In this study, a series of numerical experiments are performed on supersonic/hypersonic flows over an adiabatic flat plate with transitionally and fully rough surfaces. The Mach numbers simulated are 4, 5, 6, and 7; the flight heights considered are 20, 24, 28, 32, and 36 km. First, a modified roughness correction is proposed and validated with the measured data for low-speed flat-plate cases. It is verified that for the equivalent sand grain heights in the intermediate and fully rough regimes, there is a good agreement with the semi-empirical formula available in the open literature. Then, this roughness correction is applied to high-speed flow regime to investigate the effects of flight heights and Mach numbers on drag for rough-wall flat-plate cases. It is found that within the roughness measured in real flight, the roughness height change has little effect on drag compared to the variations of both flight heights and Mach numbers. The drag coefficient derivation between rough-wall and smooth-wall conditions, achieves the maximum value of 0.79% for the 60 cases selected.

The DLR inverse design code computes the wing geometry for a prescribed target pressure distribution. It is based on the numerical solution of the integral inverse transonic small perturbation (TSP) equations. In this work, several extensions and modifications of the inverse design code are described. Results are validated with corresponding redesign test cases. The first modification concerns applications for high transonic Mach numbers or cases with strong shocks. The introduced modifications enable converged design solutions for cases where the original method failed. The second modification is the extension of the code to general non-planar wings. Previously, the design code was restricted to non-planar wing designs with small dihedral or to nacelle design. A third modification concerns aerofoil/wings designed for wind-tunnel design. In order to design a swept wing between two wind-tunnel walls, the solution method was extended to two symmetry planes. The introduced extensions and modifications have increased the robustness and range of applicability of the inverse design code.

Cranfield University's National Flying Laboratory Centre (NFLC) has developed a Bulldog light aircraft into a flight test facility. The facility is being used to research advanced in-flight instrumentation including fibre optic pressure and strain sensors. During the development of the test bed, Computational Fluid Dynamics (CFD) has been used to assist the flight test design process, including the sensor requirements. This paper describes the development of the Bulldog flight test facility, including an overview of the design and certification process, the in-flight data taken using the installed fibre optic sensor systems and lessons learned from the development programme, including potential further applications of the sensors.

In this paper, a novel thrusting manoeuvre control scheme is proposed for a small spacecraft which is based only on the gimbaled thrust vector control (TVC) system. The spacecraft structure is composed of a body and a gimbaled thruster where common attitude control systems such as reaction control system (RCS) and spin stabilisation are not employed. A nonlinear two-body model is considered for the characterisation of the gimbaled-nozzle spacecraft where the gimbal actuator provides the only active control input. The spacecraft attitude is affected by a large exogenous disturbance torque which is generated by a thrust vector misalignment from the centre of mass (CM). To achieve some performance goals in the both transient and steady-state modes, a new control scheme is derived based on the combination of two linear and nonlinear controllers. The proposed method ensures the attitude and thrust vector stability during an impulsive orbital manoeuvre while eliminating and rejecting an exogenous disturbance torque. The numerical simulations illustrate the applicability of this method for using in a small spacecraft and its efficiency in sustained operation.

This paper develops a method to determine the lift-drag ratio of air cushion craft and specifically the maximum value of the lift-drag ratio, its associated design speed, and related performance and economic measures. The method works from well-established equations for drag and power for air cushion craft. Such values are required to determine the performance and economic efficiency of the craft in many modes of operation. The method covers both amphibious and non-amphibious craft of the types used in both military and commercial operation. A basis is developed for an optimisation procedure to design future craft for maximum efficiency.

The article by Manso(1) was published online with a number of errors. The errors are shown below, along with the corrected article.

Gas path diagnostics is one of the most effective condition monitoring techniques in supporting condition-based maintenance of gas turbines and improving availability and reducing maintenance costs of the engines. The techniques can be applied to the health monitoring of different gas path components and also gas path measurement sensors. One of the most important measurement sensors is that for the engine control, also called the power setting sensor, which is used by the engine control system to control the operation of gas turbine engines. In most of the published research so far, it is rarely mentioned that faults in such sensors have been tackled in either engine control or condition monitoring. The reality is that if such a sensor degrades and has a noticeable bias, it will result in a shift in engine operating condition and misleading diagnostic results.

