This work deals with tailoring of adaptive material included at the roots of hingeless helicopter rotor blades to be used in individual blade control (IBC) strategies.

Usually, IBC strategies involving the use of adaptive materials either consider adaptive material embedded in the blade structure for inducing strain deformations, or apply adaptive actuators for controlling segments of the blade (e.g. for moving trailing-edge flaps).

Here, the adaptive material is used to provide augmentation of modal damping in a passive control approach, that can be conveniently tuned so as to make it the most suitable for the actual rotor configuration under examination. The presentation of a procedure for tailoring this ‘smart spring’ is the aim of the paper.

The aeroelastic blade model considered consists of a cantilever slender beam undergoing flap, lead-lag and torsional motion, coupled with a strip theory approach for the prediction of the aerodynamic loads, based on the very low frequency approximation of the pulsating-free-stream Greenberg’s theory.

Starting from this model and applying the Galërkin method, generalised mass, damping and stiffness matrices of the basic blade, as well as the incremental generalised mass, damping and stiffness matrices due to the ‘smart spring’ have been determined, the latter depending on the ‘smart spring’ inertial and elastic characteristics.

It will be shown that the application of an optimal control criterion, followed by a low frequency-approximation observer, yields the identification of the most suitable ‘smart spring’ characteristics for augmentation of rotor blade aeroelastic stability. The validity of this procedure will be demonstrated by numerical results concerning the stability analysis of two hovering blade configurations.

An experimental investigation has been conducted on the mixing layers produced from the merging of two non-parallel streams. To study the effect of velocity ratio on its development, mixing layers were produced with velocity ratios 0·7, 0·8 and 0·9. The boundary layers were untripped and initially turbulent in all the cases. For each velocity ratio, measurements were made at six streamwise locations. It was found that for velocity ratios 0·7 and 0·8, mixing layers appear to attain self-similarity but failed for 0·9 within the measurement domain. The development distance of the mixing layer flow was increased with increasing velocity ratio and the splitter wake was found to play an important role in its development. The mixing layer growth was found to decrease with increasing velocity ratio though the spread of the layer was larger at the high speed side.

This paper deals with the methods of physical modelling of Flight Control Systems (FCS) by means of Dynamically Similar free-flying Models (DSM) for the investigation of stability and controllability of aircraft at subsonic flight speeds. The subsonic flight regime allows us to avoid Mach number similarity considerations. The large scale of the DSM meets autosimilarity of Reynolds numbers, whilst Froude similarity is assured during the development and manufacturing of the DSM. This paper proves the presence of necessary and sufficient conditions of similarity for the FCS of an aircraft, and that of its DSM. The existence of necessary conditions have been proved both mathematically (by means of the π – Theorem from the theory of dimensions), and with equations involving physical quantities. The generalised scale coefficients for transitioning from the FCS’s aircraft gain factors, to those of the FCS of the DSM have been obtained, and it is shown that the coefficients depend only on the linear scale of the DSM.

The subsonic doublet lattice method (DLM) has been the industry standard method for the calculation of unsteady air loads for the past three decades. A recent comparison between the DLM and the subsonic constant pressure panel code ZONA6 suggested that the DLM lacked robustness. However, the DLM code used in the comparison was flawed and not representative of the DLM as such. Results from a contemporary DLM code are presented for the same test cases that were used in the misleading comparison. These new results show that the DLM can produce valid results for all the test cases considered. Where feasible, a comparison is made between results from the present DLM code and the published ZONA6 results.

A new semi-empirical one-dimensional approach for the prediction of aeroengine combustor emissions is presented. The model features a high level of versatility, resulting in a remarkable correlation quality of measurement data of different combustors. This combustion model was modified to reflect the design characteristics of staged combustors. After implementation in an optimisation procedure it was used for staged combustor design studies. For a typical aeroengine application an optimised combustor design was derived with respect to a predefined criteria. Potential exchange rates of combustion efficiency versus NOx emissions by rematching the fuel split for different operating conditions at extreme points of the environmental envelope (ISA/hot day/cold day) of a modern commercial aircraft were investigated. Conclusions regarding the suitable operation of the combustor were drawn. Finally the effect of changed design constrains on the optimisation result was investigated.