Advanced design offices have traditionally applied conceptual design techniques based on semi-empirical methods in an attempt to develop an accurate prediction of aircraft designs at the end of the development process. Continuing advances in computer capability and rotorcraft analysis software present an opportunity to re-think conceptual design to include the greater use of physics-based analyses. A roadmap for developing this capability is outlined, taking into account techniques and ideas from Model-Based Systems Engineering, Design Thinking and Multidisciplinary Optimisation. Recent activities that demonstrate some of these desired capabilities are briefly described along with lessons learned.

There is a wide variety of CFD grid types including Cartesian, structured, unstructured and hybrids, as well as, numerous methodologies of combining these to reduce the time required to generate high-quality grids around complex configurations. If the grid methodologies were implemented in different codes, they should be written in such a way as to obtain the maximum performance from the available computer resources. A common interface should also be required to allow for ease of use. However, it is very time consuming to develop, maintain and add extra functionally to different codes. This paper examines the possibility of taking an existing CFD solver, the Helicopter Multi-Block (HMB) CFD method, and implementing a new grid type while reusing as much as possible the original code base. The paper presents some of the challenges encountered in extending the code which was written for a single mesh type, to a more flexible solver that is still computationally efficient but can cope with a variety of grid types.

Turbulent airflow in a supersonic intake model of rectangular cross-section is studied numerically. Instability of shock waves formed in the intake and ahead of the entrance is examined. Solutions of the Reynolds-averaged Navier–Stokes equations are obtained with a finite-volume solver of second-order accuracy. The expulsion and the swallowing of the shocks with a variation of free-stream parameters are studied at both subsonic and supersonic conditions prescribed at the outlet. Hysteresis in the dependence of 2D flow on the free-stream velocity and angle-of-attack is documented. An influence of 3D effects on the flow is examined.

A novel vortex generator design positioned upstream of a normal shock followed by a subsequent diffuser was investigated using large eddy simulations. In particular, “ramped-vane” flow control devices with three different heights relative to the incoming boundary layer thickness (0.34δ, 0.52δ and 0·75δ) were placed in a supersonic boundary layer with a freestream Mach number of 1.3 and a Reynolds number of 2400 based on the momentum thickness. This is the first numerical study to investigate the size effect of the ramped-vane for flow control device in terms of shape factor, flow separation and flow unsteadiness. The results showed that these devices generated strong streamwise vortices that entrained high-momentum fluid to the near-wall region and increased turbulent mixing. The devices also decreased shock-induced flow separation, which resulted in a higher downstream skin friction in the diffuser. In general, the largest ramped-vane (0.75δ) produced the largest reductions in flow separation, shape factor and overall unsteadiness. These results and a careful review of the literature study also determined the quantitative correlation of optimum VG height with Mach number, whereby h/δ~1 is often optimum for incompressible flows while higher Mach numbers lead to small optimum heights, tending towards h/δ~0.45 at M=2.5.

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.

In conversion flight, the aeroelastic modelling of tiltrotor aircraft needs to consider the unsteady effect of the rotor wake bending due to the rotor tilting. In this paper, the unsteady models of the rotor wake bending and dynamic inflow have been introduced into the aeroelastic modelling of the tiltrotor aircraft in conversion flight by using Hamilton’s generalized principle. The method for solving the aeroelastic stability of tiltrotor aircraft in conversion flight has been established by using the small perturbation theory and the Floquet theory. The influences of unsteady dynamic inflow on trim control inputs and aeroelastic stability of a tiltrotor aircraft in conversion flight were calculated and analysed. The calculation results show that the required collective pitch increases with the pylon tilting forward and the unsteady inflow is trimmed primarily by the lateral cyclic pitch of the rotor. The wake bending unsteady dynamic inflow can obviously reduce the stability of the flapping modes of the rotor, and have no obvious influence on the lag modes of the rotor and the motion modes of the wing. The instability of tiltrotor occurs in the chordwise bending mode of the wing when the pylon tilts to a certain angle in high speed forward flight.

We report on the flight test results of an Airbus 320 during novel advanced required navigation performance (RNP) procedures which contain a fixed radius turn that delivers the aircraft onto a short instrument landing system (ILS) precision final. Moreover, the advanced RNP part contains altitude constraints and/or a coded vertical path angle. The approaches were flown automatically with guidance and autothrust as computed by the flight management system. We investigated the use of the fixed radius in conjunction with vertical path options onto (a) the performance of the speed profile for arrival time optimisation, (b) the vertical path during the RNP part of the procedure and (c) the performance of the autoland capability after a curved transition onto an ILS.

For the trials, we designed five instrument approaches to a runway equipped with ILS. A RF curve terminates at the ILS intercept point at heights of 550, 750, 1000, 1500 and 2000 ft above aerodrome level and each approach had four different initial approach fixes which corresponded to a track angle change of 30°, 60°, 90° and 180° during the constant radius turn-to-final. We constructed the procedure such that the altitude constraints at initial, intermediate and final approach fix describe a continuous vertical path with minus 2° inclination, transitioning to the –3° glide path of the ILS and intercepting the glide path from below. In all cases, the land mode of the flight guidance computer became active between 316 and 381ft radar altitude. The vertical path following error depended on the coding of the procedure in the database. With coded vertical path angle and altitude constraints, the vertical path following error was never greater than +57 m (above desired flight path) during the RNP part when flown by the automatic flight guidance system without any pilot intervention.

The laminar flow over a slender delta wing at incidence has been extensively studied both experimentally and theoretically using vortex sheet methods. These vortex sheet methods have generally been successful apart from the prediction of the secondary boundary-layer separation induced by the primary vortex. This paper revisits the problem using computational fluid dynamics (CFD) and focusses on the effects of the secondary flow separation. The modelling approach is briefly summarised, and the results are compared with flow measurements and results from vortex sheet methods. The computations show very good agreement with measurements for the surface pressures and total head contours. The results help to understand the complex structure of the leading edge vortex flow, and the associated secondary separation of the boundary layer. They indicate that inviscid mechanisms dominate the larger scale features, and highlight a possible mechanism for the development of an instability in the leading edge vortex sheet.