In this paper, the phenomenon of a power-setting sensor fault has been discussed and a gas path diagnostic method based on a Genetic Algorithm (GA) has been proposed for the detection of power-setting sensor fault with and without the existence of engine component degradation and other gas path sensor faults. The developed method has been applied to the diagnostic analysis of a model aero turbofan engine in several case studies. The results show that the GA-based diagnostic method is able to detect and quantify the power-setting sensor fault effectively with the existence of single engine component degradation and single gas path sensor fault. An exceptional situation is that the power-setting sensor fault may not be distinguished from a component fault if both faults have the same fault signature. In addition, the measurement noise has small impact on prediction accuracy. As the GA-based method is computationally slow, it is only recommended for off-line applications. The introduced GA-based diagnostic method is generic so it can be applied to different gas turbine engines.

Thermal plasma wind tunnels with power of 0.4 MW and 2.4 MW have been constructed at Chonbuk National University (CBNU) in Korea. This facility is capable of producing a heat flux greater than 10 MW/m2, a level that is relevant for testing thermal protection materials that are used for re-entry vehicles in space transportation. A segmented arc plasma torch was adopted as a plasma source; this was designed to have high thermal efficiency and long life, and to produce a supersonic plasma flow with enthalpy greater than 10 MJ/kg. We investigated the characteristics of the supersonic plasma flow using intrusive and non-intrusive diagnostic systems. Ablation characteristics of potential thermal protection materials such as carbon/carbon composites and graphite were investigated with the plasma wind tunnel. Cracks and pores in the materials accelerated the erosion. For carbon/carbon composites, the pores grew and the cracks which occurred at the interfaces between the carbon fibres and the matrix propagated, while for the graphite, the erosion started at the pores and peeled off the surface.

This research explores the modification and implementation of a Detached-Eddy Simulation (DES) in a high-order compressible solver and its application to automotive aerodynamics. This was conducted on a 20° SAE Reference Notchback Model with a Reynolds number of 2.23 × 105. This DES algorithm implemented within FLAMENCO, which is finite-volume research code operating over multi-block meshes, was used for all the simulations. The primary objectives were to capture unsteady flow features, separated coherent structures and also relax the meshing requirements to improve accessibility to turbulence-resolving methods for realistic configurations. This also aims to better understand the separated flow physics, especially around the base surfaces of the car. Simulations for three mesh refinement levels were compared to wind-tunnel measurements. Even on relatively coarse meshes (~7 m cells) for DES, time-averaged Cp was obtained with maximum errors of <8%.

Lean Blow-Out (LBO) limits are critically important in the operation of aero engines. Previously, Lefebvre's LBO empirical correlation has been extended to the flame volume concept by the authors. Flame volume takes into account the effects of geometric configuration, spatial interaction of mixing jets, turbulence, heat transfer and combustion processes inside the gas turbine combustion chamber. For these reasons, LBO predictions based on flame volume are more accurate. Although LBO prediction accuracy has improved, it poses a challenge associated with Vf estimation in real gas turbine combustors. This work extends the approach of flame volume prediction based on fuel iterative approximation with cold flow simulations to reactive flow simulations. Flame volume for 11 combustor configurations were simulated and validated against experimental data. To make prediction methodology robust, as required in preliminary design stage, reactive flow simulations were carried out with the combination of presumed Probability Density Function (PDF) and discrete phase model (DPM) in Fluent 15.0 The criterion for flame identification was defined. Two important parameters—critical injection diameter (Dp,crit) and critical temperature (Tcrit)—were identified and their influence on reactive flow simulation was studied for Vf estimation. Results exhibit ±15% error in Vf estimation with experimental data.

In this paper, an efficient approach to design and optimize a flight controller of a small-scale unmanned helicopter is proposed. Given the identified helicopter model, the Linear Quadratic Gaussian/Loop Transfer Recovery (LQG/LTR) robust control method is applied for trajectory tracking and attitude control of the helicopter with a two-loop hierarchical control architecture. Since the performance of the controller extremely depends on its weighting matrices, the Artificial Bee Colony (ABC) algorithm is introduced to automatically select the parameters of the matrices. Comparative studies between optimal algorithms are also carried out. A series of flight experiments and simulations are conducted to investigate the effectiveness and robustness of the proposed optimised controller.

