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The purpose of this paper is to present the results of the preliminary design and optimization of the air-intake system and the engine nacelle. The work was conducted as part of an integration process of a turboprop engine in a small aircraft in a tractor configuration.

The preliminary design process was performed using a parametric, interactive design approach. The parametric model of the aircraft was developed using the PARADES™ in-house software. The model assumed a high level of freedom concerning shaping all the components of aircraft important from the point of view of the engine integration. Additionally, the software was used to control the fulfillment of design constraints and to analyze selected geometrical properties. Based on the developed parametric model, the preliminary design was conducted using the interactive design and optimization methodology. Several concepts of the engine integration were investigated in the process. All components of the aircraft propulsion system were designed simultaneously to ensure their compliance with each other.

The concepts of the engine integration were modified according to changes in the design and technological constraints in the preliminary design process. For the most promising configurations, computational fluid dynamics (CFD) computations were conducted using commercial Reynolds-averaged Navier–Stokes solver FLUENT™ (ANSYS). The simulations tested the flow around the nacelle and inside the air-delivery system which consists of the air-intake duct, the foreign-particles separator and the auxiliary ducts delivering air to the cooling and air-conditioning systems. The effect of the working propeller was modeled using the Virtual Blade Model implemented in the FLUENT code. The flow inside the air-intake system was analyzed from the point of view of minimization of pressure losses in the air-intake duct, the quality of air stream delivered to the engine compressor and the effectiveness of the foreign particles separator.

Based on results of the CFD analyses, the final concept of the turboprop engine integration has been chosen.

The presented results of preliminary design process are valuable to achieve the final goal in the ongoing project.

This paper aims to set up and assess a new method to collaboratively mature the requirements for engine development in a more efficient way during the preliminary design phase.

A collaborative process has been set up in which detailed information on the behaviour of designed engines has been integrated into the aircraft preliminary sizing process by means of surrogate modelling.

The engine surrogate model has been invoked as a black box from within the aircraft preliminary design optimisation loops. The surrogate model reduces the uncertainty of coarse-grain formulas and may result in more competitive aircraft and engine designs. The surrogate model has been integrated in a collaborative cross-organisational workflow between aircraft manufacturer, engine manufacturer and simulation service providers to prepare for its deployment in industrial preliminary design processes.

The new collaborative way of working between aircraft manufacturer, engine manufacturer and simulation service providers could contribute to remove time consuming rework cycles in early and later design stages within delivering the optimal aircraft-engine combination.

The assessed process, based on an innovative collaboration standard, provides the opportunity to introduce useful design iterations with much more enriched information than in the classical design process as performed today. Specifically, the application of an engine surrogate model is advantageous, as it allows for extensive trade-off studies on aircraft level because of the low computational effort, while the intellectual property of the engine manufacturer (the engine preliminary design process) is respected and kept in-house.

The purpose of this paper is to present a multi‐objective and multi‐point optimization method to support the preliminary design of an unmixed turbofan mounted on a sample UAV/UCAV aircraft.

An in‐house multi‐objective evolutionary algorithm, a flight simulator and a validated engine simulator are implemented and joined together using object‐oriented programming.

Optimal values are found of the pressure ratio and corrected mass flow of both the engine fan and compressor as they operate in on/off design conditions (multipoint approach), as well as the engine by‐pass ratio, that contextually minimize time and engine fuel consumption required to cover a fixed trajectory (mission profile). Furthermore, the optimal distribution of the thermodynamic quantities along the trajectory is determined.

The research deals with a preliminary design of an engine, therefore no detailed engine geometry can be found.

The paper shows how a multiobjective and multipoint approach to the design of an engine can affect the choice of the engine architecture. In particular, major practical implications regard how the mission profile can affect the choice of the design point: in fact, there is no longer a definitive design point but the design of a UAV/UCAV should be addressed as a function of the mission profile.

The paper presents a multiobjective and multipoint approach to engine optimization as a function of the mission profile.

The aim of this work was to focus on the creation of a mathematical model for the optimization problem and the methodology for selection of the most important parameters. The engine mount design optimization problem requires a selection of the most important parameters, which influence the final solution.

A two-tier optimization algorithm was considered. The geometry of the engine mount is optimized by first taking into account the sum of loads, and the weight is minimized next.

Several parameters were selected as the essential set of factors for the engine mount design and for optimization to get the minimal weight for the system of the engine and engine mount.

As a case study, the single turboprop engine attached to the front part of fuselage – the so-called tractor propeller configuration – was considered.

This paper contains the proposed methodology, which is going to be used in the efficient systems and propulsion for small aircraft (ESPOSA) project to optimize the newly designed engine mount.

This paper presents a new developed methodology that could be useful for engine mount design.

This paper aims to investigate the cycle performance of a small size turbojet engine used in unmanned aerial vehicles at 0–5,000 m altitude and 0–0.8 Mach flight speeds with real component maps.

The engine performance calculations were performed for both on-design and off-design conditions through an in-house code generated for simulating the performance of turbojet engines at different flight regimes. These calculations rely on input parameters in which fuel composition are obtained through laboratory elemental analysis.

