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Insight in the temperature distribution on the internal and external surface of the nacelle is of great importance during the design phase of an aircraft. However, detailed information is not always needed. In a preliminary project stage or during parametric optimization, short analysis times are often more crucial than high accuracy. In such cases, the global insight in the temperature levels suffices to gain understanding of the relevance and influence of certain parameters. Nevertheless, estimating the maximum temperature for the most adverse conditions should also be done before a prototype is built. Therefore, this study aims to present and compare a simplified and an advanced methodology for the analysis of engine bay cooling and ventilation systems as well as heat transfer in the nacelle in a small airplane equipped with a turboprop engine in the tractor arrangement.

Both approaches included conductive, convective and radiative heat transfer in the engine bay of the small airplane I-23 as well as heat conduction in the nacelle made of material with anisotropic thermal conductivity. The one-dimensional (1D) model assumed that the nacelle with the air flow and engine was represented by a lumped thermal model in which heat was exchanged between the different lumped segments (the nodes) and the flowing air and engine. The three-dimensional (3D) model was based on the continuous control volume approach for heat, fluid flow and thermal radiation as well as on realizable k-ε turbulence model. Both models used commercial software.

The temperature distribution at the internal and external surface of the top nacelle was calculated. The 1D model predicted a temperature per node (per segment) while the 3D model was able to determine its values accurately and find the location of hot spots. Considering the complex geometry of the engine bay and nacelle and the assumed simplification, the obtained 1D and 3D results agreed quite well.

Both models will help in the development of new ventilation and cooling systems of the engine bay and nacelle as well as in the selection of materials for parts of the top cowling in the newly redesigned airplane I-23 equipped with a turboprop engine. In addition, the methodology presented in this paper might be applied in the development of other airplanes.

The 1D and 3D models of complex heat transfer inside the engine bay and in the nacelle of the newly re-designed airplane I-23 were elaborated and compared.

This paper aims to build an accurate mathematical model which is necessary for control design and attitude estimation of a miniature unmanned rotorcraft and its subsequent conversion to an autonomous vehicle.

Frequency-domain system identification of a small-size flybar-less remote controlled helicopter is carried out based on the input–output data collected from flight tests of the instrumented vehicle. A complete six degrees of freedom quasi-steady dynamic model is derived for hover and cruise flight conditions.

The veracity of the developed model is ascertained by comparing the predicted model responses to the actual responses from flight experiments and from statistical measures. Dynamic stability analysis of the vehicle is carried out using eigenvalues and eigenvectors. The identified model represents the vehicle dynamics very well in the frequency range of interest.

The model needs to be augmented with additional terms to represent the high-frequency dynamics of the vehicle.

Control algorithms developed using the first principles model can be easily reconfigured using the identified model, because the model structure is not altered during identification.

This paper gives a practical solution for model identification and stability analysis of a small-scale flybar-less helicopter. The estimated model can be easily used in developing control algorithms.

At present work is going ahead in Holland on the construction of the very promising Infrared Astronomical Satellite IRAS. The Dutch aircraft industry has been encouraged by the great success of the first Netherlands Astronomical Satellite, ANS, to develop space‐flight technology programmes. The existence of an aerospace industry in Holland is also due to other factors such as active government support, the commercial success of the Dutch Fokker F27 and F28 aircraft, and the intensive research programmes of the astronomical departments of several Dutch universities. This article deals briefly with various developments in the industry.

Traditional dynamic and flutter analysis demands a detailed finite element model of the aircraft in terms of its mass and stiffness distribution. However, in absence of these details, modal parameters obtained from experimental tests can be used to predict the flutter characteristics of an aircraft. The purpose of this paper is to develop an improved and reliable method to predict the flutter characteristics of an aircraft structure of unknown configuration under an anticipated aerodynamic loading using software such as MSC Nastran and experimental modal parameters (such as mode shapes, natural frequencies and damping) from ground vibration tests.

A finite element model with nodes representing the test points on the aircraft is created with appropriate boundary constraints. A direct matrix abstraction program has been written for NASTRAN software that carries out a normal modes analysis and replaces the mass normalized eigenvalues and vectors with the experimentally obtained modal parameters. The flutter analysis proceeds with the solution of the flutter equation in the flutter module of NASTRAN.

The method has been evaluated for a light composite aircraft and its results have been compared with flight flutter tests and the flutter speeds obtained from the finite element model with actual stiffness and mass distributions of the aircraft.

The methodology developed helps in the realistic prediction of flutter characteristics of a structure with known geometric configuration and does not need material properties, mass or stiffness distributions. However, experimental modal parameters of each configuration of the aircraft are required for flutter speed estimation.

The proposed methodology requires experimental modal parameters of each configuration of the aircraft for flutter speed estimation.

The paper shows that an effective method to predict flutter characteristics using modal parameters from ground vibration tests has been developed.

Ground vibration testing (GVT) results can be used as system parameters for predicting flutter, which is essential for aeroelastic clearance. This paper aims to compute GVT-based flutter in time domain, using unsteady air loads by matrix polynomial approximations.

The experimental parameters, namely, frequencies and mode shapes are interpolated to build an equivalent finite element model. The unsteady aerodynamic forces extracted from MSC NASTRAN are approximated using matrix polynomial approximations. The system matrices are condensed to the required shaker location points to build an aeroelastic reduced order state space model in SIMULINK.

The computed aerodynamic forces are successfully reduced to few input locations (optimal) for flutter simulation on unknown structural system (where stiffness and mass are not known) through a case study. It is demonstrated that GVT data and the computed unsteady aerodynamic forces of a system are adequate to represent its aeroelastic behaviour.

Airforce of every nation continuously upgrades its fleet with advanced weapon systems (stores), which demands aeroelastic flutter clearance. As the original equipment manufacturers does not provide the design data (stiffness and mass) to its customers, a new methodology to build an aeroelastic system of unknown aircraft is devised.

A hybrid approach is proposed, involving GVT data to build an aeroelastic state space system, using rationally approximated air loads (matrix polynomial approximations) computed on a virtual FE model for ground flutter simulation.

Introduction This report deals with two important causes of service failures in high strength steels: stress corrosion and hydrogen embrittlement. These failure modes are often difficult to distinguish from one another fractographically, and sometimes the macroscopic and metallagraphic interpretation is also open to doubt.

The present paper aims to describe the model updating of a small aircraft dynamic finite element model (FEM) to improve its agreement with ground vibration test (GVT) data.

An automatic updating method using an optimization procedure is carried out. Instead of using dedicated updating tools, the procedure is implemented using standard MSC/NASTRAN because of wide availability of the software in small aircraft industries. The objective function is defined to minimize the differences in the natural frequency and the differences in the mode shape between the analytical model and the GVT data. Provision has been made to include the quantification of confidence in both the GVT data and in the initial model. Parameter grouping is carried out to reduce the number of design parameters during the optimization process.

The optimization module of standard finite element (FE) software can be effectively used to reduce the differences between the GVT and the FEM in terms of frequency and mode shape satisfactorily. The strategy to define the objective function based on minimizing the mode shape error can reduce the improvement in the frequency error. The required user interference can be kept low.

The most important contribution of the present paper concerns the combination of strategies to define the objective function and selection of the parameters.