Changing airline scenarios and aggravated congestion have led to an increasing demand for new regional aircraft which will cover ranges of more than 1500 nm and accommodate 80 to 130 passengers. These far exceed the capability of today's commuter types. In the required thrust range of 15-5-20 klb only a very limited number of older engines with high SFC is available. Recently the manufacturers have begun offering new turbofans based on modern core technology and internal mixing (long duct) with lower SFC.

The underwing turbofan installation aspects are well understood from earlier wide and narrowbody aircraft but installing a long-duct nacelle on such a small aircraft means a significant increase in interference problems which may lead to close coupling and new methods for minimisation of the mutual interference.

After a survey of installation problems on previous jet transports up to the Airbus family the state of the art of experimental and theoretical methods for interference analysis is discussed. By means of an inverse design method a local lofting change on a given wing design confined to 20% span area was derived in order to sustain the isobar pattern of the clean wing with the pod/pylon on.

The design modification was checked by means of a 3D-Euler code and verified on a large half-model in the NLR-HST wind tunnel. From the analysis of the resulting conclusions concerning the feasibility of such a lofting change, further improvements are made.

A new rotor model for helicopter flight mechanics simulation is presented. The rotor dynamics are described using multiblade coordinates, and the aerodynamic loads include nonlinear effects such as stall and compressibility. These loads are described as harmonic series. The number of harmonics in the series determines the model accuracy. Thus by changing the number of harmonics from one to a large number, it is possible to obtain models that range between a quasi tip path plane approach and an accurate blade element model. The user of the model can very easily change the model accuracy and consequently its efficiency. The new rotor model is investigated and its application for trim and manoeuvre calculations is presented and discussed.

A pressure instrument has been designed for the measurement of local skin friction in turbulent boundary layers on smooth and non-smooth walls. The instrument consists of three pitot tubes placed in the logarithmic mean velocity region of a turbulent boundary layer.

An instrument was built and tested in the boundary layer produced on both smooth and rough walls. The rough walls consisted of two sandpaper abrasives of sizes #40 and #24. The experiments were performed over a range of momentum thickness Reynolds numbers, Rθ, from about 5 × 103 to 18 × 103, and in a small favourable pressure gradient given by –0·0005 < (υ/ρU3τ)(dp/dx) <–0·0001. The non-dimensional equivalent sand roughness, ks+ ≡ ksUτ/υ varied from zero to approximately 85, thus spanning all three roughness regimes from smooth to fully rough. Skin friction coefficients calculated using the three tube instrument were compared to those obtained using a skin friction balance. For the smooth wall measurements, skin friction relations and Preston tubes were also used as a basis of comparison. The results were found to agree fairly well for both the smooth and non-smooth walls, and was independent of the instrument's position within the logarithmic region. However, due to measurement difficulties the scatter and uncertainty were larger for the three tube instrument than the other techniques.

Numerical smoothing of the results indicated that the agreement is expected to improve with better measurement of the pressure differences between the pitot tubes, or direct measurement of the velocity differences.

Low speed wind tunnel tests have been made to investigate the flow around a leading edge vortex flap at the maximum LID condition. Tests were also made to measure the performance of a vortex plate. The force measurements and flow visualisation tests were conducted on a 0·53 m span 60° delta wing model. Results indicate that the lift to drag ratio is a maximum for any given flap deflection angle that the flow comes smoothly onto the deflected vortex flap without forming a large leading edge separation vortex on the flap surface. The benefit of the vortex plate is seen in the drag results which are smaller than those for the datum wing. This benefit is due to leading edge suction acting on the forward facing region between the delta wing and the vortex plate.