact of the Mach number should be considered during the aerodynamic design of the Rocket Plane. The main goal of the investigation was to determine the sensitivity of the Rocket Plane aerodynamic characteristics to the Mach number during the optimisation of the LEX geometry. The paper includes results of the optimisation process for Mach number from the range Ma = 0.5 to Ma = 2.5. These results reveal that the aerodynamic characteristics of models optimised for the subsonic and transonic regime of Mach numbers (up to Ma = 1) were also improved for the supersonic speed regime. However, in the case of models optimised for the supersonic flight regime the aerodynamic characteristics in subsonic flight regime, are inferior compared to the model before the optimisation process.

The use of a hybrid powertrain for a conventional single main rotor helicopter is investigated, with the objective of assessing its feasibility and its potential impact on improving safety, especially for single-engine rotorcraft. The study is focused on the characteristics of the powertrain and required battery pack. It is based on a simple analysis of power required in forward flight and the estimate of the total energy required for a powered landing manoeuvre after thermal engine failure. Current technologies are considered as well as expected improvements, especially as far as energy density and power density of the battery are concerned. The latter analysis is based on current trends for battery and motors technologies, in order to determine the technological breakthrough limit.

Optimized electric motors are lighter and smaller than conventional piston engines. As a result, new airplane configurations are feasible as motors can be placed in unconventional positions. Through careful aircraft design higher aerodynamic efficiencies of airframe, propeller, and propeller integration can be achieved. The energy density of current batteries, however, still limits strongly the range of purely battery powered aircraft. But if the energy is stored in liquid fuel and converted by a generator into electric energy, then the advantages of electric propelled airplanes and conventional combustion engines can be combined. But which combustion engine is optimal for such a serial-hybrid electric aircraft? In this new propulsion chain, other boundary conditions apply to the combustion engine than in conventional aircraft designs. These boundary conditions interact with the characteristics of combustion engines. An example for an engine characteristic is that different kinds of piston engines exist. It can be observed that technologies, which result in lighter piston engines, are associated with lower efficiencies and vice versa. In this paper it will be shown through considerations on aircraft level, that the optimal combustion engine for an electric-hybrid airplane should be heavier and more efficient than the optimal combustion engine for a conventional aircraft.

Nowadays, optimization is a very popular tool used to improve existing projects. The optimization covers different disciplines by linking them into multidisciplinary process of design. Existing software tools allow to very effectively solve particular problems giving high quality solutions which were previously very hard to achieve. Aeronautical engineering is a domain/field which links many disciplines: aerodynamics, stability, control, structural analysis, materials, propulsion systems, avionics, etc. Therefore, the multidisciplinary optimization results in very significant progress not only in aircraft design but also in air transport, which links technical aspects with economical questions. The paper presents selected aspects of using the multidisciplinary optimization in aeronautical engineering with special focus on multidisciplinary aircraft design.