Exemplarily, according to comparative results between in-house developed performance code and commercially available software, there is 0.25% of the difference in thrust value at on-design conditions.

Once the on-design performance parameters and fluid properties were determined, the off-design operation calculations were performed based on the compressor and turbine maps and scaling methodology. This method enables predicting component maps and fitting them to real conditions.

A method to be used easily by researchers on turbojet engine performance calculations which best fits to real conditions.

This paper aims to use the crankshaft-bearing system of a four-cylinder internal combustion engine as the studying object, and develop a multi-objective optimization design of the crankshaft-bearing. In the current optimization design of engine crankshaft-bearing, only the crankshaft-bearing was considered as the studying object. However, the corresponding relations of major structure dimensions exist between the crankshaft and the crankshaft-bearing in internal combustion engine, and there are the interaction effects between the crankshaft and the crankshaft-bearing during the operation of internal combustion engine.

The crankshaft mass and the total frictional power loss of crankshaft-bearing s are selected as the objective functions in the optimization design of crankshaft-bearing. The Particle Swarm Optimization algorithm based on the idea of decreasing strategy of inertia weight with the exponential type is used in the optimization calculation.

The total frictional power loss of crankshaft-bearing and the crankshaft mass are decreased, respectively, by 26.2 and 5.3 per cent by the multi-objective optimization design of crankshaft-bearing, which are more reasonable than the ones of single-objective optimization design in which only the crankshaft-bearing is considered as the studying object.

The crankshaft-bearing system of a four-cylinder internal combustion engine is taken as the studying object, and the multi-objective optimization design of crankshaft-bearing based on the crankshaft-bearing system is developed. The results of this paper are helpful to the design of the crankshaft-bearing for engine. There is universal significance to research the multi-objective optimization design of crankshaft-bearing based on the crankshaft-bearing system. The research method of the multi-objective optimization design of crankshaft-bearing based on the crankshaft-bearing system can be used to the optimization design of the bearing in the shaft-bearing system of ordinary machinery.

An aircraft engine control system consists of a large scale of control parameters and variables because of the complex structure of aero-engine. Monitoring and adjusting control variables and parameters such as detecting, isolating and reconfiguring the system faults/failures depend on the controller design. Developing a robust controller is based on an accurate mathematical model.

In this study, a small-scale turboprop engine is modeled. Simulation is carried out on MATLAB/Simulink for design and off-design operating conditions. Both steady-state and transient conditions (from idle to maximum thrust levels) are tested. The performance parameters of compressor and turbine components are predicted via trained Neuro-Fuzzy model (ANFIS) based on component maps. Temperature, rotational speed, mass flow, pressure and other parameters are generated by using thermodynamic formulas and conservation laws. Considering these calculated values, error calculations are made and compared with the cycle data of the engine at the related simulation conditions.

Simulation results show that the designed engine model’s simulation values have acceptable accuracy for both design and off-design conditions from idle to maximum power operating envelope considering cycle data. The designed engine model can be adapted to other types of gas turbine engines.

Different from other literature studies, in this work, a small-scale turboprop engine is modeled. Furthermore, for performance prediction of compressor and turbine components, ANFIS structure is applied.

A combustor design is a particularly important and difficult task in the development of gas turbine engines. During studies for accurate and easy combustor design, reasonable design methodologies have been established and used in engine development. The purpose of this paper is to review the design methodology for combustor in development of advanced gas turbine engines. The advanced combustor development task can be successfully achieved in less time and at lower cost by adopting new and superior design methodologies.

The review considers the main technical problems (combustion, cooling, fuel injection and ignition technology) in the development of modern combustor design and deals with combustor design methods by dividing it into preliminary design, performance evaluation, optimization and experiment. The advanced combustion and cooling technologies mainly used in combustor design are mentioned in detail. In accordance with the modern combustor design method, the design mechanisms are considered and the methods used in every stage of the design are reviewed technically.

The improved performances and strict emission limits of gas turbine engines require the application of advanced technologies when designing combustors. The optimized design mechanism and reasonable performance evaluation methods are very important in reducing experiments and increasing the effectiveness of the design.

This paper provides a comprehensive review of the design methodology for the advanced gas turbine engine combustor.

AIRCRAFT production is so much in everyone's mind at the present time that inevitably much discussion has arisen around its relationship to design. Schools of thought range from that which would subjugate design entirely to production to that which puts design for maximum performance before all other considerations. It is intended here to review the design and production of current types of aero‐engines primarily from the engine designer's point of view.

IN all cases, the estimation of airscrew‐engine performance must start with a clear specification of the primary condition for which the airscrew is to be designed. In the majority of cases up to date this primary design condition has been that the airscrew is to absorb the maximum b.h.p. of the engine at maximum permissible r.p.m. at maximum boost height and maximum aircraft speed. (The aircraft speed is obviously an estimate only, but we may assume for the time being that it is reliable.) Under these conditions the airscrew efficiency is required to be the maximum possible.