Higher-Order Spectra (HOS) are used to characterise the nonlinear aeroelastic behaviour of a plunging and pitching 2-degree-of-freedom aerofoil system by diagnosing structural and/or aerodynamic nonlinearities via the nonlinear spectral content of the computed displacement signals. The nonlinear aeroelastic predictions are obtained from high-fidelity viscous fluid-structure interaction simulations. The power spectral, bi-spectral and tri-spectral densities are used to provide insight into the functional form of both freeplay and inviscid/viscous aerodynamic nonlinearities with the system displaying both low- and high-amplitude Limit Cycle Oscillation (LCO). It is shown that in the absence of aerodynamic nonlinearity (low-amplitude LCO) the system is characterised by cubic phase coupling only. Furthermore, when the amplitude of the oscillations becomes large, aerodynamic nonlinearities become prevalent and are characterised by quadratic phase coupling. Physical insights into the nonlinearities are provided in the form of phase-plane diagrams, pressure coefficient distributions and Mach number flowfield contours.

The semi-analytical prediction of pollutants emissions from gas turbines in the conceptual design phase is addressed in this paper. The necessity of this work arose from an urgent need for a comprehensive model that can quickly provide data in the conceptual design phase. Based on the available inputs data in the initial phases of the design process, a chemical reactor network (CRN) is defined to model the combustion with a detailed chemistry. In this way, three different chemical mechanisms are studied for Jet-A aviation fuel. Furthermore, the droplet evaporation for liquid fuel and the non-uniformity in fuel-air mixture are modelled. The results of a developed augmented modelling tool are compared with the pollutants data of two annular engine's combustors. The CRN results have good agreement with the actual engine test rig emissions output. In conclusion, the augmented CRN has shown to be efficient in predicting engine emissions with a very short executing time (few seconds) using a small CPU requirement such as a personal computer.

Propellers are the predominant propulsion source for small unmanned aerial vehicles. At low advance ratios, large sections of the propeller blade can be stalled, and the Reynolds number faced by each blade can be low. This leads to difficulties in modelling propeller performance, as the aerodynamic models coupled with blade element methods usually only provide aerodynamic data for an assumed aerofoil section, for a small angle-of-attack range and for a single Reynolds number, while rotational effects are often ignored. This is specifically important at low advance ratios, and a consistent evaluation of the applicability of various methods to improve aerodynamic modelling is not available. To provide a systematic appraisal, three-dimensional (3D) scanning is used to obtain the aerofoil sections that make up a propeller blade. An aerodynamic database is formed using each extracted aerofoil section, across a wide range of angles of attack and Reynolds numbers. These databases are then modified to include the effects of rotation. When compared with experimental results, significant improvement in modelling accuracy is shown at low advance ratios relative to a generic blade element-momentum model, particularly for smaller propellers. Notably, when considering small propeller performance, efficiency modelling is improved from within 30% relative to experimental data to within 5% with the use of the extended blade element momentum theory method. The results show that combining Viterna and Corrigan flat plate theory with the Corrigan and Schillings stall delay model consistently yields the closest match with experimental data.

This paper deals with the development of a simulation platform for the dynamic analysis of systems characterised by different physical domains. The research has been carried out in the context of the EC-funded Clean Sky Joint Technology Initiative (Green Regional Aircraft/All-Electric Aircraft domain). In particular, the objective of the research is focused on the on-board systems of new All-Electric Aircraft, where a crucial design point is related to the electrical energy management. In the “all-electric” concept, where pneumatic and hydraulic power systems are eliminated to improve aviation costs and environmental impact, the dynamics of electrical power absorptions is to be characterised and managed to avoid excessive peaks with respect to generators capabilities. The paper describes the architecture of a Matlab/Simulink simulation platform developed in order to design and validate of the electrical energy management logics, which lead up to 32% reduction of the maximum power request for the case study considered. Thanks to an approach based on a mixing of co-simulation and S-function compiling, the platform integrates models coming from different environments (AMESim, Dymola/Modelica), and developed by various partners/specialists.

The pressure oscillation and terminal shock motion in a two dimensional inlet, which was designed for tandem configuration turbine-based combined cycle propulsion systems was investigated experimentally and numerically, respectively. The inlet was characterised by a bleed cavity upstream the inlet throat, an S-shape rectangular-to-circular diffuser and flowpaths for a turbine and a ramjet engine. The terminal shock motion was calculated through a second-order unsteady Reynolds-averaged Navier-Stokes scheme. The pressure and the terminal shock were unsteady when the combined cycle inlet operated at different conditions. With the terminal shock located in the throat and at the shoulder of the third ramp of the TBCC inlet, the pressure oscillation was significant and the shock exhibited unsteady streamwise motion with an oscillatory pattern. The amplitude of shock oscillation at these two conditions was 6mm and 12mm, respectively. When the shock was located downstream of the throat and upstream of the cowl lip, it oscillated in a small range. We defined this motion as the “shake” of the shock. This unsteady behaviour of the shock was caused by flow separation in the combined cycle inlet diffuser.

Installation effects on the S-duct intakes of a high power turboprop were investigated by comparing three different nacelle configurations as the channel wing, over-the-wing and conventional tractor wing using computational fluid dynamic methods. The interaction of the propeller, the wing, and the nacelle on the scoop type turboprop intakes were identified in terms of recovery and distortion on the aerodynamic interface planes. An actuator disc model was used to simulate the propeller downstream effects. The channel wing installation of S-ducts showed approximately 2% higher total pressure recovery than the other configurations. The channel wing configuration with a wrap-around S-duct experienced a 2% lower circumferential distortion intensity, whereas over-the-wing had an increase by 3% in comparison to the tractor wing. It was also observed that the S-duct type had a greater influence on the radial distortion intensity than the installation type. The swirl coefficient distribution showed that multiple peaks occurred as opposed to swirl results based on a single sector of the aerodynamic interface plane. The difference between swirl peaks of shaft penetration S-duct was reduced by 14% when installed on channel wing, whereas the peaks of wrap-around experienced 90° phase shift, and their difference showed 10% increase on channel wing.

In this paper, a droplet to film interaction model technique is presented. In the proposed approach, the liquid and gas continua are modelled using an enhanced Volume-of-Fluid (VoF) technique while the droplets are tracked using a Lagrangian framework and are coupled to the Eulerian phases using source terms. The eventual target application is an aeroengine bearing chamber in which oil is found as droplets, shed from the bearings, splashing on impact, separated from wall surfaces at obstacles or simply re-entrained, and as a continuum oil film coating the bearing chamber outer walls which it also cools. In finite volume Computational Fluid Dynamics (CFD) techniques, a prohibitively large number of cells would be required to describe the details of the droplet impact phenomenon. Based on published correlations, the splashing droplets are created and tracked as Lagrangian particles. The flowing film and the gas continua are handled with an enhanced VoF technique.

Circulation control via blowing over Coanda surfaces at transonic freestream Mach numbers is investigated using numerical simulations. The performance and sensitivity of several circulation control devices applied to a supercritical aerofoil are assessed. Different Coanda devices were studied to assess the effect of Coanda radius-to-slot height ratio, nozzle shape and Coanda surfaces with a step. The range of operating conditions for which a supersonic Coanda jet remained attached at transonic freestream conditions were extended by increasing the radius of curvature at the slot exit for Coanda devices with a converging nozzle. Additional improvements were found by reducing the strength of shock boundary-layer interactions on the Coanda surface by expanding the jet flow using a converging-diverging nozzle and also by introducing a step between the Coanda surface and the nozzle exit. The performance when using a converging-diverging nozzle can be matched using a simple stepped Coanda device. It is shown that circulation control has the potential to match the performance of traditional control surfaces during regimes of attached flow at transonic speeds, up to an equivalent aileron deflection angle of 10°. In addition, lift augmentation ratios ΔCl/Cμ of over 100 were achieved.